CN108931987B - Attitude control system design method - Google Patents
Attitude control system design method Download PDFInfo
- Publication number
- CN108931987B CN108931987B CN201810743622.1A CN201810743622A CN108931987B CN 108931987 B CN108931987 B CN 108931987B CN 201810743622 A CN201810743622 A CN 201810743622A CN 108931987 B CN108931987 B CN 108931987B
- Authority
- CN
- China
- Prior art keywords
- initial value
- mass
- control system
- propellant
- attitude control
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000000034 method Methods 0.000 title claims abstract description 33
- 238000013461 design Methods 0.000 title claims abstract description 24
- 230000002068 genetic effect Effects 0.000 claims abstract description 20
- 239000003380 propellant Substances 0.000 claims description 69
- 238000002485 combustion reaction Methods 0.000 claims description 28
- 238000003860 storage Methods 0.000 claims description 16
- 238000004364 calculation method Methods 0.000 claims description 15
- 239000000446 fuel Substances 0.000 claims description 8
- 230000008602 contraction Effects 0.000 claims description 6
- 239000007800 oxidant agent Substances 0.000 claims description 6
- 239000007921 spray Substances 0.000 claims description 6
- 239000000463 material Substances 0.000 claims description 5
- 238000012216 screening Methods 0.000 claims description 4
- 230000001133 acceleration Effects 0.000 claims description 3
- 239000011159 matrix material Substances 0.000 claims description 3
- 238000005096 rolling process Methods 0.000 claims description 3
- 230000001590 oxidative effect Effects 0.000 claims description 2
- 238000012827 research and development Methods 0.000 abstract description 7
- 238000011160 research Methods 0.000 abstract description 5
- 235000015842 Hesperis Nutrition 0.000 abstract description 3
- 235000012633 Iberis amara Nutrition 0.000 abstract description 3
- 238000012360 testing method Methods 0.000 description 4
- 238000004891 communication Methods 0.000 description 3
- 230000008878 coupling Effects 0.000 description 3
- 238000010168 coupling process Methods 0.000 description 3
- 238000005859 coupling reaction Methods 0.000 description 3
- IJGRMHOSHXDMSA-UHFFFAOYSA-N Atomic nitrogen Chemical compound N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 description 2
- 238000005457 optimization Methods 0.000 description 2
- 230000008569 process Effects 0.000 description 2
- 230000002035 prolonged effect Effects 0.000 description 2
- 201000004569 Blindness Diseases 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 239000007789 gas Substances 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 239000007788 liquid Substances 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 229910052757 nitrogen Inorganic materials 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 239000004449 solid propellant Substances 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
Landscapes
- Engineering & Computer Science (AREA)
- Remote Sensing (AREA)
- Automation & Control Theory (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Radar, Positioning & Navigation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The invention provides a method for designing an attitude control system, which relates to the technical field of small carrier rockets and comprises the following steps: receiving mission parameters for configuring the small launch vehicle to perform a mission; generating an initialization model according to the task parameters; calculating a plurality of groups of initial values which are preset for variables in the initialization model by using a genetic algorithm to obtain an optimal initial value; judging whether the optimal initial value meets a preset engineering implementation rule or not, if so, outputting an initialization model using the optimal initial value as a variable to obtain an optimal design scheme of the attitude control system of the small carrier rocket, so as to solve the technical problem that the optimal design scheme of the attitude control system of the small carrier rocket cannot be determined in the prior art, and through the design scheme of the system, the research and development period is shortened while a large amount of scientific research cost is saved.
Description
Technical Field
The invention relates to the technical field of small carrier rockets, in particular to a design method of an attitude control system.
Background
With the trend of low cost and civil aviation in the development direction of the world aerospace industry, more and more universities and colleges begin to research and develop small low-cost carrier rockets. In the existing small carrier rocket, the design scheme of the attitude control system mostly adopts the improvement of the existing scheme, and the finally determined scheme is not the optimal scheme in the executable scheme, so that excessive redundancy is brought or the control requirement cannot be well realized.
For the proposal of a new design scheme of a small carrier rocket attitude control system, a large amount of tests are required to try and verify, so that great blindness exists, the research and development cost is greatly increased, the research and development period is prolonged, and whether the optimal test scheme is a true optimal scheme cannot be determined. An improper attitude control system can greatly increase the ineffective load and reduce the performance of the small carrier rocket.
Disclosure of Invention
In view of this, the present invention aims to provide a method for designing an attitude control system, so as to solve the technical problem that the optimal design scheme of the attitude control system of a small launch vehicle cannot be determined in the prior art.
