CN112362289A - Aircraft split test method and device, computer equipment and readable storage medium - Google Patents

Aircraft split test method and device, computer equipment and readable storage medium Download PDF

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CN112362289A
CN112362289A CN202110043161.9A CN202110043161A CN112362289A CN 112362289 A CN112362289 A CN 112362289A CN 202110043161 A CN202110043161 A CN 202110043161A CN 112362289 A CN112362289 A CN 112362289A
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test
booster
preset
parameter
aircraft
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CN112362289B (en
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郭鹏
钱丰学
高鹏
刘奇
易国庆
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Ultra High Speed Aerodynamics Institute China Aerodynamics Research and Development Center
High Speed Aerodynamics Research Institute of China Aerodynamics Research and Development Center
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Ultra High Speed Aerodynamics Institute China Aerodynamics Research and Development Center
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
    • G01M9/02—Wind tunnels
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
    • G01M9/06—Measuring arrangements specially adapted for aerodynamic testing
    • G01M9/062—Wind tunnel balances; Holding devices combined with measuring arrangements
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
    • G01M9/08—Aerodynamic models

Abstract

The application discloses an aircraft split test method, an aircraft split test device, computer equipment and a readable storage medium, and relates to the technical field of wind tunnel tests, wherein the aircraft split test method comprises the following steps: generating an initialization signal and sending the initialization signal to the mechanism; sending a starting signal to the wind tunnel, and collecting aerodynamic force data of the test model; judging whether a preset constraint formula is established or not according to the centroid position of the test model at a first preset moment; if so, calculating to obtain a first test parameter of the test model at a second preset moment; if not, calculating to obtain a second test parameter of the test model at a second preset time; inputting a first test parameter or a second test parameter into the mechanism; and repeating the step of sending an initialization signal to the mechanism to input the first test parameter or the second test parameter into the mechanism, and stopping the test and acquiring the target parameter when the preset stop moment is reached, so that the target parameter can be reliably acquired.

Description

Aircraft split test method and device, computer equipment and readable storage medium
Technical Field
The application relates to the field of wind tunnel tests, in particular to an aircraft split test method, an aircraft split test device, computer equipment and a readable storage medium.
Background
An important link in the flight process of the multi-stage reusable carrier is interstage separation, and the aim is to reduce the overall mass and further improve the carrying capacity under the condition of limited cost by separating and discarding a task completion part. In the prior art, a parallel layout is often adopted between two stages of a new generation lift body layout reusable vehicle, such as a Gryphon (magson) vehicle system proposed by the united states air force, a tsto (two Stage to one track) vehicle of NASA (National Aeronautics and Space Administration, abbreviated as the National Space and flight Administration), and a new generation vehicle proposed by Japan office of Aerospace research and development, because the two stages have similar sizes, the pneumatic interference existing in the interstage separation process is not negligible for any Stage, so a special separation mechanism design is usually adopted, such as a typical aircraft has a spherical hinge double support in the two stages, so as to avoid task failure caused by contact between the two stages in the separation process, and the similar special design also adds additional conditions or constraints to the separation process, similarly, taking the aircraft separation process as an example, after the separation starts, the two-stage front part can be firstly separated under the action of aerodynamic force, the tail ball-like hinge mechanism is still in a contact state and can generate support reaction force, the size of the support reaction force can change along with the relative motion state of the two-stage aircraft and the aerodynamic force borne by each stage, then the carrier and the booster start to move under the action of the respective aerodynamic force, thrust, gravity and contact position support reaction force, the support reaction force of the contact position is gradually reduced to 0 along with the separation process, the carrier and the booster are gradually separated at the tail ball-like hinge and reach a certain distance, and then the separation process can be considered to be completed. In the conventional wind tunnel interstage separation device or a CTS (positive track System, abbreviated as a motion Trajectory capturing System), after separation, both stages are considered to be in a free flight state, and no constraint force exists between the two stages, so that the research requirements on the separation problem cannot be met at all.
In view of the above, how to provide a reliable split test scheme for an aircraft is a problem to be solved by those skilled in the art.
Disclosure of Invention
The embodiment of the application provides an aircraft split test method, an aircraft split test device, computer equipment and a readable storage medium.
In a first aspect, an embodiment of the present application provides an aircraft split test method, which is applied to a computer device, where the computer device is electrically connected to a mechanism, a wind tunnel, and a test model, the mechanism is used to control a posture of the test model, the wind tunnel is used to provide a test environment of the test model, and the method includes:
generating an initialization signal and sending the initialization signal to the mechanism so that the mechanism performs initial position setting on the test model according to the initialization signal;
sending a starting signal to the wind tunnel so as to start the wind tunnel according to a preset Mach number, and acquiring aerodynamic force data and test data of the test model;
judging whether a preset constraint formula is established or not according to the centroid position of the test model at a first preset moment;
if yes, inputting the test data into a constraint kinetic equation determined according to the aerodynamic force data, and calculating to obtain a first test parameter of the test model at a second preset moment;
if not, inputting the test data into a non-constrained kinetic equation determined according to the aerodynamic force data, and calculating to obtain a second test parameter of the test model at a second preset time;
inputting the first test parameter or the second test parameter into the mechanism so that the mechanism can adjust the position according to the first test parameter or the second test parameter;
and repeating the steps of sending the initialization signal to the mechanism and inputting the first test parameter or the second test parameter into the mechanism, and stopping the test and acquiring the target parameter when the preset stop moment is reached.
