CN109634306B

CN109634306B - Aircraft control parameter determination method and device

Info

Publication number: CN109634306B
Application number: CN201811625481.XA
Authority: CN
Inventors: 不公告发明人
Original assignee: Beijing Interstellar Glory Space Technology Co Ltd
Current assignee: Beijing Star Glory Space Technology Co Ltd
Priority date: 2018-12-28
Filing date: 2018-12-28
Publication date: 2020-09-08
Anticipated expiration: 2038-12-28
Also published as: CN109634306A

Abstract

The invention relates to the technical field of aircrafts, in particular to a method and a device for determining aircraft control parameters, wherein the method comprises the steps of obtaining interference force and interference torque of at least one preset component; determining flight control parameters according to the disturbance force and the disturbance moment; wherein the flight control parameters include a control force and a control moment. In the process of determining the flight control parameters, the method combines the interference force and the interference moment of the preset component, namely, combines the interference caused by the structural deviation of the aircraft into the determination of the aircraft control parameters, thereby ensuring the accuracy of the determined aircraft control parameters.

Description

Aircraft control parameter determination method and device

Technical Field

The invention relates to the technical field of aircrafts, in particular to a method and a device for determining aircraft control parameters.

Background

In actual flight, an aircraft is often subjected to various disturbance forces and disturbance moments from the aircraft itself and the external environment, and deviates from a predetermined flight state. Therefore, in order to ensure that the aircraft can fly according to the preset track, some parameters of the aircraft need to be controlled. The most important control parameters are control force and control torque.

Disclosure of Invention

In view of this, embodiments of the present invention provide a method and an apparatus for determining aircraft control parameters, so as to solve the problem of determining aircraft control parameters.

According to a first aspect, an embodiment of the present invention provides a method for determining an aircraft control parameter, including:

acquiring interference force and interference torque of at least one preset component;

determining flight control parameters according to the disturbance force and the disturbance moment; wherein the flight control parameters include a control force and a control moment.

According to the method for determining the aircraft control parameters, the interference force and the interference moment of the preset component are combined in the process of determining the flight control parameters, namely, the interference caused by the structural deviation of the aircraft is combined in the determination of the aircraft control parameters, so that the accuracy of the determined aircraft control parameters can be guaranteed.

With reference to the first aspect, in a first implementation manner of the first aspect, the acquiring the interference force of the preset component includes:

acquiring installation parameters of the preset component and an aerodynamic coefficient corresponding to the preset component; wherein the mounting parameters include at least one of a mounting angle and a dihedral angle;

based on the installation parameters, the aerodynamic coefficient is disassembled to determine the disturbance force of the preset component.

According to the method for determining the aircraft control parameters, the aerodynamic coefficient corresponding to the preset component is split by using the installation parameters of the preset component, and the interference force of the preset component is obtained; namely, the mounting angle and the dihedral angle of the preset component are equivalent to the additional attack angle and the sideslip angle of the preset component, the interference force of the preset component is determined through the splitting of the aerodynamic coefficient, and the accuracy of the determined interference force can be guaranteed.

With reference to the first embodiment of the first aspect, in a second embodiment of the first aspect, the acquiring a disturbance torque of the preset component includes:

acquiring the relative position between the center of mass of the aircraft and the pressure center of the preset component;

and determining the disturbance moment of the preset component by using the relative position and the disturbance force of the preset component.

According to the method for determining the control parameters of the aircraft, provided by the embodiment of the invention, the action distance of the interference force can be determined by utilizing the relative relation between the center of mass of the aircraft and the pressure center of the preset component, and the interference force of the preset component is combined, so that the interference torque of the preset component can be determined.

With reference to the first aspect, in a third implementation manner of the first aspect, the determining a flight control parameter according to the disturbance force and the disturbance moment includes:

calculating the mean square sum of the interference force of each preset component and the mean square sum of the interference torque of each preset component, wherein the mean square sum of the interference force of each preset component is the interference force of all the preset components, and the mean square sum of the interference torque of each preset component is the interference torque of all the preset components;

determining the flight control parameters based on the disturbance forces of all the preset components and the disturbance moments of all the preset components; the control force is larger than the interference force of all the preset components, and the control torque is larger than the interference torque of all the preset components.

According to the method for determining the aircraft control parameters, provided by the embodiment of the invention, under the condition that the interference of each preset component obeys normal distribution, the mean square sum of the interference force and the interference moment of each preset component is calculated, so that the interference force and the interference moment caused by the structural deviation can be determined, and the control force and the control moment of the aircraft can be determined.

With reference to the first embodiment of the first aspect, in the fourth embodiment of the first aspect, the preset parts include a cabin section part and a tail wing, and the cabin section part includes at least one of a fairing, an instrument cabin, a stage section, an engine cabin and a tail cabin; the installation parameters of the preset part comprise at least one of a normal installation angle of the cabin section, a lateral installation angle of the cabin section, an installation angle of the tail wing and a dihedral angle of the tail wing; the aerodynamic coefficient includes at least one of a lift line slope and a side force line slope.

