CN112364432B - Control method for carrier hanging and throwing separation process - Google Patents
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Abstract
The invention provides a control method for a flying and throwing separation process of a carrier, which comprises the steps of firstly obtaining overall parameters and throwing separation parameters of an aircraft as design data, calculating initial values of rudder deflection angles and start control time required for keeping stable separation postures, calculating and determining a posture angle instruction value in the separation process, then verifying the initial values of the rudder deflection angles and the start control time through separation dynamic track simulation calculation, designing a posture angle control law according to a classical control method, forming a throwing separation control scheme, and finally verifying the scheme effectiveness through Monte Carlo simulation. Compared with the prior art, the aircraft launching separation safety control method has the advantages that in the separation control design, aiming at the pneumatic interference of the aircraft carrier in the existing plane symmetry aircraft separation process, the initial control plane deflection angle is increased, the aircraft launching separation safety problem with the airfoil is effectively solved, meanwhile, the fastest separation attitude angle instruction is increased in the separation process, the risk of collision with the aircraft carrier is greatly reduced, and the launching separation safety is improved.
Description
Technical Field
The invention relates to a control method for a carrier hanging and throwing separation process, which is applicable to the field of hanging and throwing control.
Background
The unpowered world round-trip plane symmetrical aircraft is an important field of high-tech development of various large countries of the world in the present stage, and through continuous efforts of various large countries of the world, the plane symmetrical round-trip plane aircraft makes great progress on key technical verification, but any technology forming process of product put into use is explored through tortuous scientific technology, and before development of the space aircraft, a demonstration verification test of overhanging and putting in is carried out by adopting a carrier such as a white knight and the like in the United states so as to fully verify the key technology. For an unpowered world round-trip plane symmetrical aircraft, hanging and throwing are a common key technology flight verification test method.
As with external missile delivery, the problem of safety in separating a facing aircraft from a carrier is also a key problem and a difficult problem in separating the aircraft from the carrier, wherein the separation of the aircraft from the carrier is guided not to exceed the design limit of the missile or the carrier or other airborne objects, and the carrier, the suspension device or other suspended objects cannot be damaged or collided, or adverse side effects are generated on the carrier, the suspension device or other suspended objects. Unlike axisymmetric missiles, however, face symmetric aircraft suffer from more severe carrier aerodynamic interference problems, which are more pronounced with face symmetric aircraft that rely on aerodynamic effect face flight and control, and therefore effective flight control methods within the separation zone are particularly important.
Disclosure of Invention
The invention solves the technical problems that: the method for controlling the separation process of the flying and throwing of the carrier solves the problem of safety control of the surface symmetry aircraft on the outside hanging and separation, and provides a proper separation control scheme for the aircraft.
The technical scheme of the invention is as follows: a control method for a carrier hanging and throwing separation process comprises the following steps:
1) Acquiring overall parameters and release separation parameters of the aircraft as design data;
2) According to the pneumatic interference characteristic of the plane symmetrical aircraft of the carrier in the separation area, calculating the deflection angle of the control rudder required for keeping stable attitude, determining the initial deflection angle of the separation rudder, and simultaneously giving an initial value of the start control time according to experience;
3) Calculating the relation between the inclination angle of the separation track and the attack angle, and determining the attitude angle instruction value in the separation process;
4) Carrying out simulation calculation on the separated dynamic track to obtain the relation between the gesture and track parameters before starting control under the conditions of the initial rudder deflection angle and the starting control time determined in the step 2), and confirming that the gesture before starting control is stable and does not collide with the carrier;
5) Designing an attitude angle control law according to a classical control method;
6) And (3) performing Monte Carlo simulation analysis according to the initial rudder deflection angle and the control time determined in the step (2) and the control law designed in the step (5), and controlling the gesture in the separation process to be stable and not collide with the carrier.
The specific process of the step 1) is as follows: the method comprises the steps of obtaining overall parameters of an aircraft and release separation parameters as design data, wherein the overall parameters of the aircraft comprise appearance parameters, quality and inertia characteristic parameters of the aircraft, and the overall parameters comprise aerodynamic data of carrier separation area interference; the launch separation parameters include launch height, velocity, mach number, angle of attack/sideslip angle, attitude angular velocity, while including deviations in the parameters.