Optionally, the initialization signal includes a mechanism initialization motion compensation amount, and the step of generating the initialization signal includes:
calculating to obtain an initial attack angle according to preset ballistic parameters and preset attitude information;
and calculating to obtain the mechanism initialization motion compensation quantity according to the initial attack angle.
Optionally, the test model includes a carrier and a booster, and the preset constraint formula is:
Figure 891655DEST_PATH_IMAGE001
wherein, | | RK1-RK2| is 2 norm calculation formula, reIn order to constrain the distance, the distance is constrained,
Figure 954289DEST_PATH_IMAGE002
to constrain the angle, phieTo constrain the angle threshold, RK1Is the centroid position, R, of the carrier at a first predetermined momentK2Is the center of mass position v of the booster at a first preset momentxK1Is the speed of the vehicle in a first direction in the inertial system, vyK1Is the speed, v, of the vehicle in a second direction in the inertial systemxK2Is the speed, v, of the booster in a first direction in the inertial systemyK2Is the speed of the booster in a second direction in the inertial system, theta1Is the attitude angle of the carrier in the inertial system.
Optionally, the inputting the test data into a constraint kinetic equation determined from the aerodynamic data:
Figure 579436DEST_PATH_IMAGE003
a first test parameter is calculated, wherein,
Figure 284087DEST_PATH_IMAGE004
=
Figure 560348DEST_PATH_IMAGE005
Figure 793883DEST_PATH_IMAGE006
Figure 155594DEST_PATH_IMAGE007
=
Figure 867198DEST_PATH_IMAGE008
wherein m is1Is the mass of the carrier, /)1The distance from the center of mass of the carrier to the connection point of the carrier and the booster (x)1，y1) Is the coordinate of the center of mass of the aircraft in the inertial system, J1Moment of inertia being the centre of mass of the aircraft, Fx1Is a generalized active force vector of the first direction of the vehicle in the inertial system, Fy1Is a generalized active force vector of the second direction of the vehicle in the inertial system, M1Is a generalized mass matrix of the carrier, g is the acceleration of gravity,
Figure 276926DEST_PATH_IMAGE009
the angular velocity, m, of the carrier in the inertial system2To the mass of the booster, /)2Is the distance from the center of mass of the booster to the point of attachment of the carrier to the booster, J2Moment of inertia being the centre of mass of the booster, Fx2Is a generalized primary force vector of the booster in a first direction in the inertial system, Fy2Is a generalized primary force vector, θ, of the booster in a second direction in the inertial system2For the attitude angle, M, of the booster in the inertial system2In order to provide a generalized mass matrix for the booster,
Figure 681363DEST_PATH_IMAGE010
to determine the angular velocity of the booster in the inertial system,
Figure 530370DEST_PATH_IMAGE004
in order to be a quality matrix,
Figure 45665DEST_PATH_IMAGE011
in the form of a matrix of accelerations,
Figure 296518DEST_PATH_IMAGE012
in the form of a matrix of aerodynamic forces,
Figure 137435DEST_PATH_IMAGE011
as a matrix of position coordinates
Figure 473738DEST_PATH_IMAGE013
Obtained after two successive derivations.
Optionally, the inputting the test data into an unconstrained kinetic equation determined from the aerodynamic data:
Figure 543457DEST_PATH_IMAGE014
calculating to obtain a second test parameter, wherein M1Is an airborne generalized quality matrix and is,
Figure 648816DEST_PATH_IMAGE015
generalized acceleration of the vehicle, F1Is the generalized active force vector of the aircraft, G1Is the gravity vector of the aircraft, M2In order to provide a generalized mass matrix for the booster,
Figure 395055DEST_PATH_IMAGE016
for generalized acceleration of the booster, F2Is the generalized primary force vector of the booster, G2Is the gravity vector of the booster.
Optionally, the first test parameter includes a compensation amount of the mechanism, and the step of inputting the first test parameter into the mechanism to adjust the position of the mechanism according to the first test parameter includes:
and inputting the mechanism compensation quantity into the mechanism so as to enable the mechanism to carry out position adjustment according to the first test parameter, wherein the mechanism compensation quantity comprises the mass center position, the speed, the attitude angle and the angular speed of the test model at a second preset moment.
In a second aspect, an embodiment of the present application provides an aircraft components of a whole that can function independently test device, is applied to computer equipment, computer equipment is connected with mechanism, wind-tunnel and test model electricity respectively, the mechanism is used for control test model's gesture, the wind-tunnel is used for providing test model's experimental environment, the device includes:
the generating module is used for generating an initialization signal and sending the initialization signal to the mechanism so that the mechanism can perform initial position setting on the test model according to the initialization signal;
the transmitting module is used for transmitting a starting signal to the wind tunnel so as to start the wind tunnel according to a preset Mach number and acquiring aerodynamic force data and test data of the test model;
the judging module is used for judging whether a preset constraint formula is established or not according to the centroid position of the test model at a first preset moment; if yes, inputting the test data into a constraint kinetic equation determined according to the aerodynamic force data, and calculating to obtain a first test parameter of the test model at a second preset moment; if not, inputting the test data into a non-constrained kinetic equation determined according to the aerodynamic force data, and calculating to obtain a second test parameter of the test model at a second preset time;
the adjusting module is used for inputting the first test parameter or the second test parameter into the mechanism so as to enable the mechanism to adjust the position according to the first test parameter or the second test parameter;
and the test module is used for repeatedly sending the initialization signal to the mechanism and inputting the first test parameter or the second test parameter into the mechanism, and when the preset stop moment is reached, stopping the test and acquiring the target parameter.