With reference to the fourth embodiment of the first aspect, in the fifth embodiment of the first aspect, the splitting the aerodynamic coefficient to determine the disturbance force of the preset component based on the installation parameter includes:

splitting a lifting force line slope corresponding to each cabin section component by using a normal installation angle of each cabin section component so as to determine a normal interference force of the cabin section components;

splitting the gradient of a lifting line of the tail wing by utilizing the mounting angle of the tail wing so as to determine a normal interference force brought by the mounting angle of the tail wing;

splitting the gradient of a lifting line of the tail wing by using the dihedral angle of the tail wing so as to determine a normal interference force brought by the dihedral angle of the tail wing;

using the lateral installation angle of each cabin segment component, disassembling the slope of a lateral force line corresponding to the cabin segment component to determine the lateral interference force of the cabin segment component;

splitting the slope of the lateral force line of the tail wing by using the mounting angle of the tail wing so as to determine the lateral interference force caused by the mounting angle of the tail wing;

and splitting the slope of the lateral force line of the tail wing by using the dihedral angle of the tail wing so as to determine the lateral interference force brought by the dihedral angle of the tail wing.

With reference to the fifth embodiment of the first aspect, in the sixth embodiment of the first aspect, the disturbance torque of the preset component includes a normal disturbance torque of each cabin segment component, a normal disturbance torque due to the tail wing installation angle, a normal disturbance torque due to the tail wing dihedral, a lateral disturbance torque of each cabin segment component, a lateral disturbance torque due to the tail wing installation angle, a lateral disturbance torque due to the tail wing dihedral, a roll disturbance torque due to the tail wing installation angle, and a roll disturbance torque due to the tail wing dihedral.

With reference to the first aspect, or any one of the first to sixth embodiments of the first aspect, in a seventh embodiment of the first aspect, the method further includes:

obtaining interference torque brought by the centroid deviation of the aircraft; wherein the disturbance moment caused by the centroid deviation comprises at least one of a normal disturbance moment caused by the centroid deviation, a lateral disturbance moment caused by the centroid deviation and a roll disturbance moment caused by the centroid deviation;

the determining of the control moment in the flight control parameters according to the disturbance moment comprises: and determining the control moment in the flight control parameters according to the disturbance moment of each preset component and the disturbance moment brought by the mass center deviation.

The method for determining the aircraft control parameters provided by the embodiment of the invention can further improve the accuracy of the determined control torque by combining the disturbance torque brought by the centroid deviation of the aircraft in the determination of the disturbance torque.

According to a second aspect, an embodiment of the present invention further provides an aircraft control parameter determination apparatus, including:

the acquisition module is used for acquiring the interference force and the interference torque of at least one preset component;

the determining module is used for determining flight control parameters according to the interference force and the interference moment; wherein the flight control parameters include a control force and a control moment.

According to the determining device for the aircraft control parameters, provided by the embodiment of the invention, in the process of determining the flight control parameters, the interference force and the interference moment of the preset component are combined, namely, the interference caused by the structural deviation of the aircraft is combined in the determination of the aircraft control parameters, so that the accuracy of the determined aircraft control parameters can be ensured.

According to a third aspect, the present invention provides an electronic device, including a memory and a processor, where the memory and the processor are communicatively connected to each other, the memory stores computer instructions, and the processor executes the computer instructions to perform the method for determining aircraft control parameters according to the first aspect or any implementation manner of the first aspect.

According to a fourth aspect, an embodiment of the present invention provides a computer-readable storage medium storing computer instructions for causing a computer to perform the method for determining aircraft control parameters according to the first aspect or any one of the embodiments of the first aspect.

Drawings

In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.

FIG. 1 is a flow chart of a method of determining aircraft control parameters according to an embodiment of the invention;

FIG. 2 is a flow chart of a method of determining aircraft control parameters according to an embodiment of the invention;

FIG. 3 is a flow chart of a method of determining aircraft control parameters according to an embodiment of the invention;

FIG. 4 is a block diagram of an apparatus for determining aircraft control parameters, in accordance with an embodiment of the present invention;

fig. 5 is a schematic diagram of a hardware structure of an electronic device according to an embodiment of the present invention.

Detailed Description

In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. 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 invention.

For convenience of the following description, technical terms involved in the present invention are explained as follows:

(1) arrow coordinate system: origin O₁Is the center of mass of the rocket; o is₁x₁The shaft points along the longitudinal axis of the arrow body towards the head; o is₁y₁The axis is in the longitudinal symmetry plane of the arrow body and is parallel to O₁x₁The axis is vertical and points to arrow body III quadrant line; o is₁z₁A coordinate system with axes determined by the right hand rule;

(2) aerodynamic coefficient: the aerodynamic force is the force acting on the aircraft when the aircraft and the air move relatively; aerodynamic coefficients include lift coefficients, lateral force coefficients, drag coefficients, and the like.