In a first aspect, an embodiment of the present invention provides a method for designing an attitude control system, where the attitude control system is applied to a small launch vehicle, and the method includes the following steps:
receiving mission parameters for configuring the small launch vehicle to perform a mission;
generating an initialization model according to the task parameters;
calculating a plurality of groups of initial values which are preset for variables in the initialization model by using a genetic algorithm to obtain an optimal initial value;
and judging whether the optimal initial value meets a preset engineering realization rule or not, and if the optimal initial value meets the preset engineering realization rule, outputting an initialization model using the optimal initial value as a variable.
With reference to the first aspect, an embodiment of the present invention provides a first possible implementation manner of the first aspect, where the method further includes:
if the optimal initial value does not meet the preset engineering realization rule, generating a limiting condition for screening the initial value according to the optimal initial value;
and calculating the initial value meeting the limiting condition by using a genetic algorithm to obtain a new optimal initial value, and continuously judging whether the new optimal initial value meets a preset engineering realization rule or not until the obtained new optimal initial value meets the preset engineering realization rule.
With reference to the first aspect, an embodiment of the present invention provides a second possible implementation manner of the first aspect, where the task parameter includes: track parameters, the quality of each sub-level structure, the structure size and the type of propellant; the generating of the initialization model according to the task parameters comprises:
determining an orbit judgment condition of an initialization model according to the orbit parameters of the small carrier rocket;
determining a quality judgment condition of an initialization model according to the quality of each sub-level structure of the small carrier rocket;
determining a size judgment condition of an initialization model according to the structure size of the small carrier rocket;
determining propellant judgment conditions of an initialization model according to the propellant type of the small carrier rocket;
and generating the initialization model according to the track judgment condition, the quality judgment condition, the size judgment condition and the propellant judgment condition.
With reference to the first aspect, an embodiment of the present invention provides a third possible implementation manner of the first aspect, where the quality determination condition includes:
wherein s.t. is a constraint condition, g is a gravitational acceleration, a1iFor maximum pitch overload of attitude control system, a2iFor maximum yaw overload of attitude control system, a3iV is a variable matrix and the total mass m of the front j-stage small carrier rocket structure for the maximum rolling overload of the attitude control system0jTotal mass m of i-th small carrier rocket structure0iAnd the i-th small carrier rocket main engine mass mmotoriAs a constant term in the constraint condition;
m is0jM is the same as0iAnd m is saidmotoriCalculated by the following way:
m0i＝msi+mmotori+mti
mmotori＝mmotorsi+mmotori,f+mmotori,o
wherein m is0jThe total mass m of the front j-stage small carrier rocket structureuFor payload mass, m0iIs the total mass, m, of the i-th small launch vehicle structuresiIs the structural mass, m, of the small-sized I-stage carrier rocket enginemotoriFor the main engine mass, m, of the i-th small carrier rockettiConnecting structural mass m for the i-th small carrier rocket cabin sectionmotorsiFor negative mass of small-sized launch vehicle engine, mmotori,fMass of main engine fuel, mmotori,oIs mainly startedMass of organic oxidizer.
With reference to the first aspect, an embodiment of the present invention provides a fourth possible implementation manner of the first aspect, where the size determination condition includes:
wherein d isi＝kdi×Di，diControl system Engine diameter for the i-th attitude, DiIs the primary engine diameter of class i, kliIn a ratio of the twoi＝kli×Di，liEngine length for the i-th attitude control system, DiIs the primary engine diameter of class i, kdiK is the ratio of the twomotor,dAnd kmotor,lIs an engineering empirical coefficient.
With reference to the first aspect, an embodiment of the present invention provides a fifth possible implementation manner of the first aspect, where the calculating, by using a genetic algorithm, a plurality of sets of initial values that are set for variables in the initialization model in advance to obtain an optimal initial value includes:
selecting initial values corresponding to preset thruster layout parameters from the initial value groups according to each group of initial values, and determining a thruster layout scheme corresponding to the obtained initial values and the number of thrusters included in the thruster layout scheme;
selecting an initial value corresponding to a preset engine parameter from the initial value group in the initial value group, determining the engine type corresponding to the obtained initial value, and calculating the mass of the engine;
selecting an initial value corresponding to a preset propellant type parameter from the initial value group in the initial value group, determining the propellant type corresponding to the obtained initial value, and calculating the mass of the propellant and the storage tank;
selecting an initial value corresponding to a preset combustion chamber shape parameter from the initial value group in the initial value group, determining the combustion chamber shape corresponding to the obtained initial value, and calculating the mass of the combustion chamber;
selecting an initial value corresponding to a preset propellant supply device parameter from the initial value group in the initial value group, determining a propellant supply device corresponding to the obtained initial value, and calculating the mass of the propellant supply device;
selecting an initial value corresponding to a preset injector parameter from the initial value group in the initial value group, determining an injector corresponding to the obtained initial value, and calculating the quality of the injector;
selecting an initial value corresponding to a preset engine spray pipe type parameter from the initial value group, determining the engine spray pipe type corresponding to the obtained initial value, and calculating the mass of the engine spray pipe;
and adding all the masses, multiplying the added masses by the number of thrusters to obtain the total mass of the attitude control system, and performing iterative calculation on a plurality of obtained total masses by using a genetic algorithm optimizer until an optimal initial value is obtained.