The initial value of the starting control time in the step 2) is 1s.
In the step 2), the calculation formula of the deflection angle of the separated initial control surface is as follows:
C l (Ma * ,α * ,β * ,δ a * ,δ e * ,δ r * )＝0
C m (Ma * ,α * ,β * ,δ a * ,δ e * ,δ r * )＝0
C n (Ma * ,α * ,β * ,δ a * ,δ e * ,δ r * )＝0
wherein: c (C) l (.)、C m (.)、C n (.) roll, pitch, yaw moment with respect to Mach number Ma, angle of attack α, sideslip angle β, aileron δ, respectively a Elevator delta e And rudder delta r A functional expression of (2); values of mach number, angle of attack, sideslip angle, elevator, aileron, rudder when roll, pitch and yaw moments are zero, wherein mach number, angle of attack, sideslip angle are given by step 1);
and solving the three equations to obtain values of the aileron, the elevator and the rudder.
The specific calculation method of the step 3) comprises the following steps: the particle dynamics equation for an unpowered aircraft in the wind axis is described as:
the particle dynamics equation for an aircraft in the wind axis is described as:
wherein V is speed, m is mass, g is gravitational acceleration, gamma is track inclination angle, D and L respectively represent aerodynamic drag and lift force acting on the unpowered aircraft, and the method is further rewritten as:
respectively representing aerodynamic drag and lift force acting on an unpowered aircraft, S being a reference area, C L For lift coefficient, C D The resistance coefficient, ma is Mach number, alpha is attack angle, dynamic pressure +.>ρ is the atmospheric density and V is the airspeed of the aircraft; and (3) determining the relation between the track inclination angle and the attack angle according to the separation parameter in the step (1), the overall aircraft parameter and the equation, and taking the corresponding attack angle when the track inclination angle is maximum to separate according to the altitude, the speed, the Mach number and the overall aircraft parameter of the separation parameter, so as to finish the gesture angle instruction design.
In the step 4), a separation dynamic track simulation calculation is carried out, specifically, the initial rudder deflection angle and the starting control time determined in the step 2) are transmitted to a pneumatic specialty, the relation between the gesture and track parameters before starting control is calculated with numerical values, if the calculation shows that the gesture in the separation process diverges or collides with a carrier, the calculation of the initial trimming rudder deflection angle is carried out again, and meanwhile, the starting control time initial value is adjusted.
The specific calculation method of the step 5) comprises the following steps: and (3) carrying out bias pulling on the overall data of the aircraft and the initial separation parameters, carrying out Monte Carlo simulation analysis, if the gesture can be ensured not to diverge under various deviation conditions and the separation process can not interfere with the carrier, ending, obtaining separation start control time after the design is ended, and separating the initial control surface deflection angle and gesture control law to form a separation control scheme, otherwise returning to the step (2), and restarting.
Compared with the prior art, the invention has the advantages that: in the design of separation control, aiming at the aerodynamic interference of the carrier in the existing plane symmetry aircraft separation process, the initial control plane deflection angle is increased, and the problem of safety in the throwing and separation of the aircraft with the airfoil is effectively solved. Meanwhile, the fastest separation attitude angle instruction is added in the separation process, so that the risk of collision with the carrier is greatly reduced, and the release separation safety is improved.
Drawings
FIG. 1 is a flow chart of the method of the present invention.
Detailed Description
As shown in fig. 1, a method for controlling a carrier flying and throwing separation process comprises the following steps:
1. the method comprises the steps of obtaining overall parameters of an aircraft and release separation parameters as design data, wherein the overall parameters of the aircraft comprise appearance parameters, quality and inertia characteristic parameters of the aircraft, and the overall parameters comprise aerodynamic data of carrier separation area interference; the launch separation parameters include launch height, velocity, mach number, angle of attack/sideslip angle, attitude angular velocity, while including deviations in the parameters.