(3) Lifting force: the component of the pneumatic resultant force vector in the longitudinal symmetrical plane of the aircraft, which is vertical to the direction of the flying speed;

(4) resistance force: a component force of the aerodynamic forces acting on the aircraft in a direction parallel to the flight direction;

(5) normal force: component of resultant aerodynamic force on vertical axis of arrow coordinate system (i.e. along O)₁y₁A component of the axis, or a component in the Y direction);

(6) lateral force: the pneumatic resultant force is on the longitudinal axis of the arrow coordinate systemComponent (i.e., along O)₁z₁A component of the axis, or a component in the Z-direction);

(7) lift coefficient: the ratio of the lift experienced by the object to the product of the aerodynamic pressure and the reference area;

(8) coefficient of lateral force: the ratio of the lateral force experienced by the object to the product of the dynamic airflow pressure and the reference area;

(9) slope of the lifting line: lift coefficient generated by unit angle of attack;

(10) side force line slope: the lateral force coefficient generated per unit sideslip angle;

(11) ballistic: rocket mass center flight motion trail;

(12) pressing the core: namely the pressure center, the action point of aerodynamic resultant;

(13) dynamic pressure q: the kinetic energy per unit volume of fluid, varying along the trajectory;

(14) reference area s: also called characteristic area, which plays a reference role in pneumatic analysis, representative geometric area (in the embodiment of the present invention, the arrow body cross-sectional area is taken as an example);

(15) angle of attack α, also known as angle of attack, with velocity vector in rocket longitudinal plane O₁x₁y₁Inner projection and rocket O₁x₁The included angle between the axes;

(16) sideslip angle β rocket velocity vector in rocket longitudinal plane O₁x₁z₁Inner projection and rocket O₁x₁The angle between the axes.

The inventor finds that in the process of researching an aircraft (taking a carrier rocket as an example), the carrier rocket generally adopts a swinging nozzle control scheme, and because the control force provided by the swinging nozzle is large, the interference caused by structural deviation is not considered when the carrier rocket is designed in the scheme. In addition, when the structure of the carrier rocket is designed, the allowed precision deviation of the structure is given according to the past model experience and the precision level which can be reached by the structure production, and the precision deviation is used as the basis for the distribution of the structure size chain. Further, for small-sized launch vehicles, in order to reduce the cost of the small-sized launch vehicle, a control scheme of an aerodynamic control surface and an attitude control nozzle is generally adopted. In order to reduce drag and reduce the dry weight of the structure in order to improve carrying capacity, the aerodynamic control surface and the attitude control nozzle are preferably smaller, and the control force provided by the aerodynamic control surface and the attitude control nozzle is correspondingly reduced. Therefore, the interference caused by the self structure of the aircraft needs to be considered in the determination process of the control parameters, and on the basis, the invention provides a method for determining the control parameters of the aircraft, wherein the interference force and the interference moment of a preset component are combined when the flight control parameters are determined; the method for determining the disturbance force and the disturbance moment of the preset component is described in detail, and a theoretical basis is provided for determining reasonable aircraft control parameters.

The aircraft in the embodiment of the invention may be a launch vehicle, or may be various missiles (ballistic missiles, air-to-air missiles, air-to-ground missiles, rocket missiles), space-exploring rockets, and the like, and the following description will describe the launch vehicle as an example in detail.

In accordance with an embodiment of the present invention, there is provided an aircraft control parameter determination method embodiment, it being noted that the steps illustrated in the flowchart of the drawings may be performed in a computer system, such as a set of computer-executable instructions, and that while a logical order is illustrated in the flowchart, in some cases the steps illustrated or described may be performed in an order different than presented herein.

In the present embodiment, a method for determining an aircraft control parameter is provided, which may be used in an electronic device, and fig. 1 is a flowchart of a method for determining an aircraft control parameter according to an embodiment of the present invention, as shown in fig. 1, where the flowchart includes the following steps:

and S11, acquiring the interference force and the interference torque of at least one preset component.

The predetermined components include the structure of the aircraft, such as the cowling, tail tanks and empennages, etc. The electronic equipment can directly acquire the interference force and the interference torque of at least one preset component from the outside, and can also calculate the corresponding interference torque by using the interference force of the preset component; or the interference force and the interference torque of the preset component are stored in the electronic device, and when the flight control parameters need to be determined, the interference force and the interference torque can be directly extracted from the memory. The acquisition mode of the interference force and the interference torque is not limited, and only the interference force and the interference torque of the preset component can be acquired by the electronic equipment.

The interference force of the preset component may be calculated by using the installation parameters of the preset component, or determined according to an empirical value. In the embodiment shown below, the calculation of the disturbance force of the preset member using the installation parameters of the preset member will be described in detail. In addition, it should be noted that the number of the preset components and the specific preset components may be specifically set according to the design requirements of the control parameters of the actual aircraft, and are not limited herein.

And S12, determining flight control parameters according to the disturbance force and the disturbance moment.

Wherein the flight control parameters include a control force and a control moment. After the interference force and the interference moment of each preset component are obtained, the electronic equipment can determine the flight control force and the flight control moment by combining the interference force and the interference moment brought by environmental factors. Specifically, the determined control force includes disturbance force and disturbance moment caused by the structure of the aircraft and disturbance force and disturbance moment caused by environmental factors.

The control force of the aircraft can be multiplied by a preset multiple on the basis of the interference force of the preset component, or the interference force of the preset component is combined with the interference force caused by environmental factors, and the determined control force of the aircraft is larger than the combined interference force of the preset component and the environmental factors; accordingly, the determination method of the control torque may be the same as the determination method of the control force, i.e. the acting distance of the corresponding acting force is multiplied on the basis of the control force.

According to the method for determining the aircraft control parameters, in the process of determining the flight control parameters, the disturbance force and the disturbance moment of the preset component are combined, that is, the disturbance caused by the structural deviation of the aircraft is combined in the determination of the aircraft control parameters, so that the accuracy of the determined aircraft control parameters can be ensured.