With reference to the first aspect, an embodiment of the present invention provides a sixth possible implementation manner of the first aspect, where, in the initial value set, an initial value corresponding to a preset propellant type parameter is selected from the initial value set, a propellant type corresponding to the obtained initial value is determined, and the calculating of the mass of the propellant and the tank includes:
calculating the mass of the propellant according to the propellant type;
and selecting a storage tank according to the propellant type, and calculating the mass of the storage tank.
With reference to the first aspect, an embodiment of the present invention provides a seventh possible implementation manner of the first aspect, where a mass calculation formula of the propellant and the tank is:
mtank＝dtankStankAtank
wherein m ispIs the mass flow of propellant, mp,oIs the mass flow of the oxidizing agent, mp,fIs fuel mass flow, r is oxygen-fuel ratio, F is single attitude control engine thrust, t is working time, IsIs specific impulse, mtankFor tank quality, dtankIs the density of the storage tank material, StankFor the wall thickness of the tank, AtankIs the surface area of the tank.
With reference to the first aspect, an embodiment of the present invention provides an eighth possible implementation manner of the first aspect, where the mass calculation formula of the combustion chamber is:
wherein m iscFor combustion chamber mass, DcIs the combustion chamber diameter, ScIs the wall thickness of the combustion chamber, LcIs the length of the combustion chamber, CFIs a thrust coefficient, [ e ]]cIs the limit of stress intensity, L*Is a characteristic length, PcIs the combustion chamber pressure, XeIs the nozzle contraction ratio, thetaiThe nozzle contraction half angle.
With reference to the first aspect, an embodiment of the present invention provides a ninth possible implementation manner of the first aspect, wherein a mass calculation formula of the injector is as follows:
wherein m isinjIs the injector mass, SinjIs the injector thickness, Δ PinjFor injector pressure drop, [ e ]]injIs the injector stress intensity limit.
The embodiment of the invention has the following beneficial effects: the embodiment of the invention receives the task parameters for configuring the small carrier rocket to execute the flight task; generating an initialization model according to the task parameters; calculating a plurality of groups of initial values which are preset for variables in the initialization model by using a genetic algorithm to obtain an optimal initial value; judging whether the optimal initial value meets a preset engineering implementation rule or not, if so, outputting an initialization model using the optimal initial value as a variable to obtain an optimal design scheme of the attitude control system of the small carrier rocket, so as to solve the technical problem that the optimal design scheme of the attitude control system of the small carrier rocket cannot be determined in the prior art, and through the design scheme of the system, the research and development period is shortened while a large amount of scientific research cost is saved.
Additional features and advantages of the invention will be set forth in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. The objectives and other advantages of the invention will be realized and attained by the structure particularly pointed out in the written description and claims hereof as well as the appended drawings.
In order to make the aforementioned and other objects, features and advantages of the present invention comprehensible, preferred embodiments accompanied with figures are described in detail below.
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 flowchart of a method for designing an attitude control system according to an embodiment of the present invention;
FIG. 2 is a flow chart of an initialization model establishment for the attitude control system according to the embodiment of the present invention;
FIG. 3 is a flowchart of calculating an optimal initial value according to an embodiment of the present invention;
fig. 4 is a flowchart of another method for designing an attitude control system according to an embodiment of the present invention.
Detailed Description
To make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings, and it is apparent 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.
At present, a large amount of tests are required to try and verify the new design scheme of the attitude control system of the small carrier rocket, so that the blind is high, the research and development cost is greatly increased, the research and development period is prolonged, and whether the optimal scheme of the test is the true optimal scheme cannot be determined.