2. This step is mainly to take into account the strong aerodynamic disturbances of the aircraft on the opposite side of the symmetrical aircraft in the separation zone. According to the pneumatic interference characteristics of the aircraft on the opposite symmetry of the separation area, the initial control surface deflection angle required for maintaining stable attitude is calculated, and an initial value of the start control time is given empirically, and is generally 1s. The calculation formula of the initial control surface deflection angle is as follows:
C l (Ma * ,α * ,β * ,δ a * ,δ e * ,δ r * )＝0
C m (Ma * ,α * ,β * ,δ a * ,δ e * ,δ r * )＝0
C n (Ma * ,α * ,β * ,δ a * ,δ e * ,δ r * )＝0
wherein: c (C) l (.)、C m (.)、C n (.) roll, pitch, yaw moment with respect to Mach number Ma, angle of attack α, sideslip angle β, aileron δ, respectively a Elevator delta e And rudder delta r A functional expression of (2); the values of mach number, angle of attack, sideslip angle, elevator, aileron, rudder when the roll, pitch and yaw moments are zero are given by step 1), and the values of aileron, elevator and rudder are solved by three equations.
3. And calculating the relation between the separation track inclination angle and the separation attitude angle instruction, and determining the attitude angle instruction. The specific calculation method comprises the following steps: the particle dynamics equation for an unpowered aircraft in the wind axis is described as:
the particle dynamics equation for an aircraft in the wind axis can be described as:
in the above formula, V is velocity, m is mass, g is gravitational acceleration, γ is track inclination angle, D, L respectively represent aerodynamic drag and lift force acting on an unpowered aircraft, which can be further rewritten as:
respectively representing aerodynamic drag and lift force acting on an unpowered aircraft, S being a reference area, C L For lift coefficient, C D The resistance coefficient, ma is Mach number, alpha is attack angle, dynamic pressure +.>ρ is the atmospheric density and V is the airspeed of the aircraft. According to the separation parameter in the step 1), the overall aircraft parameter and the four equations, the relation between the track inclination angle and the attack angle can be determined according to the height, the speed, the Mach number and the overall aircraft parameter of the separation parameter, and the corresponding attack angle is obtained when the maximum separation track inclination angle is taken, so that the gesture angle instruction design is completed.
4. And (3) carrying out simulation calculation of the separation dynamic track, specifically transmitting the initial rudder deflection angle and the starting control time determined in the step (2) to a pneumatic specialty, carrying out numerical calculation on the relation between the gesture and the track parameter before starting control along with time, and if the calculation shows that the gesture in the separation process diverges or collides with the carrier, carrying out calculation on the initial trimming rudder deflection angle again, and simultaneously adjusting the starting control time initial value.
5. And (3) completing the design of the attitude control law according to a classical engineering control method, wherein the specific design method is a general method in the field.
6. And (3) carrying out bias pulling on the overall data of the aircraft and the initial separation parameters, carrying out Monte Carlo simulation analysis, if the gesture can be ensured not to diverge under various deviation conditions and the separation process can not interfere with the carrier, ending, obtaining separation start control time after the design is ended, and separating the initial control surface deflection angle and gesture control law to form a separation control scheme, otherwise returning to the step (2), and restarting.
The invention is not described in detail in the field of technical personnel common knowledge.
Claims (7)
1. A control method for a carrier hanging and throwing separation process is characterized by comprising the following steps:
1) Acquiring overall parameters and release separation parameters of the aircraft as design data;
2) According to the pneumatic interference characteristic of the plane symmetrical aircraft of the carrier in the separation area, calculating the deflection angle of the control rudder required for keeping stable attitude, determining the initial deflection angle of the separation rudder, and simultaneously giving an initial value of the start control time according to experience;
3) Calculating the relation between the inclination angle of the separation track and the attack angle, and determining the attitude angle instruction value in the separation process;
4) Carrying out simulation calculation on the separated dynamic track to obtain the relation between the gesture and track parameters before starting control under the conditions of the initial rudder deflection angle and the starting control time determined in the step 2), and confirming that the gesture before starting control is stable and does not collide with the carrier;
5) Designing an attitude angle control law according to a classical control method;
6) And (3) performing Monte Carlo simulation analysis according to the initial rudder deflection angle and the control time determined in the step (2) and the control law designed in the step (5), and controlling the gesture in the separation process to be stable and not collide with the carrier.
2. The method for controlling the on-board fly release separation process according to claim 1, wherein the method comprises the following steps: the specific process of the step 1) is as follows: the method comprises the steps of obtaining overall parameters of an aircraft and release separation parameters as design data, wherein the overall parameters of the aircraft comprise appearance parameters, quality and inertia characteristic parameters of the aircraft, and the overall parameters comprise aerodynamic data of carrier separation area interference; the launch separation parameters include launch height, velocity, mach number, angle of attack/sideslip angle, attitude angular velocity, while including deviations in the parameters.