In this embodiment, a method for determining an aircraft control parameter is further provided, and may be used in an electronic device, fig. 2 is a flowchart of a method for determining an aircraft control parameter according to an embodiment of the present invention, as shown in fig. 2, where the flowchart includes the following steps:

and S21, acquiring the interference force and the interference torque of at least one preset component.

In this embodiment, the electronic device calculates the disturbance force and the disturbance moment of the preset component using the installation parameters of the preset component. That is, the electronic device needs to acquire the installation parameters of the preset component, and the installation parameters may be acquired from the outside or may be stored in advance in the memory of the electronic device.

And S211, acquiring installation parameters of preset components and aerodynamic coefficients corresponding to the preset components.

Wherein the mounting parameter includes at least one of a mounting angle and a dihedral angle. Specifically, the installation angle is the installation angle of the structural component of each cabin section, and the dihedral angle is the composition of the reflection and the transverse movement of the tail wing on the stabilizer. The aerodynamic coefficients correspond to preset components one to one, and different preset components correspond to different aerodynamic coefficients.

And S212, based on the installation parameters, disassembling the aerodynamic coefficient to determine the interference force of the preset part.

The aerodynamic force parameter of the predetermined element is the force acting on the predetermined element, which is generated when the aircraft moves relative to the air, and therefore, the aerodynamic force parameter is related to the angle of action between the air and the aircraft, which mainly includes the angle of attack α and the angle of sideslip β, in addition to being related to environmental factors. Further, due to the predetermined component being mounted, there is an angle between the predetermined component and the aircraft body, for example, a mounting angle, a dihedral angle, etc. Then, the action angle between the air and the aircraft includes an attack angle α and a sideslip angle β, a certain angle exists between the preset component and the aircraft body, and the aerodynamic parameters can be split by using the installation parameters of the preset component (which is equivalent to the fact that the installation parameters of the preset component are equivalent to the additional attack angle and the sideslip angle of each preset component), so that the action force acting on each preset component is split, and the action force can be equivalent to an interference force.

Therefore, the electronic equipment splits the corresponding aerodynamic coefficient by using the installation parameters of the preset components, and the interference force of the preset components can be determined.

And S213, acquiring the relative position between the center of mass of the aircraft and the pressure center of the preset component.

The distance between the center of mass of the aircraft and the pressure center of the preset component is the acting distance of the interference force corresponding to the preset component, and the subsequent electronic equipment can determine the interference torque of the preset component by utilizing the relative position relation.

S214, determining the interference torque of the preset component by using the relative position and the interference force of the preset component.

In particular, the disturbance moment of the preset component is the product of the disturbance force and the relative position of the preset component. The electronic device multiplies the interference force of the preset component by the relative position (distance) to obtain the interference torque of the preset component.

And S22, determining flight control parameters according to the disturbance force and the disturbance moment.

Wherein the flight control parameters include a control force and a control moment. The determination of the control force and the control torque may comprise the steps of:

and S221, calculating the mean square sum of the interference force of each preset component and the mean square sum of the interference moment of each preset component.

Specifically, under the condition that the interference of each preset part is subjected to normal distribution, the sum of the mean square of the interference force and the interference moment of each preset part is calculated, so that the interference force and the interference moment caused by the structural deviation of the aircraft can be determined, and the control force and the control moment of the aircraft can be determined. The calculated mean square sum of the interference force of each preset component is the interference force of all the preset components, and the calculated mean square sum of the interference torque of each preset component is the interference torque of all the preset components.

For example, in the process of determining the interference force caused by the structural deviation of the aircraft, the interference force caused by the structural deviation of 4 preset components (A, B, C, D) is obtained by respectively utilizing the installation parameters of the preset components A to D and splitting aerodynamic coefficients corresponding to the preset components A to D; and then, the interference force caused by the structural deviation of the aircraft is obtained by the mean square sum of the interference forces from the preset part A to the preset part D. Correspondingly, the disturbance moments caused by the structural deviations of the aircraft can be obtained in the same way of calculation (mean square sum).

S222, determining flight control parameters based on the interference force of all the preset components and the interference torque of all the preset components.

The control force is larger than the interference force of all the preset components, and the control torque is larger than the interference torque of all the preset components.

Compared with the embodiment shown in fig. 1, in the method for determining the aircraft control parameter provided by this embodiment, the aerodynamic coefficient corresponding to the preset component is split by using the installation parameter of the preset component, so as to obtain the disturbance force of the preset component; namely, the mounting angle and the dihedral angle of the preset component are equivalent to the additional attack angle and the sideslip angle of the preset component, the interference force of the preset component is determined by splitting the aerodynamic coefficient, and the accuracy of the determined interference force can be ensured; in addition, the action distance of the interference force can be determined by utilizing the relative relation between the center of mass of the aircraft and the pressure center of the preset component, and the interference force of the preset component is combined, so that the interference torque of the preset component can be determined.

In this embodiment, a method for determining an aircraft control parameter is further provided, and may be used in an electronic device, fig. 3 is a flowchart of a method for determining an aircraft control parameter according to an embodiment of the present invention, as shown in fig. 3, where the flowchart includes the following steps:

and S31, acquiring the interference force and the interference torque of at least one preset component.

Specifically, the preset component comprises a cabin section component and an empennage, and the cabin section component comprises at least one of a fairing, an instrument cabin, a stage section, an engine cabin and a tail cabin.