To facilitate understanding of the present embodiment, first, a detailed description is given to a method for designing an attitude and attitude control system disclosed in the present embodiment, as shown in a flowchart of a method for designing an attitude control system shown in fig. 1, where the method includes the following steps:
step S101, receiving task parameters for configuring the small carrier rocket to execute a flight task;
in an embodiment of the present invention, the task parameters include: track parameters, quality of each sub-level structure, structure size and propellant type. The specific task parameter content is determined according to the actual operation requirement, and is not limited herein
Step S102, generating an initialization model according to the task parameters;
in an embodiment of the present invention, as shown in fig. 2, the creating an initialization model according to the task parameter includes:
step S201, determining an orbit judgment condition of an initialization model according to the orbit parameters of the small carrier rocket;
step S202, determining a quality judgment condition of an initialization model according to the quality of each sub-level structure of the small carrier rocket;
in an embodiment of the present invention, the quality determination condition includes:
wherein s.t. is a constraint condition, g is a gravitational acceleration, a1iFor maximum pitch overload of attitude control system, a2iFor maximum yaw overload of attitude control system, a3iV is a variable matrix and the total mass m of the front j-stage small carrier rocket structure for the maximum rolling overload of the attitude control system0jTotal mass m of i-th small carrier rocket structure0iAnd the i-th small carrier rocket main engine mass mmotoriAs a constant term in the constraint condition;
m is0jM is the same as0iAnd m is saidmotoriCalculated by the following way:
m0i＝msi+mmotori+mti
mmotori＝mmotorsi+mmotori,f+mmotori,o
wherein m is0jThe total mass m of the front j-stage small carrier rocket structureuFor payload mass, m0iIs the total mass, m, of the i-th small launch vehicle structuresiIs the structural mass, m, of the small-sized I-stage carrier rocket enginemotoriFor the main engine mass, m, of the i-th small carrier rockettiConnecting structural mass m for the i-th small carrier rocket cabin sectionmotorsiFor negative mass of small-sized launch vehicle engine, mmotori,fIs the main engine fuelMass of material, mmotori,oPrimary engine oxidant mass.
Step S203, determining a size judgment condition of an initialization model according to the structure size of the small carrier rocket;
in an embodiment of the present invention, the size determination condition includes:
wherein d isi＝kdi×Di，diControl system Engine diameter for the i-th attitude, DiIs the primary engine diameter of class i, kliIn a ratio of the twoi＝kli×Di，liEngine length for the i-th attitude control system, DiIs the primary engine diameter of class i, kdiK is the ratio of the twomotor,dAnd kmotor,lIs an engineering empirical coefficient.
Step S204, determining propellant judgment conditions of an initialization model according to the propellant type of the small carrier rocket;
in the embodiment of the present invention, the propellant determination condition is determined according to the kind of propellant, i.e., the supply system thereof. For example: selecting a bipropellant liquid propellant, two storage tanks and two propellant supply systems are needed, and a corresponding mass model is selected as a mass model of the attitude control system; if the solid propellant is selected, a propellant supply system is not needed, and the mass model of the attitude control system selects a corresponding mass model. Meanwhile, the existence of a shared part element is judged, if the shared part element exists, such as a cold air propulsion system, a propellant can directly shunt nitrogen from a pressurization device of a main engine supply system, and an independent gas cylinder is not needed as a container; if the attitude control engine propellant is the same as the main engine propellant, a set of propellant supply system can be shared, and a separate propellant supply system does not need to be pushed. Special constraints may be set up in the initialization model.
Step S205, generating the initialization model according to the trajectory determination condition, the quality determination condition, the size determination condition, and the propellant determination condition.
Wherein the task parameters include: the method comprises the following steps of determining judgment conditions according to each specific parameter in task parameters, generating an initialization model according to all the judgment conditions, completing establishment of the initialization model, using the initialization model as a basic model, and applying to subsequent optimization calculation. By the method, the initialization model is generated by adopting a system type design method, a large amount of experimental work is omitted, and the initialization model with high accuracy can be obtained.
Step S103, calculating a plurality of groups of initial values which are preset for variables in the initialization model by using a genetic algorithm to obtain an optimal initial value;
in the embodiment of the present invention, the calculating, by using a genetic algorithm, a plurality of sets of initial values preset for variables in the initialization model to obtain an optimal initial value, as shown in an optimal initial value calculation flowchart shown in fig. 3, includes:
step S301, aiming at each group of initial values, selecting initial values corresponding to preset thruster layout parameters from the initial value groups, and determining a thruster layout scheme corresponding to the obtained initial values and the number of thrusters included in the thruster layout scheme;
step S302, selecting an initial value corresponding to a preset engine parameter from the initial value group in the initial value group, determining the engine type corresponding to the obtained initial value, and calculating the mass of the engine;
step S303, selecting an initial value corresponding to a preset propellant type parameter from the initial value group in the initial value group, determining the propellant type corresponding to the obtained initial value, and calculating the mass of the propellant and the storage box;
in an embodiment of the present invention, the selecting an initial value corresponding to a preset propellant type parameter from an initial value group, determining a propellant type corresponding to the obtained initial value, and calculating the mass of the propellant and the storage tank includes:
calculating the mass of the propellant according to the propellant type;
and selecting a storage tank according to the propellant type, and calculating the mass of the storage tank.