3. The method for controlling the on-board fly release separation process according to claim 1, wherein the method comprises the following steps: the initial value of the starting control time in the step 2) is 1s.
4. The method for controlling the on-board fly release separation process according to claim 1, wherein the method comprises the following steps: in the step 2), the calculation formula of the deflection angle of the separated initial control surface is as follows:
wherein: c (C) l (.)、C m (.)、C n (.) roll, pitch, yaw moment with respect to Mach number Ma, angle of attack α, sideslip angle β, aileron δ, respectively a Elevator delta e And rudder delta r A functional expression of (2); values of mach number, angle of attack, sideslip angle, elevator, aileron, rudder when roll, pitch and yaw moments are zero, wherein mach number, angle of attack, sideslip angle are given by step 1);
and solving the three equations to obtain values of the aileron, the elevator and the rudder.
5. The method for controlling the on-board fly release separation process according to claim 1, wherein the method comprises the following steps: the specific calculation method of the step 3) comprises the following steps: the particle dynamics equation for an unpowered aircraft in the wind axis is described as:
the particle dynamics equation for an aircraft in the wind axis is described as:
wherein V is speed, m is mass, g is gravitational acceleration, gamma is track inclination angle, D and L respectively represent aerodynamic drag and lift force acting on the unpowered aircraft, and the method is further rewritten as:
d, L respectively represent aerodynamic drag and lift force acting on the unpowered aircraft, S is a reference area, C L For lift coefficient, C D The resistance coefficient, ma is Mach number, alpha is attack angle, dynamic pressure +.>ρ is the atmospheric density and V is the airspeed of the aircraft; and (3) determining the relation between the track inclination angle and the attack angle according to the separation parameter in the step (1), the overall aircraft parameter and the equation, and taking the corresponding attack angle when the track inclination angle is maximum to separate according to the altitude, the speed, the Mach number and the overall aircraft parameter of the separation parameter, so as to finish the gesture angle instruction design.
6. The method for controlling the on-board fly release separation process according to claim 1, wherein the method comprises the following steps: in the step 4), a separation dynamic track simulation calculation is carried out, specifically, the initial rudder deflection angle and the starting control time determined in the step 2) are transmitted to a pneumatic specialty, the relation between the gesture and track parameters before starting control is calculated with numerical values, if the calculation shows that the gesture in the separation process diverges or collides with a carrier, the calculation of the initial trimming rudder deflection angle is carried out again, and meanwhile, the starting control time initial value is adjusted.
7. The method for controlling the on-board fly release separation process according to claim 1, wherein the method comprises the following steps: the specific calculation method of the step 5) comprises the following steps: and (3) carrying out bias pulling on the overall data of the aircraft and the initial separation parameters, carrying out Monte Carlo simulation analysis, if the gesture can be ensured not to diverge under various deviation conditions and the separation process can not interfere with the carrier, ending, obtaining separation start control time after the design is ended, and separating the initial control surface deflection angle and gesture control law to form a separation control scheme, otherwise returning to the step (2), and restarting.
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| CN113609594B (en) * | 2021-08-18 | 2022-03-15 | 北京空间飞行器总体设计部 | Method for determining safe separation condition of heat-proof outsole |
| CN114940264B (en) * | 2022-05-27 | 2024-07-12 | 中国人民解放军国防科技大学 | Safe suspension object and carrier separation attitude control method |
| CN116266238B (en) * | 2022-12-28 | 2024-05-03 | 中国航天科工飞航技术研究院（中国航天海鹰机电技术研究院） | Supersonic near-ground parallel interstage separation method with preset rudder deflection characteristics |
| CN118013893B (en) * | 2024-04-08 | 2024-06-18 | 中国空气动力研究与发展中心计算空气动力研究所 | Method and system for judging variant throwing safety separation condition of variant aircraft |
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| IT1392259B1 (en) * | 2008-12-11 | 2012-02-22 | Alenia Aeronautica Spa | ESTIMATION PROCEDURE OF THE INCIDENCE ANGLE AND THE DERAPATE CORNER OF AN AIRCRAFT |
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