And S311, acquiring the installation parameters of the preset components and the aerodynamic coefficients corresponding to the preset components.

Wherein the mounting parameter includes at least one of a mounting angle and a dihedral angle. The installation parameters of the preset part comprise at least one of a normal installation angle of the cabin section, a lateral installation angle of the cabin section, an installation angle of the tail wing and a dihedral angle of the tail wing. The aerodynamic coefficient includes at least one of a lift line slope and a side force line slope.

And S312, based on the installation parameters, disassembling the aerodynamic coefficient to determine the interference force of the preset component.

Wherein the interference force of the preset member includes: normal disturbance force, lateral disturbance force. When the normal disturbance force is determined, calculation is performed for each of the cabin components and the tail fin. Specifically, for each bay component, the normal incidence angle of each bay component is utilized to break down the lift line slope corresponding to the bay component to determine the normal disturbance force of the bay component. In the case of the tail wing, the normal disturbance force corresponding to the tail wing includes a normal disturbance force due to the installation angle of the tail wing, and a normal disturbance force due to the dihedral angle of the tail wing; namely, the mounting angle of the empennage is utilized to split the gradient of the lifting line of the empennage so as to determine the normal interference force brought by the mounting angle of the empennage; and (3) utilizing the dihedral angle of the tail wing to divide the slope of the lifting line of the tail wing so as to determine the normal interference force caused by the dihedral angle of the tail wing.

Further, when determining the lateral disturbance force, calculation is performed for each of the cabin components and the tail wing. Specifically, for each of the deck section components, the lateral force line slope corresponding to the deck section component is split using the lateral stagger angle of each of the deck section components to determine the lateral disturbance force of the deck section component. In the case of the tail wing, the lateral disturbance force corresponding to the tail wing includes a lateral disturbance force due to the installation angle of the tail wing and a lateral disturbance force due to the dihedral angle of the tail wing; namely, the mounting angle of the empennage is utilized to split the slope of the lateral force line of the empennage so as to determine the lateral interference force caused by the mounting angle of the empennage; and (3) utilizing the dihedral angle of the tail wing to divide the slope of the lateral force line of the split tail wing so as to determine the lateral disturbance force caused by the dihedral angle of the tail wing.

The manner in which the specific normal and lateral disturbance forces are calculated will be described in detail below.

And S313, acquiring the relative position between the center of mass of the aircraft and the pressure center of the preset component.

Please refer to S213 in fig. 2 for details, which are not described herein.

And S314, determining the interference torque of the preset component by using the relative position and the interference force of the preset component.

Wherein, predetermine the disturbance moment of part and include: normal disturbance torque, lateral disturbance torque, and roll disturbance torque. The device comprises a normal disturbance moment of each cabin section part, a normal disturbance moment brought by an empennage installation angle, a normal disturbance moment brought by an empennage dihedral angle, a lateral disturbance moment of each cabin section part, a lateral disturbance moment brought by the empennage installation angle, a lateral disturbance moment brought by the empennage dihedral angle, a rolling disturbance moment brought by the empennage installation angle and a rolling disturbance moment brought by the empennage dihedral angle.

Taking a carrier rocket as an example, the normal interference force and the normal interference moment aiming at the preset part are adopted; lateral disturbance force, lateral disturbance moment; and roll torque are described in detail:

(1) normal disturbance force and normal disturbance moment

a) Fairing

ΔM_{z_1}＝ΔF_{y_1}(x_c-x_{q_1})

Wherein, Δ F_{y_1}In order to provide normal disturbing forces to the fairing,

is the slope of the lifting line of the fairing, q is the dynamic pressure, s is the reference area,

is normal to the fairingMounting angle, Δ M_{z_1}For normal disturbance moment of fairing, x_cIs the centroid position of the whole arrow along the X direction, X_{q_} ¹The position of the pressure center of the fairing along the X direction.

b) Instrument cabin

ΔM_{z_2}＝ΔF_{y_2}(x_c-x_{q_2})

Wherein, Δ F_{y_2}In order to disturb the force in the normal direction of the instrument chamber,

is the slope of the instrument pod lift line,

is the normal installation angle of an instrument chamber, delta M_{z_2}For normal disturbance moment, x, of the instrument chamber_{q_2}The instrument chamber is in the pressure center position along the X direction.

c) Stage section

ΔM_{z_3}＝ΔF_{y_3}(x_c-x_{q_3})

Wherein, Δ F_{y_3}In order to provide a normal disturbance force for the stage section,

for the slope of the lift line of the stage interval,

is a normal mounting angle of the stage section, Δ M_{z_3}For normal disturbance moment, x, of the stage section_{q_3}The position of the core of the stage section along the X direction.

d) Engine compartment

ΔM_{z_4}＝ΔF_{y_4}(x_c-x_{q_4})

Wherein, Δ F_{y_4}In order to interfere with the force in the normal direction of the engine compartment,

is the slope of the engine compartment lift line,

is the normal installation angle of the engine compartment, Δ M_{z_4}For engine compartment normal disturbance moment, x_{q_4}The pressure center position of the engine compartment along the X direction.

e) Tail cabin

ΔM_{z_5}＝ΔF_{y_5}(x_c-x_{q_5})