The mass calculation formula of the propellant and the storage tank is as follows:
mtank＝dtankStankAtank
wherein m ispIs the mass flow of propellant, mp,oIs the mass flow of the oxidizing agent, mp,fIs fuel mass flow, r is oxygen-fuel ratio, F is single attitude control engine thrust, t is working time, IsIs specific impulse, mtankFor tank quality, dtankIs the density of the storage tank material, StankFor the wall thickness of the tank, AtankIs the surface area of the tank.
Step S304, selecting an initial value corresponding to a preset combustion chamber shape parameter from the initial value group in the initial value group, determining the combustion chamber shape corresponding to the obtained initial value, and calculating the mass of the combustion chamber;
in the embodiment of the invention, the mass calculation formula of the combustion chamber is as follows:
wherein m iscFor combustion chamber mass, DcIs the combustion chamber diameter, ScIs the wall thickness of the combustion chamber, LcIs the length of the combustion chamber, CFIs a thrust coefficient, [ e ]]cIs the limit of stress intensity, L*Is a characteristic length, PcIs the combustion chamber pressure, XeIs the nozzle contraction ratio, thetaiThe nozzle contraction half angle.
Step S305, selecting an initial value corresponding to a preset propellant supply device parameter from the initial value group in the initial value group, determining a propellant supply device corresponding to the obtained initial value, and calculating the mass of the propellant supply device;
step S306, selecting an initial value corresponding to a preset injector parameter from the initial value group in the initial value group, determining an injector corresponding to the obtained initial value, and calculating the quality of the injector;
in an embodiment of the present invention, a mass calculation formula of the injector is as follows:
wherein m isinjIs the injector mass, SinjIs the injector thickness, Δ PinjFor injector pressure drop, [ e ]]injIs the injector stress intensity limit.
Step S307, selecting an initial value corresponding to a preset engine nozzle type parameter from the initial value group in the initial value group, determining the engine nozzle type corresponding to the obtained initial value, and calculating the mass of the engine nozzle;
and S308, adding all the masses, multiplying the added masses by the number of thrusters to obtain the total mass of the attitude control system, and performing iterative computation on a plurality of obtained total masses by using a genetic algorithm optimizer until an optimal initial value is obtained.
In the embodiment of the invention, the genetic algorithm utilizes a roulette selection mechanism to select the individuals with high fitness to be inherited to the next generation, and the fitness of each generation depends on the total quality of the attitude control system obtained by adding all the qualities and multiplying the sum by the number of thrusters. In the genetic algorithm, the fitness is in direct proportion to the value calculated by the objective function, but in the step, the attitude control system with the lowest total mass is required to be selected, so that the objective function (the reciprocal of the total mass) is obtained by taking the reciprocal method to achieve the purposes that the total mass of the attitude control system is the lowest and the fitness is higher. And each generation aims at eliminating the initial value group corresponding to the value with high total quality, and storing the initial value group with high value probability corresponding to the value with low total quality until the optimal initial value group is obtained. And the outlet of the genetic algorithm is set as three outlets, wherein the condition is as follows: the difference between each generation of the objective function is less than 1 e-3; and a second condition: the difference between each generation of each variable in the initial value set is less than 1 e-3; and (3) carrying out a third condition: and stopping the calculation when the maximum preset step number is reached. And when the condition I and the condition II are met, outputting the optimal initial value, and when the condition III is met, outputting the currently output initial value as the optimal initial value no matter whether the condition I and the condition II are met. The optimal initial value is obtained through calculation of a genetic algorithm, a large amount of experimental work is reduced, the scientific research period is shortened, and the optimal initial value can be assigned to the attitude control system initialization model.
Step S104, judging whether the optimal initial value meets a preset engineering realization rule, and if the optimal initial value meets the preset engineering realization rule, outputting an initialization model using the optimal initial value as a variable.