Optionally, the initialization signal includes a mechanism initialization motion compensation amount, and the generating module is specifically configured to:
calculating to obtain an initial attack angle according to preset ballistic parameters and preset attitude information; and calculating to obtain the mechanism initialization motion compensation quantity according to the initial attack angle.
In a third aspect, embodiments of the present application provide a computer device, where the computer device includes a processor and a non-volatile memory storing computer instructions, and when the computer instructions are executed by the processor, the computer device executes the aircraft split test method.
In a fourth aspect, an embodiment of the present application provides a readable storage medium, where the readable storage medium includes a computer program, and the computer program controls, when running, a computer device on which the readable storage medium is located to perform the aircraft split test method described above.
Compared with the prior art, the beneficial effects provided by the application comprise: by adopting the aircraft split test method, the aircraft split test device, the computer equipment and the readable storage medium, the mechanism performs initial position setting on the test model according to the initialization signal by generating the initialization signal and sending the initialization signal to the mechanism; sending a starting signal to the wind tunnel so as to start the wind tunnel according to a preset Mach number, and acquiring aerodynamic force data and test data of the test model; then judging whether a preset constraint formula is established or not according to the centroid position of the test model at a first preset moment; if yes, inputting the test data into a constraint kinetic equation determined according to the aerodynamic force data, and calculating to obtain a first test parameter of the test model at a second preset moment; if not, inputting the test data into a non-constrained kinetic equation determined according to the aerodynamic force data, and calculating to obtain a second test parameter of the test model at a second preset time; inputting the first test parameter or the second test parameter into the mechanism so that the mechanism can adjust the position according to the first test parameter or the second test parameter; and repeating the steps from sending the initialization signal to the mechanism to inputting the first test parameter or the second test parameter into the mechanism, stopping the test and acquiring the target parameter when the preset stop moment is reached, and reliably acquiring the data required by the aircraft separation test.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings needed to be used in the embodiments will be briefly described below. It is appreciated that the following drawings depict only certain embodiments of the application and are therefore not to be considered limiting of its scope. For a person skilled in the art, it is possible to derive other relevant figures from these figures without inventive effort.
FIG. 1 is a schematic flow chart illustrating steps of a method for split testing an aircraft according to an embodiment of the present application;
FIG. 2 is a schematic diagram of a separation system for test models according to an embodiment of the present disclosure;
FIG. 3 is a simplified schematic diagram of a separation system for test models according to an embodiment of the present disclosure;
FIG. 4 is a schematic structural diagram of an aircraft split test device provided in the embodiment of the present application;
fig. 5 is a schematic structural diagram of a computer device according to an embodiment of the present application.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. It is to be understood that the embodiments described are only a few embodiments of the present application and not all embodiments. The components of the embodiments of the present application, generally described and illustrated in the figures herein, can be arranged and designed in a wide variety of different configurations.
Thus, the following detailed description of the embodiments of the present application, presented in the accompanying drawings, is not intended to limit the scope of the claimed application, but is merely representative of selected embodiments of the application. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.
It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, it need not be further defined and explained in subsequent figures.
In the description of the present application, it is to be understood that the terms "upper", "lower", "inner", "outer", "left", "right", and the like, refer to orientations or positional relationships that are based on the orientations or positional relationships shown in the drawings, or the orientations or positional relationships that the products of the application conventionally position when in use, or the orientations or positional relationships that are conventionally understood by those skilled in the art, and are used for convenience in describing the present application and simplifying the description, but do not indicate or imply that the referenced devices or elements must have a particular orientation, be constructed in a particular orientation, and be operated, and therefore, should not be construed as limiting the present application.
Furthermore, the terms "first," "second," and the like are used merely to distinguish one description from another, and are not to be construed as indicating or implying relative importance.
In the description of the present application, it is also to be noted that, unless otherwise explicitly stated or limited, the terms "disposed" and "connected" are to be interpreted broadly, for example, "connected" may be a fixed connection, a detachable connection, or an integral connection; can be mechanically or electrically connected; the connection may be direct or indirect via an intermediate medium, and may be a communication between the two elements. The specific meaning of the above terms in the present application can be understood by those of ordinary skill in the art as appropriate.
The following detailed description of embodiments of the present application will be made with reference to the accompanying drawings.
At present, an important link in the flight process of the multistage reusable carrier is interstage separation, and the aim is to reduce the overall mass and further improve the carrying capacity under the condition of limited cost by separating and discarding a task completion part. Based on this, the embodiment of the application provides an aircraft split test method, which is applied to computer equipment, the computer equipment is respectively and electrically connected with a mechanism, a wind tunnel and a test model, the mechanism is used for controlling the posture of the test model, the wind tunnel is used for providing the test environment of the test model, and as shown in fig. 1, the aircraft split test method comprises the following steps:
step 201, generating an initialization signal and sending the initialization signal to the mechanism, so that the mechanism performs initial position setting on the test model according to the initialization signal.
Step 202, sending a starting signal to the wind tunnel so as to start the wind tunnel according to the preset Mach number, and collecting aerodynamic force data and test data of the test model.
And 203, judging whether a preset constraint formula is established or not according to the centroid position of the test model at the first preset moment.
If yes, go to step 204.