Wherein, Δ F_{y_5}Is the normal interference force of the tail cabin,

is the slope of the tail tank lifting line,

is the normal installation angle of the tail cabin, delta M_{z_5}For normal disturbance moment of the tail tank, x_{q_5}The core pressing position of the tail cabin along the X direction.

f) Single-piece empennage installation angle (taking X-shaped empennage as an example)

ΔM_{z_W}＝ΔF_{y_W}(x_c-x_{q_W})

Wherein, Δ F_{y_W}The normal interference force brought by the installation angle of the tail wing,

is the slope of the lift line of the tail wing, delta phi is the mounting angle of the tail wing, and delta M_{z_W}Normal disturbance moment, x, for the mounting angle of the tail_{q_W}The tail wing is in the pressing core position along the X direction.

g) Single tail dihedral angle and sideslip (taking X-type tail as an example)

ΔM_{z_WF}＝ΔF_{y_WF}(x_c-x_{q_w})

Wherein, Δ F_{y_WF}Normal disturbance force, Δ M, for the dihedral angle of the tail_{z_WF}For normal disturbance moment brought by the dihedral angle of the tail wing, delta psi is the dihedral angle of the stabilizer combined with the sideslip, and the angle of attack α is the actual angle of attack along the trajectory.

(2) Lateral disturbance force and lateral disturbance moment

a) Fairing

ΔM_{y_1}＝ΔF_{z_1}(x_c-x_{q_1})

Wherein, Δ F_{z_1}In order to provide a lateral disturbing force to the fairing,

is the slope of the fairing side force line, q is the dynamic pressure, s is the reference area,

for fairing side-mounted angle, Δ M_{y_1}For a fairing side disturbance moment, x_cIs the center of mass, X, of the whole arrow along the X direction_{q_1}The position of the pressure center of the fairing along the X direction.

b) Instrument cabin

ΔM_{y_2}＝ΔF_{z_2}(x_c-x_{q_2})

Wherein, Δ F_{z_2}In order to interfere with the force laterally of the instrument pod,

is the slope of the force line on the side of the instrument pod,

for the lateral setting angle of the instrument chamber, Δ M_{y_2}For side disturbing moment, x, of the instrument chamber_{q_2}The instrument chamber is in the pressure center position along the X direction.

c) Stage section

ΔM_{y_3}＝ΔF_{z_3}(x_c-x_{q_3})

Wherein, Δ F_{z_3}In order to provide a level section side interference force,

for the slope of the side force line of the stage section,

is a lateral mounting angle of the stage section, Δ M_{y_3}For step-interval side disturbance moment, x_{q_3}The position of the core of the stage section along the X direction.

d) Engine compartment

ΔM_{y_4}＝ΔF_{z_4}(x_c-x_{q_4})

Wherein, Δ F_{z_4}In order to provide a side-to-side disturbing force to the engine compartment,

is the slope of the engine compartment side force line,

for engine nacelle side-mounting angle, Δ M_{y_4}For engine nacelle side disturbing moment, x_{q_4}The pressure center position of the engine compartment along the X direction.

e) Tail cabin

ΔM_{y_5}＝ΔF_{z_5}(x_c-x_{q_5})

Wherein, Δ F_{z_5}Is the side interference force of the tail cabin,

is the slope of the force line on the lateral side of the tail tank,

is a lateral installation angle of the tail cabin, delta M_{y_5}For side disturbing moment of stern compartment, x_{q_5}The core pressing position of the tail cabin along the X direction.

ΔM_{y_W}＝ΔF_{z_W}(x_c-x_{q_W})

Wherein, Δ F_{z_W}The side interference force brought by the installation angle of the tail wing,

is the slope of the lift line of the tail wing, delta phi is the mounting angle of the tail wing, and delta M_{y_W}Side disturbance moment, x, for the mounting angle of the tail_{q_W}The tail wing is in the pressing core position along the X direction.

g) Single tail dihedral angle and sideslip (taking X-type tail as an example)

ΔM_{y_WF}＝ΔF_{z_WF}(x_c-x_{q_w})

Wherein, Δ F_{z_WF}Side interference, Δ M, for the dihedral of the tail_{y_WF}For the side disturbance moment brought by the dihedral of the tail wing, Δ ψ is the combined dihedral of the stabilizer and the sideslip, and the sideslip angle β is the actual sideslip angle along the trajectory.

(3) Rolling disturbance torque

a) Single-piece empennage installation angle (taking X-shaped empennage as an example)

ΔM_{x_W}＝ΔF_{z_W}(y_c-y_{q_W})+ΔF_{y_W}(z_c-z_{q_W})

Wherein, Δ M_{x_W}Roll disturbance moment y of the single-flight tail angle_c、z_cThe centroid of the whole arrow along the y-direction and the z-direction, respectively, y_{q_W}、z_{q_W}The tail wing pressure center positions along the Y direction and the Z direction respectively, wherein the reasonable interference force direction can be selected to enable the tail wings to generate the moment in the same direction.

b) Single tail upper reverse and lateral shift angle (taking X-type tail as an example)

ΔM_{x_WF}＝ΔF_{z_WF}(y_c-y_{q_W})+ΔF_{y_WF}(z_c-z_{q_W})

Wherein, Δ M_{x_WF}The roll interference torque brought by the dihedral angle of the single-chip empennage can be selected from reasonable interference force directions so that the empennage generates torque in the same direction.