In the embodiment of the present invention, the engineering implementation rule is a library of artificially established judgment conditions, which is established according to a large amount of experimental and empirical judgment data, for example: judging the slenderness ratio of the ith-level attitude control engine,
wherein k isλThe maximum slenderness ratio of engineering experience; judging the quality of the i-th attitude control engine propellant Mpi≤kp,mMp,majoriWherein M ispiFor the ith-class attitude control engine propellant quality, Mp,majoriIs the i-th rocket main engine propellant mass, kp,mThe maximum propellant mass coefficient of the attitude control engine; judging the thickness of the wall of the storage tank of the ith-level attitude control engine Stank,i>Stank,kIn which S istank,kThe minimum wall thickness in engineering production. If any of the engineering realization rules such as the conditions can not be met, the optimization model is modified, the limiting conditions for screening the initial values are generated according to the optimal initial values, and then the genetic algorithm pair is utilizedAnd calculating the initial value meeting the limiting condition to obtain a new optimal initial value, and continuously judging whether the new optimal initial value meets a preset engineering realization rule or not until the obtained new optimal initial value meets the preset engineering realization rule. Through the iterative feedback mode, a more accurate optimal initial value is finally obtained, the technical problem that the optimal design scheme of the small carrier rocket attitude control system cannot be determined in the prior art is solved, and through the design scheme of the system, a large amount of scientific research cost is saved, and meanwhile, the research and development period is shortened.
In yet another embodiment of the present invention, the method as shown in fig. 4 comprises:
step S101, receiving task parameters for configuring the small carrier rocket to execute a flight task;
in an embodiment of the present invention, the task parameters include: track parameters, quality of each sub-level structure, structure size and propellant type. The specific task parameter content is determined according to the actual operation requirement, and is not limited herein
Step S102, generating an initialization model according to the task parameters;
step S103, calculating a plurality of groups of initial values preset for variables in the initialization model by using a genetic algorithm to obtain an optimal initial value, and judging whether the optimal initial value meets a preset engineering realization rule or not;
and S104, if the optimal initial value meets a preset engineering realization rule, outputting an initialization model using the optimal initial value as a variable.
Step S105, if the optimal initial value does not meet the preset engineering realization rule, generating a limiting condition for screening the initial value according to the optimal initial value;
step S106, calculating the initial value meeting the limiting condition by using a genetic algorithm to obtain a new optimal initial value, and continuously judging whether the new optimal initial value meets a preset engineering realization rule or not until the obtained new optimal initial value meets the preset engineering realization rule.
In the embodiment of the present invention, the engineering implementation rule is a library of artificially established judgment conditions, which is established according to a large amount of experimental and empirical judgment data, and it can be clearly understood by those skilled in the art that, for convenience and simplicity of description, the specific working process of the system and the apparatus described above may refer to the corresponding process in the foregoing method embodiment, and details are not described herein again.
In addition, in the description of the embodiments of the present invention, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, and are only for convenience of description and simplicity of description, but do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In the several embodiments provided in the present application, it should be understood that the disclosed system, apparatus and method may be implemented in other ways. The above-described embodiments of the apparatus are merely illustrative, and for example, the division of the units is only one logical division, and there may be other divisions when actually implemented, and for example, a plurality of units or components may be combined or integrated into another system, or some features may be omitted, or not executed. In addition, the shown or discussed mutual coupling or direct coupling or communication connection may be an indirect coupling or communication connection of devices or units through some communication interfaces, and may be in an electrical, mechanical or other form.
The units described as separate parts may or may not be physically separate, and parts displayed as units may or may not be physical units, may be located in one place, or may be distributed on a plurality of network units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of the embodiment.
In addition, functional units in the embodiments of the present invention may be integrated into one processing unit, or each unit may exist alone physically, or two or more units are integrated into one unit.