And 204, inputting the test data into a constraint kinetic equation determined according to the aerodynamic force data, and calculating to obtain a first test parameter of the test model at a second preset time.
If not, go to step 205.
And step 205, inputting the test data into a non-constrained kinetic equation determined according to the aerodynamic force data, and calculating to obtain a second test parameter of the test model at a second preset time.
Step 206, inputting the first test parameter or the second test parameter into the mechanism, so that the mechanism can adjust the position according to the first test parameter or the second test parameter.
And repeatedly executing the step 201 to the step 206 until the preset stop time is reached, and executing the step 207.
Step 207, stop the test and obtain the target parameters.
On the basis of the above, the initialization signal includes the mechanism initialization motion compensation amount, and the embodiment of the present application also provides an example of generating the initialization signal, which can be implemented by the following steps.
And calculating to obtain an initial attack angle according to the preset ballistic parameters and the preset attitude information.
And calculating to obtain the mechanism initialization motion compensation amount according to the initial attack angle.
Referring to fig. 2, the test model 1 includes a carrier 10 and a booster 20, and when modeling, the model may satisfy the following conditions: the separation process is carried out in a plane of longitudinal symmetry of the carrier 10 and booster 20, and parallel to the direction of the local plumb line; the heights of the carrier 10 and the booster 20 and the length of the separating mechanism can be ignored, and the connecting mechanism is only considered as one point in the modeling; once the carrier 10 and the booster 20 exceed a certain distance, the constraint action possibly generated by the separation mechanism is not considered; the friction forces present during the separation process can be neglected. On the basis, please refer to fig. 3, an inertial coordinate system can be established, wherein the carrier 10 coordinate system O1X1Y1Origin O of1Located in the center of mass, O, of the carrier 101X1The center of mass points to the tail, O, of the carrier 101Y1Perpendicular to O1X1The shaft is directed above the carrier 10. Booster 20 body coordinate system O2X2Y2Origin O of2Located in the center of mass, O, of the booster 202X2Center of mass pointing to the tail, O, of the booster 202Y2Perpendicular to O2X2The shaft is directed above the booster 20. Local inertial coordinate system OEXEYEIs the local geodetic coordinate system, O thereofEXEThe axis being located at the local level behind the pointing carrier 10 and booster 20, OEXEThe shaft points upward against the local plumb line. It can be noted that the connecting point K1 on the aircraft 10 is far from the centroid O of the aircraft 101A distance of l1The connecting point K2 on the booster 20 is far from the mass center O of the booster 202A distance of l2As is understood by definition, at the initial moment of separation, the carrier 10 and the tail of the booster 20 are in a connected state, where K1 is at the same position as K2, which is denoted as point K in fig. 2.
In the present embodiment, the kinetic equations of the vehicle 10 and the thrusters 20 can be established under constrained conditions and unconstrained conditions, respectively.
For two free-running rigid bodies (the carrier 10 and the thrusters 20) with unconstrained conditions after separation, the kinetic equations can be expressed as:
Figure 953075DEST_PATH_IMAGE017
in the formula, m1And J1Mass of the vehicle 10 and moment of inertia, m, relative to the center of mass2And J2The mass of the booster 20 and the moment of inertia of the relative center of mass, F1=（Fx1，Fy1，M1）TAnd F2=（Fx2，Fy2，M2）TThe generalized active force vectors received by the carrier 10 and the booster 20, g is the gravity acceleration of the test ground,
Figure 75752DEST_PATH_IMAGE018
in order to be the acceleration of the vehicle 10,
Figure 35618DEST_PATH_IMAGE019
is the rotational acceleration of the carrier 10,
Figure 952758DEST_PATH_IMAGE020
in order to accelerate the booster 20,
Figure 263654DEST_PATH_IMAGE021
is the rotational acceleration of the booster 20.
On the basis of the above, the experimental data can be input into an unconstrained kinetic equation determined from the aerodynamic data to be converted into a vector form as follows.
Figure 675175DEST_PATH_IMAGE022
Calculating to obtain a second test parameter, wherein M1Is an airborne generalized quality matrix and is,
Figure 755126DEST_PATH_IMAGE015
generalized acceleration of the vehicle, F1Is the generalized active force vector of the aircraft, G1Is the gravity vector of the aircraft, M2In order to provide a generalized mass matrix for the booster,
Figure 108747DEST_PATH_IMAGE016
for generalized acceleration of the booster, F2Is the generalized primary force vector of the booster, G2Is the gravity vector of the booster.