It should be noted that, for the normal disturbance force and disturbance moment caused by the installation angle of the tail wing, the lateral disturbance force and disturbance moment caused by the installation angle of the tail wing, the normal disturbance force and disturbance moment caused by the dihedral angle of the tail wing, and the lateral disturbance force and disturbance moment caused by the dihedral angle of the tail wing, all the individual tail wings are used, and for different aircraft, the disturbance force and the disturbance moment corresponding to all the tail wings can be obtained by multiplying the disturbance force and the disturbance moment corresponding to the individual tail wing by the number of the tail wings according to the specific number of the tail wings.

And S32, acquiring disturbance torque brought by the mass center deviation of the aircraft.

Wherein the disturbance torque caused by the centroid deviation comprises at least one of a normal disturbance torque, a lateral disturbance torque and a roll disturbance torque caused by the centroid deviation.

Specifically, it is expressed as follows:

(1) normal disturbance moment delta M brought by centroid deviation_{z_c}

Moment of aerodynamic force of each main part due to mass center deviation:

where angle of attack α is the actual angle of attack, Δ M, along the trajectory_{z_c}The normal disturbance moment brought by the deviation of the mass center,

is the slope of the full arrow lift line, C_xIs the drag coefficient of the whole arrow, Δ y_cIs the centroid deviation of the whole arrow normal direction (Y direction).

(2) Side interference moment delta M caused by centroid deviation_{y_c}

The moment caused by the center of mass deviation of aerodynamic force of each main part and the thrust of the solid rocket engine is considered:

where the side slip angle β is the actual side slip angle along the trajectory, Δ M_{y_c}The side interference moment caused by the deviation of the mass center,

is the coefficient of lift of the whole arrow, C_xIs the coefficient of drag of the whole arrow, Δ z_cIs the lateral (Z-direction) centroid deviation of the whole arrow.

(3) Roll interference torque brought by centroid deviation

Wherein, Δ M_{x_c}The roll disturbance moment brought by the centroid deviation, wherein, the reasonable disturbance force direction can be selected to lead the empennage to generate the moment in the same direction.

And S33, determining flight control parameters according to the disturbance force and the disturbance moment.

Wherein the flight control parameters include a control force and a control moment.

And S331, calculating the mean square sum of the interference force of each preset component, the interference moment of each preset component and the mean square sum of the interference moment brought by the mass center deviation.

Under the condition that the interference of each preset part is subjected to normal distribution, the sum of the mean square of the interference force, the interference moment and the interference moment brought by the centroid deviation of each preset part is calculated, the interference force and the interference moment caused by the structural deviation can be determined, and therefore the control force and the control moment of the aircraft can be determined.

The mean square sum of the interference force of each preset component is the interference force of all the preset components, and the mean square sum of the interference moment of each preset component and the interference moment brought by the mass center deviation is the interference moment brought by all the preset components and the mass center deviation.

Specifically, the normal disturbance force and the normal disturbance torque of the whole rocket (taking a cabin and a single-piece tail as examples) are root mean square values of the various normal disturbance forces and normal disturbance torques, that is:

wherein, Δ F_yIs the normal interference force of the whole arrow.

Wherein, Δ M_zIs the normal disturbance moment of the whole arrow.

(2) For the whole arrow, the lateral disturbance force and the lateral disturbance moment are root mean square values of the various lateral disturbance forces and the lateral disturbance moments, namely:

wherein, Δ F_zIs the lateral interference force of the whole arrow.

Wherein, Δ M_yIs the lateral disturbance moment of the whole arrow.

(3) The roll disturbance torque of the whole rocket is the root mean square value of the various roll disturbance torques, namely:

wherein, Δ M_xThe torque is the roll disturbance torque of the whole arrow.

It should be noted that, in the practical application process, when the total rocket interference force and the interference torque are determined, specific calculation is performed in combination with the number of specific cabin segments.

S332, determining flight control parameters based on the interference force of all the preset components and the interference torque of all the preset components.

Specifically, the determined control force of the aircraft is expressed as: f>ΔF_y+ΔF_z；

The determined control moment of the aircraft is expressed as: m>ΔM_z+ΔM_y+ΔM_x。

In this embodiment, a device for determining aircraft control parameters is further provided, and the device is used to implement the foregoing embodiments and preferred embodiments, which have already been described and are not described again. As used below, the term "module" may be a combination of software and/or hardware that implements a predetermined function. Although the means described in the embodiments below are preferably implemented in software, an implementation in hardware, or a combination of software and hardware is also possible and contemplated.

The present embodiment provides an aircraft control parameter determination device, as shown in fig. 4, including:

an obtaining module 41 is configured to obtain a disturbance force and a disturbance torque of at least one preset component.

A determination module 42, configured to determine a flight control parameter according to the disturbance force and the disturbance moment; wherein the flight control parameters include a control force and a control moment.

The device for determining aircraft control parameters provided by this embodiment combines the disturbance force and the disturbance moment of the preset component in the process of determining the flight control parameters, that is, combines the disturbance caused by the structural deviation of the aircraft into the determination of the aircraft control parameters, thereby being capable of ensuring the accuracy of the determined aircraft control parameters.

The means for determining aircraft control parameters in this embodiment are in the form of functional units, where a unit refers to an ASIC circuit, a processor and memory executing one or more software or fixed programs, and/or other devices that may provide the above-described functionality.