Finally, it should be noted that: the above-mentioned embodiments are only specific embodiments of the present invention, which are used for illustrating the technical solutions of the present invention and not for limiting the same, and the protection scope of the present invention is not limited thereto, although the present invention is described in detail with reference to the foregoing embodiments, those skilled in the art should understand that: any person skilled in the art can modify or easily conceive the technical solutions described in the foregoing embodiments or equivalent substitutes for some technical features within the technical scope of the present disclosure; such modifications, changes or substitutions do not depart from the spirit and scope of the embodiments of the present invention, and they should be construed as being included therein. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (9)
1. An attitude control system design method is characterized in that the attitude control system is applied to a small-sized launch vehicle, and the method comprises the following steps:
receiving mission parameters for configuring the small launch vehicle to perform a mission;
generating an initialization model according to the task parameters;
calculating a plurality of groups of initial values which are preset for variables in the initialization model by using a genetic algorithm to obtain an optimal initial value;
judging whether the optimal initial value meets a preset engineering realization rule or not, and if the optimal initial value meets the preset engineering realization rule, outputting an initialization model using the optimal initial value as a variable;
the calculating a plurality of groups of initial values preset for variables in the initialization model by using a genetic algorithm to obtain an optimal initial value comprises the following steps:
selecting initial values corresponding to preset thruster layout parameters from the initial value groups according to each group of initial values, and determining a thruster layout scheme corresponding to the obtained initial values and the number of thrusters included in the thruster layout scheme;
selecting an initial value corresponding to a preset engine parameter from the initial value group in the initial value group, determining the engine type corresponding to the obtained initial value, and calculating the mass of the engine;
selecting an initial value corresponding to a preset propellant type parameter from the initial value group in the initial value group, determining the propellant type corresponding to the obtained initial value, and calculating the mass of the propellant and the storage tank;
selecting an initial value corresponding to a preset combustion chamber shape parameter from the initial value group in the initial value group, determining the combustion chamber shape corresponding to the obtained initial value, and calculating the mass of the combustion chamber;
selecting an initial value corresponding to a preset propellant supply device parameter from the initial value group in the initial value group, determining a propellant supply device corresponding to the obtained initial value, and calculating the mass of the propellant supply device;
selecting an initial value corresponding to a preset injector parameter from the initial value group in the initial value group, determining an injector corresponding to the obtained initial value, and calculating the quality of the injector;
selecting an initial value corresponding to a preset engine spray pipe type parameter from the initial value group, determining the engine spray pipe type corresponding to the obtained initial value, and calculating the mass of the engine spray pipe;
and adding all the masses, multiplying the added masses by the number of thrusters to obtain the total mass of the attitude control system, and performing iterative calculation on a plurality of obtained total masses by using a genetic algorithm optimizer until an optimal initial value is obtained.
2. The attitude control system design method according to claim 1, characterized in that the method further comprises:
if the optimal initial value does not meet the preset engineering realization rule, generating a limiting condition for screening the initial value according to the optimal initial value;
and calculating the initial value meeting the limiting condition by using a genetic algorithm to obtain a new optimal initial value, and continuously judging whether the new optimal initial value meets a preset engineering realization rule or not until the obtained new optimal initial value meets the preset engineering realization rule.
3. The attitude control system design method according to claim 1, wherein the task parameters include: track parameters, the quality of each sub-level structure, the structure size and the type of propellant; the generating of the initialization model according to the task parameters comprises:
determining an orbit judgment condition of an initialization model according to the orbit parameters of the small carrier rocket;
determining a quality judgment condition of an initialization model according to the quality of each sub-level structure of the small carrier rocket;
determining a size judgment condition of an initialization model according to the structure size of the small carrier rocket;
determining propellant judgment conditions of an initialization model according to the propellant type of the small carrier rocket;
and generating the initialization model according to the track judgment condition, the quality judgment condition, the size judgment condition and the propellant judgment condition.
4. An attitude control system design method according to claim 3, characterized in that the quality determination condition includes:
wherein s.t. is a constraint condition, g is a gravitational acceleration, a1iFor maximum pitch overload of attitude control system, a2iFor maximum yaw overload of attitude control system, a3iV is a variable matrix and the total mass m of the front j-stage small carrier rocket structure for the maximum rolling overload of the attitude control system0jTotal mass m of i-th small carrier rocket structure0iAnd the i-th small carrier rocket main engine mass mmotoriAs a constant term in the constraint, g1i(V) is a first constraint function, g2i(V) is a second constraint function, g3i(V) is a third constraint function;
m is0jM is the same as0iAnd m is saidmotoriCalculated by the following way:
m0i＝msi+mmotori+mti
mmotori＝mmotorsi+mmotori,f+mmotori,o
wherein m is0jThe total mass m of the front j-stage small carrier rocket structureuFor payload mass, m0iIs the total mass, m, of the i-th small launch vehicle structuresiIs the structural mass, m, of the small-sized I-stage carrier rocket enginemotoriFor the main engine mass, m, of the i-th small carrier rockettiConnecting structural mass m for the i-th small carrier rocket cabin sectionmotorsiFor negative mass of small-sized launch vehicle engine, mmotori,fMass of main engine fuel, mmotori,oPrimary engine oxidant mass.
5. The attitude control system design method according to claim 3, wherein the size determination condition includes:
wherein d isi＝kdi×Di，diControl system Engine diameter for the i-th attitude, DiIs the primary engine diameter of class i, kliIn a ratio of the twoi＝kli×Di，liEngine length for the i-th attitude control system, DiIs the primary engine diameter of class i, kdiK is the ratio of the twomotor,dAnd kmotor,lIs an engineering empirical coefficient, g4i(V) is a fourth constraint function, g5i(V) is a fifth constraint function.