When there is a constraint during the separation of the carrier 10 and the booster 20, by definition, the following constraint equation can be obtained:
Figure 641360DEST_PATH_IMAGE023
wherein A is1And A2Coordinate transformation matrixes from the inertial system to the body coordinate system of the carrier 10 and the body coordinate system of the booster 20 respectively have the following expressions:
Figure 840260DEST_PATH_IMAGE024
on the basis of the foregoing, experimental data are input into a constraint kinetic equation determined from aerodynamic data:
Figure 774718DEST_PATH_IMAGE025
a first test parameter is calculated, wherein,
Figure 299240DEST_PATH_IMAGE004
=
Figure 804302DEST_PATH_IMAGE026
Figure 72472DEST_PATH_IMAGE006
Figure 127016DEST_PATH_IMAGE007
=
Figure 556860DEST_PATH_IMAGE008
wherein m is1Is the mass of the carrier, /)1The distance from the center of mass of the carrier to the connection point of the carrier and the booster (x)1，y1) Is the coordinate of the center of mass of the aircraft in the inertial system, J1Moment of inertia being the centre of mass of the aircraft, Fx1Is a generalized active force vector of the first direction of the vehicle in the inertial system, Fy1Is a generalized active force vector of the second direction of the vehicle in the inertial system, M1Is a generalized mass matrix of the carrier, g is the acceleration of gravity,
Figure 64065DEST_PATH_IMAGE027
the angular velocity, m, of the carrier in the inertial system2To the mass of the booster, /)2Is the distance from the center of mass of the booster to the point of attachment of the carrier to the booster, J2Moment of inertia being the centre of mass of the booster, Fx2Is a generalized primary force vector of the booster in a first direction in the inertial system, Fy2Is a generalized primary force vector, θ, of the booster in a second direction in the inertial system2For the attitude angle, M, of the booster in the inertial system2In order to provide a generalized mass matrix for the booster,
Figure 870347DEST_PATH_IMAGE010
to determine the angular velocity of the booster in the inertial system,
Figure 513818DEST_PATH_IMAGE004
in order to be a quality matrix,
Figure 127946DEST_PATH_IMAGE011
in the form of a matrix of accelerations,
Figure 856867DEST_PATH_IMAGE012
in the form of a matrix of aerodynamic forces,
Figure 466840DEST_PATH_IMAGE011
as a matrix of position coordinates
Figure 230397DEST_PATH_IMAGE013
Obtained after two successive derivations.
Under the constraint equation, the whole system has only 4 degrees of freedom, and at this time, the above-mentioned unconstrained kinetic equation cannot describe the motion state of the vehicle 10 and the booster 20 in space, so that the equation needs to be re-derived. In the embodiment of the present application, Lagrange's equation and Kane's method may be used to establish the kinetic equation under the constraint of the tail of the carrier 10 and the booster 20.
In the embodiment of the present application, the kinetic energy Q of the vehicle 10 is based on Lagrange's equation1Can be expressed as:
Figure 267623DEST_PATH_IMAGE028
wherein the content of the first and second substances,
Figure 749420DEST_PATH_IMAGE029
the speed of the first line of defense of the vehicle 10 in the inertial system,
Figure 897504DEST_PATH_IMAGE030
is the speed of the vehicle 10 in the second direction in the inertial system. A
And the kinetic energy Q of the booster 202Can be expressed as:
Figure 266300DEST_PATH_IMAGE031
wherein the content of the first and second substances,
Figure 208848DEST_PATH_IMAGE032
for the velocity of the booster 20 in the first direction in the inertial system,
Figure 912362DEST_PATH_IMAGE033
is the velocity of the booster 20 in the second direction in the inertial system.
The kinetic energy Q of the test model 1 can be expressed as:
Q=Q1+Q2
again according to the aforementioned existing constraint equation:
Figure 129717DEST_PATH_IMAGE023
deriving Q simultaneously for both sides of its equal sign can result in:
Figure 602286DEST_PATH_IMAGE034
the kinetic equation of the system under the constraint condition can be obtained by substituting the kinetic energy Q into Lagrange equation through simplification:
Figure 981315DEST_PATH_IMAGE003
in addition to this, selection is possible
Figure 188437DEST_PATH_IMAGE029
、
Figure 678324DEST_PATH_IMAGE035
、
Figure 270979DEST_PATH_IMAGE027
And
Figure 820909DEST_PATH_IMAGE010
as a generalized rate, the carrier 10 and the velocity can be adjusted under the aforementioned constraintsThe generalized azimuthal vector of the booster 20 is given by:
r1=（x1，y1，θ1）T
r2=r1+l1（cosθ1，sinθ1，0）T-l2（cosθ2，sinθ2，0）T+（0，0，θ2-θ1）T
the derivation of the kinetic energy Q can be found based on the two formulas:
Figure 233436DEST_PATH_IMAGE036
Figure 792593DEST_PATH_IMAGE037
according to the above formula, the speed v of the carrier 101And the velocity v of the booster 202The derivation can yield the acceleration of the vehicle 10 and the acceleration of the thrusters 20:
Figure 239755DEST_PATH_IMAGE038
Figure 711319DEST_PATH_IMAGE039
while being based on the speed v of the carrier 101And the velocity v of the booster 202Various offset velocity expressions can be obtained:
Figure 876721DEST_PATH_IMAGE040
Figure 708411DEST_PATH_IMAGE041
on the basis of the aboveAccording to Kane kinetic equation, system generalized inertia force
Figure 10079DEST_PATH_IMAGE042
And generalized main power
Figure 901812DEST_PATH_IMAGE043
It should satisfy:
Figure 554510DEST_PATH_IMAGE044
+
Figure 189891DEST_PATH_IMAGE045
=0
based on this, it can be found from the aforementioned formula:
Figure 382885DEST_PATH_IMAGE046
the same kinetic equation can be compared with a system in which the test model 1 is located under the constraint condition obtained based on the Lagrange equation and the Kane equation, and can be expressed as follows:
Figure 445519DEST_PATH_IMAGE047
wherein the content of the first and second substances,
Figure 585513DEST_PATH_IMAGE048
in order to be a quality matrix,
Figure 759006DEST_PATH_IMAGE049
in the form of a matrix of accelerations,
Figure 35266DEST_PATH_IMAGE012
is an aerodynamic matrix.