Further functional descriptions of the modules are the same as those of the corresponding embodiments, and are not repeated herein.

An embodiment of the present invention further provides an electronic device, which has the aircraft control parameter determination apparatus shown in fig. 5.

Referring to fig. 5, fig. 5 is a schematic structural diagram of an electronic device according to an alternative embodiment of the present invention, and as shown in fig. 5, the terminal may include: at least one processor 51, such as a CPU (Central Processing Unit), at least one communication interface 53, memory 54, at least one communication bus 52. Wherein a communication bus 52 is used to enable the connection communication between these components. The communication interface 53 may include a Display (Display) and a Keyboard (Keyboard), and the optional communication interface 53 may also include a standard wired interface and a standard wireless interface. The Memory 54 may be a high-speed RAM Memory (volatile Random Access Memory) or a non-volatile Memory (non-volatile Memory), such as at least one disk Memory. The memory 54 may alternatively be at least one memory device located remotely from the processor 51. Wherein the processor 51 may be in connection with the apparatus described in fig. 4, the memory 54 stores an application program, and the processor 51 calls the program code stored in the memory 54 for performing any of the above-mentioned method steps.

The communication bus 52 may be a Peripheral Component Interconnect (PCI) bus or an Extended Industry Standard Architecture (EISA) bus. The communication bus 52 may be divided into an address bus, a data bus, a control bus, and the like. For ease of illustration, only one thick line is shown in FIG. 5, but this is not intended to represent only one bus or type of bus.

The memory 54 may include a volatile memory (RAM), such as a random-access memory (RAM); the memory may also include a non-volatile memory (english: non-volatile memory), such as a flash memory (english: flash memory), a hard disk (english: hard disk drive, abbreviation: HDD), or a solid-state drive (english: SSD); the memory 54 may also comprise a combination of the above types of memories.

The processor 51 may be a Central Processing Unit (CPU), a Network Processor (NP), or a combination of a CPU and an NP.

The processor 51 may further include a hardware chip. The hardware chip may be an application-specific integrated circuit (ASIC), a Programmable Logic Device (PLD), or a combination thereof. The aforementioned PLD may be a Complex Programmable Logic Device (CPLD), a field-programmable gate array (FPGA), a General Array Logic (GAL), or any combination thereof.

Optionally, the memory 54 is also used to store program instructions. Processor 51 may invoke program instructions to implement the method of determining aircraft control parameters as shown in the embodiments of fig. 1-3 of the present application.

Embodiments of the present invention further provide a non-transitory computer storage medium, where the computer storage medium stores computer-executable instructions, and the computer-executable instructions may execute the method for determining aircraft control parameters in any of the above method embodiments. The storage medium may be a magnetic Disk, an optical Disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a Flash Memory (Flash Memory), a Hard Disk (Hard Disk Drive, abbreviated as HDD), a Solid State Drive (SSD), or the like; the storage medium may also comprise a combination of memories of the kind described above.

Although the embodiments of the present invention have been described in conjunction with the accompanying drawings, those skilled in the art may make various modifications and variations without departing from the spirit and scope of the invention, and such modifications and variations fall within the scope defined by the appended claims.

Claims

1. A method of determining aircraft control parameters, comprising:

determining flight control parameters according to the disturbance force and the disturbance moment; wherein the flight control parameters include a control force and a control moment;

wherein, the interference force of obtaining the preset component comprises:

acquiring installation parameters of the preset component and an aerodynamic coefficient corresponding to the preset component; wherein the mounting parameters include at least one of a mounting angle and a dihedral angle; the preset components comprise a cabin section component and an empennage, and the cabin section component comprises at least one of a fairing, an instrument cabin, a stage section, an engine cabin and a tail cabin; the installation parameters of the preset parts comprise at least one of a normal installation angle of the cabin part, a lateral installation angle of the cabin part, an installation angle of the tail wing and a dihedral angle of the tail wing; the aerodynamic coefficient comprises at least one of a lift line slope and a side force line slope;

splitting the aerodynamic coefficient based on the installation parameters to determine the disturbance force of the preset component;

the disassembling the aerodynamic coefficient based on the installation parameters to determine the disturbance force of the preset component comprises:

splitting the slope of the lateral force line of the tail wing by using the dihedral angle of the tail wing to determine the lateral disturbance force caused by the dihedral angle of the tail wing;

determining flight control parameters according to the disturbance force and the disturbance moment, wherein the flight control parameters comprise:

2. The method of claim 1, wherein the obtaining of the disturbance torque of the predetermined component comprises:

3. The method of claim 2, wherein the disturbance torque of the predetermined component includes a normal disturbance torque of each of the cabin components, a normal disturbance torque of the tail angle, a normal disturbance torque of the tail dihedral, a lateral disturbance torque of each of the cabin components, a lateral disturbance torque of the tail angle, a lateral disturbance torque of the tail dihedral, a roll disturbance torque of the tail angle, and a roll disturbance torque of the tail dihedral.

4. The method according to any one of claims 1-3, further comprising:

5. An electronic device, comprising:

a memory and a processor, the memory and the processor being communicatively coupled to each other, the memory having stored therein computer instructions, the processor executing the computer instructions to perform the method of determining aircraft control parameters of any of claims 1-4.

6. A computer-readable storage medium storing computer instructions for causing a computer to perform the method for determining aircraft control parameters of any of claims 1-4.