On the basis of the foregoing, an embodiment of the present application provides a method for determining a preset constraint formula according to a centroid position of a test model 1 at a first preset time, where the preset constraint formula is:
Figure 268801DEST_PATH_IMAGE001
wherein, | | RK1-RK2| is 2 norm calculation formula, reIn order to constrain the distance, the distance is constrained,
Figure 630513DEST_PATH_IMAGE002
to constrain the angle, phieTo constrain the angle threshold, RK1Is the centroid position, R, of the carrier at a first predetermined momentK2Is the center of mass position v of the booster at a first preset momentxK1Is the speed of the vehicle in a first direction in the inertial system, vyK1Is the speed, v, of the vehicle in a second direction in the inertial systemxK2Is the speed, v, of the booster in a first direction in the inertial systemyK2Is the speed of the booster in a second direction in the inertial system, theta1Is the attitude angle of the carrier in the inertial system. In the embodiment of the present application, the constraint angle threshold φ may be seteSet at 89 deg. to facilitate the test.
On the basis of the above, the step of inputting the first test parameter into the mechanism so that the mechanism performs position adjustment according to the first test parameter includes:
and inputting the mechanism compensation quantity into the mechanism so that the mechanism can carry out position adjustment according to the first test parameter, wherein the mechanism compensation quantity comprises the mass center position, the speed, the attitude angle and the angular speed of the test model 1 at a second preset moment.
Specifically, before the above-mentioned test is performed, the model may be installed and leveled according to an interstage separation test, and the two-stage model is in place at an angle of attack of 0 °, where the angle of attack is also called an attack angle, in this embodiment, the angle is an included angle between a projection of a velocity vector of the two-stage model on a longitudinal symmetric plane and a longitudinal axis of the two-stage model, the upward head is positive, the downward head is negative, and the angle of attack of 0 ° is zero, which is a state where the angle of attack is zero. It should be understood that the two-level model is test model 1.
Then, according to initial separation conditions, an initial flow field mach number and a separation height are given, a model attack angle (namely an initial attack angle) at an initial moment is calculated according to ballistic parameters and attitude information, and a mechanism initialization motion compensation quantity (namely an initialization motion compensation quantity) is calculated according to the initial attack angle.
The initial motion compensation amount of the mechanism is input into the mechanism so that the mechanism adjusts the test model 1 (i.e., the carrier 10 and the booster 20) to be in position at a given initial angle of attack to achieve the desired initial separation moment state (i.e., initial position setting).
And sending a starting signal to the wind tunnel so as to start the wind tunnel according to the preset Mach number, and acquiring aerodynamic force data of the test model 1 by a balance arranged on the test model 1. It should be noted that the acquired aerodynamic force data of the test model 1 are used for solving the two-stage aerodynamic forces of the test model 1 under the actual separation height according to the acquired aerodynamic force coefficient of the test model 1 and the ballistic information, and then the generalized active force (i.e., the updated generalized active force) of the test model 1 at the moment (i.e., the updated generalized active force) can be calculated according to the aerodynamic force
Figure 358428DEST_PATH_IMAGE050
）。
Then according to at tiAnd (3) solving the coordinates of the contact position by the two-stage centroid position of the test model 1 at the moment (namely the first moment), and judging whether the constraint condition is met or not through the formula. If so, inputting the acquired dynamic state parameters such as two-stage centroid positions, motion speeds, attitude angles, angular speeds and the like into a dynamic equation under the constraint condition for solving; if not, inputting the acquired dynamic state parameters such as two-stage centroid positions, motion speeds, attitude angles, angular speeds and the like into the dynamic equation under the unconstrained condition for solving.
T can be obtained by solving the kinetic equation under the constraint condition or the kinetic equation under the unconstrained conditioniThe centroid position, velocity, attitude angle, and angular velocity of the two-stage test model 1 at the time + Δ t (second time). Whether the termination condition is met can be judged, in the embodiment of the application, the termination condition can be that the test is carried out to the preset stop moment, or the test is stopped by the user actively inputting a stop instruction. When it is satisfied withAt the termination condition, the data collected at that time (i.e., the target parameter) is recorded, which may be included at tiThe + Δ t time (second time) is data of the centroid position, velocity, attitude angle, angular velocity, and the like of the two-stage test model 1. When the termination condition is not satisfied, according to tiAt the + Δ t moment (second moment), the centroid position, the velocity, the attitude angle and the angular velocity of the two-stage test model 1 are calculated to obtain a new compensation amount of the mechanism (i.e. as a new initial compensation amount), and then the above test is repeated until the termination condition is satisfied.
The embodiment of the application provides an aircraft components of a whole that can function independently test device 110, is applied to computer equipment, and computer equipment is connected with mechanism, wind-tunnel and test model 1 electricity respectively, and the mechanism is used for controlling test model 1's gesture, and the wind-tunnel is used for providing test model 1's experimental environment, and as shown in fig. 4, the device includes:
a generating module 1101, configured to generate an initialization signal and send the initialization signal to the mechanism, so that the mechanism performs initial position setting on the test model 1 according to the initialization signal.
The sending module 1102 is configured to send a start signal to the wind tunnel, so that the wind tunnel is started according to a preset mach number, and aerodynamic data of the test model 1 is collected.
The judging module 1103 is configured to judge whether the preset constraint formula is satisfied according to a centroid position of the test model 1 at a first preset time; if yes, inputting the test data into a constraint kinetic equation determined according to the aerodynamic force data, and calculating to obtain a first test parameter of the test model 1 at a second preset moment; and if not, inputting the test data into a non-constrained kinetic equation determined according to the aerodynamic force data, and calculating to obtain a second test parameter of the test model 1 at a second preset time.
And the adjusting module 1104 is used for inputting the first test parameter or the second test parameter into the mechanism so as to adjust the position of the mechanism according to the first test parameter or the second test parameter.
The test module 1105 is configured to repeat the steps from sending the initialization signal to the mechanism to inputting the first test parameter or the second test parameter into the mechanism, and when a preset stop time is reached, stop the test and obtain the target parameter.
Further, the initialization signal includes a mechanism initialization motion compensation amount, and the generating module 1101 is specifically configured to:
calculating to obtain an initial attack angle according to preset ballistic parameters and preset attitude information; and calculating to obtain the mechanism initialization motion compensation amount according to the initial attack angle.
In the embodiment of the present application, reference may be made to the implementation principle of the aircraft split test method for the specific implementation principle of the aircraft split test device 110, and details are not described herein again.
The embodiment of the application provides a computer device 100, where the computer device 100 includes a processor and a nonvolatile memory storing computer instructions, and when the computer instructions are executed by the processor, the computer device 100 executes the aircraft split test method. As shown in fig. 5, fig. 5 is a block diagram of a computer device 100 according to an embodiment of the present disclosure. The computer apparatus 100 includes an aircraft split test apparatus 110, a memory 111, a processor 112, and a communication unit 113. The double arrows in fig. 5 indicate that interactive communication is possible between the various devices.
To achieve data transmission or interaction, the memory 111, the processor 112 and the communication unit 113 are electrically connected to each other directly or indirectly. For example, the components may be electrically connected to each other via one or more communication buses or signal lines. The aircraft split test device 110 includes at least one software function module which can be stored in the memory 111 in the form of software or firmware (firmware) or is fixed in an Operating System (OS) of the computer device 100. The processor 112 is configured to execute executable modules stored in the memory 111, such as software functional modules and computer programs included in the aircraft split test apparatus 110.
The embodiment of the application provides a readable storage medium, which comprises a computer program, and when the computer program runs, the computer program controls computer equipment where the readable storage medium is located to execute the aircraft split test method.
In summary, by using the aircraft split test method, the aircraft split test device, the aircraft split test computer equipment and the readable storage medium provided by the embodiment of the application, the mathematical model of the separation process between the lower levels of the constraint condition is deduced by using a multi-body dynamics method, a reasonable constraint failure condition is established, if the constraint condition is established in the separation process, the mathematical model under the constraint condition is used for solving the kinetic parameters of the two-level separation process, and if the constraint is not established, the mathematical model under the constraint-free condition is used for solving the kinetic parameters of the separation process; and then, upgrading the program of the wind tunnel interstage separation device based on a newly derived equation to realize the test simulation of the problems. The simulation of the separation process under the constraint condition can be realized by depending on the input quantity of the conventional wind tunnel interstage separation device without independently increasing the constraint force as the input quantity.
The above description is only a preferred embodiment of the present application and is not intended to limit the present application, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, improvement and the like made within the spirit and principle of the present application shall be included in the protection scope of the present application.

Claims (10)

1. The aircraft split test method is characterized by being applied to computer equipment, wherein the computer equipment is respectively and electrically connected with a mechanism, a wind tunnel and a test model, the mechanism is used for controlling the posture of the test model, the wind tunnel is used for providing the test environment of the test model, and the method comprises the following steps:
generating an initialization signal and sending the initialization signal to the mechanism so that the mechanism performs initial position setting on the test model according to the initialization signal;
sending a starting signal to the wind tunnel so as to start the wind tunnel according to a preset Mach number, and acquiring aerodynamic force data and test data of the test model;
judging whether a preset constraint formula is established or not according to the centroid position of the test model at a first preset moment;
if yes, inputting the test data into a constraint kinetic equation determined according to the aerodynamic force data, and calculating to obtain a first test parameter of the test model at a second preset moment;
if not, inputting the test data into a non-constrained kinetic equation determined according to the aerodynamic force data, and calculating to obtain a second test parameter of the test model at a second preset time;
inputting the first test parameter or the second test parameter into the mechanism so that the mechanism can adjust the position according to the first test parameter or the second test parameter;
and repeating the steps of sending the initialization signal to the mechanism and inputting the first test parameter or the second test parameter into the mechanism, and stopping the test and acquiring the target parameter when the preset stop moment is reached.
2. The aircraft split trial method of claim 1, wherein the initialization signal comprises a mechanism initialization motion compensation quantity, and the step of generating the initialization signal comprises:
calculating to obtain an initial attack angle according to preset ballistic parameters and preset attitude information;
and calculating to obtain the mechanism initialization motion compensation quantity according to the initial attack angle.
3. The split test method for the aircraft according to claim 1, wherein the test model comprises an aircraft and a booster, and the preset constraint formula is as follows:
Figure 932515DEST_PATH_IMAGE001
wherein, | | RK1-RK2| is 2 norm calculation formula, reIn order to constrain the distance, the distance is constrained,
Figure 884290DEST_PATH_IMAGE002
to constrain the angle, phieTo constrain the angle threshold, RK1For the carrier in the first runLet the centroid position at the moment, RK2Is the center of mass position v of the booster at a first preset momentxK1Is the speed of the vehicle in a first direction in the inertial system, vyK1Is the speed, v, of the vehicle in a second direction in the inertial systemxK2Is the speed, v, of the booster in a first direction in the inertial systemyK2Is the speed of the booster in a second direction in the inertial system, theta1Is the attitude angle of the carrier in the inertial system.
4. The aircraft split test method of claim 3, wherein said inputting said test data into a constrained dynamics equation determined from said aerodynamic data:
Figure 622439DEST_PATH_IMAGE003
a first test parameter is calculated, wherein,
Figure 735889DEST_PATH_IMAGE004
=
Figure 926699DEST_PATH_IMAGE005
Figure 416586DEST_PATH_IMAGE006
Figure 759974DEST_PATH_IMAGE007
=
Figure 309904DEST_PATH_IMAGE008
wherein m is1Is the mass of the carrier, /)1The distance from the center of mass of the carrier to the connecting point of the carrier and the booster,（x1，y1) Is the coordinate of the center of mass of the aircraft in the inertial system, J1Moment of inertia being the centre of mass of the aircraft, Fx1Is a generalized active force vector of the first direction of the vehicle in the inertial system, Fy1Is a generalized active force vector of the second direction of the vehicle in the inertial system, M1Is a generalized mass matrix of the carrier, g is the acceleration of gravity,
Figure 784747DEST_PATH_IMAGE009
the angular velocity, m, of the carrier in the inertial system2To the mass of the booster, /)2Is the distance from the center of mass of the booster to the point of attachment of the carrier to the booster, J2Moment of inertia being the centre of mass of the booster, Fx2Is a generalized primary force vector of the booster in a first direction in the inertial system, Fy2Is a generalized primary force vector, θ, of the booster in a second direction in the inertial system2For the attitude angle, M, of the booster in the inertial system2In order to provide a generalized mass matrix for the booster,
Figure 812746DEST_PATH_IMAGE010
to determine the angular velocity of the booster in the inertial system,
Figure 259908DEST_PATH_IMAGE004
in order to be a quality matrix,
Figure 731472DEST_PATH_IMAGE011
in the form of a matrix of accelerations,
Figure 896874DEST_PATH_IMAGE012
in the form of a matrix of aerodynamic forces,
Figure 728564DEST_PATH_IMAGE011
as a matrix of position coordinates
Figure 295811DEST_PATH_IMAGE013
Obtained after two successive derivations.
5. The aircraft split test method of claim 1, wherein said inputting said test data into an unconstrained dynamics equation determined from said aerodynamic data:
Figure 187544DEST_PATH_IMAGE014
calculating to obtain a second test parameter, wherein M1Is an airborne generalized quality matrix and is,
Figure 122133DEST_PATH_IMAGE015
generalized acceleration of the vehicle, F1Is the generalized active force vector of the aircraft, G1Is the gravity vector of the aircraft, M2In order to provide a generalized mass matrix for the booster,
Figure 757514DEST_PATH_IMAGE016
for generalized acceleration of the booster, F2Is the generalized primary force vector of the booster, G2Is the gravity vector of the booster.
6. The aircraft split trial method of claim 1, wherein the first trial parameter comprises a mechanism offset, and the step of inputting the first trial parameter into the mechanism to cause the mechanism to make a position adjustment based on the first trial parameter comprises:
and inputting the mechanism compensation quantity into the mechanism so as to enable the mechanism to carry out position adjustment according to the first test parameter, wherein the mechanism compensation quantity comprises the mass center position, the speed, the attitude angle and the angular speed of the test model at a second preset moment.
7. The utility model provides an aircraft components of a whole that can function independently test device which characterized in that is applied to computer equipment, computer equipment is connected with mechanism, wind-tunnel and test model electricity respectively, the mechanism is used for control test model's gesture, the wind-tunnel is used for providing test model's experimental environment, the device includes:
the generating module is used for generating an initialization signal and sending the initialization signal to the mechanism so that the mechanism can perform initial position setting on the test model according to the initialization signal;
the transmitting module is used for transmitting a starting signal to the wind tunnel so as to start the wind tunnel according to a preset Mach number and acquiring aerodynamic force data and test data of the test model;
the judging module is used for judging whether a preset constraint formula is established or not according to the centroid position of the test model at a first preset moment; if yes, inputting the test data into a constraint kinetic equation determined according to the aerodynamic force data, and calculating to obtain a first test parameter of the test model at a second preset moment; if not, inputting the test data into a non-constrained kinetic equation determined according to the aerodynamic force data, and calculating to obtain a second test parameter of the test model at a second preset time;
the adjusting module is used for inputting the first test parameter or the second test parameter into the mechanism so as to enable the mechanism to adjust the position according to the first test parameter or the second test parameter;
and the test module is used for repeatedly sending the initialization signal to the mechanism and inputting the first test parameter or the second test parameter into the mechanism, and when the preset stop moment is reached, stopping the test and acquiring the target parameter.
8. The aircraft split test apparatus according to claim 7, wherein the initialization signal includes a mechanism initialization motion compensation amount, and the generating module is specifically configured to:
calculating to obtain an initial attack angle according to preset ballistic parameters and preset attitude information; and calculating to obtain the mechanism initialization motion compensation quantity according to the initial attack angle.
9. A computer device comprising a processor and a non-volatile memory storing computer instructions that, when executed by the processor, perform the aircraft split test method of any one of claims 1-6.
10. A readable storage medium, characterized in that the readable storage medium comprises a computer program which, when running, controls a computer device on which the readable storage medium is located to perform the aircraft split test method according to any one of claims 1-6.
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