6. An attitude control system design method according to claim 1, wherein said selecting an initial value corresponding to a preset propellant kind parameter from among the initial value groups in the initial value group, determining a propellant kind corresponding to the obtained initial value, and calculating the mass of the propellant and the tank comprises:
calculating the mass of the propellant according to the propellant type;
and selecting a storage tank according to the propellant type, and calculating the mass of the storage tank.
7. An attitude control system design method according to claim 6, characterized in that the mass calculation formula of the propellant and the tank is:
mtank＝dtankStankAtank
wherein m ispIs the mass flow of propellant, mp,oIs the mass flow of the oxidizing agent, mp,fIs fuel mass flow, r is oxygen-fuel ratio, F is single attitude control engine thrust, t is working time, IsIs specific impulse, mtankFor tank quality, dtankIs the density of the storage tank material, StankFor the wall thickness of the tank, AtankIs the surface area of the tank.
8. The attitude control system design method according to claim 1, characterized in that a mass calculation formula of a combustion chamber:
wherein m iscFor combustion chamber mass, DcIs the combustion chamber diameter, ScIs the wall thickness of the combustion chamber, LcIs the length of the combustion chamber, CFIs a thrust coefficient, [ e ]]cIs the limit of stress intensity, L*Is a characteristic length, PcIs the combustion chamber pressure, XeIs the nozzle contraction ratio, thetaiThe nozzle contraction half angle, d the combustion chamber material density, and F the thrust.
9. A method of designing a pose control system according to claim 1, wherein the mass calculation formula of the injector is:
wherein m isinjIs the injector mass, SinjIs the injector thickness, Δ PinjFor injector pressure drop, [ e ]]injIs the injector stress intensity limit.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810743622.1A CN108931987B (en) | 2018-07-09 | 2018-07-09 | Attitude control system design method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810743622.1A CN108931987B (en) | 2018-07-09 | 2018-07-09 | Attitude control system design method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN108931987A CN108931987A (en) | 2018-12-04 |
| CN108931987B true CN108931987B (en) | 2020-06-19 |
Family
ID=64447950
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201810743622.1A Active CN108931987B (en) | 2018-07-09 | 2018-07-09 | Attitude control system design method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN108931987B (en) |
Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109669470B (en) * | 2018-12-05 | 2021-08-10 | 北京航天自动控制研究所 | Kinematics constraint conversion method for vertical take-off and landing rocket online trajectory planning |
| CN110104218B (en) * | 2019-04-19 | 2021-04-30 | 北京航天自动控制研究所 | Pre-deflection angle nonlinear compensation method and device for rocket engine frame deformation angle |
| CN110531622B (en) * | 2019-09-05 | 2022-04-05 | 沈阳航空航天大学 | Thrust control method of solid rocket engine based on radial basis function neural network |
| CN114326767A (en) * | 2021-12-14 | 2022-04-12 | 宁波天擎航天科技有限公司 | Aircraft attitude control method, aircraft attitude control system and aircraft |
| CN114253282B (en) * | 2021-12-21 | 2023-09-22 | 航天科工火箭技术有限公司 | Carrier rocket attitude control method, device, equipment and storage medium |
| CN115310211B (en) * | 2022-10-11 | 2022-12-20 | 中国人民解放军63921部队 | Overall design method and device for vertical take-off and landing reusable launch vehicle |
Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103955557B (en) * | 2014-03-31 | 2017-01-04 | 北京航空航天大学 | Carrier rocket multidisciplinary comprehensive design optimization method and system |
| CN103942401B (en) * | 2014-05-14 | 2017-03-15 | 哈尔滨工业大学 | A kind of spacecraft trajectory multiple constraint track tool kit and method |
| US9770873B2 (en) * | 2014-08-22 | 2017-09-26 | The Boeing Company | System and method for optimizing composite laminate structures |
| CN106021784B (en) * | 2016-05-31 | 2019-01-11 | 北京航空航天大学 | A kind of full track mark optimum design method based on bilevel optimization strategy |
| CN106354901B (en) * | 2016-08-12 | 2019-10-22 | 北京宇航系统工程研究所 | A kind of carrier rocket mass property and dynamics key parameter on-line identification method |
-
2018
- 2018-07-09 CN CN201810743622.1A patent/CN108931987B/en active Active
Also Published As
| Publication number | Publication date |
|---|---|
| CN108931987A (en) | 2018-12-04 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN108931987B (en) | Attitude control system design method | |
| CN107092731B (en) | Inter-stage specific thrust trajectory integrated optimization method for sub-orbital launch vehicle | |
| CN110276159B (en) | Multi-model fusion-based satellite system multidisciplinary optimization method | |
| CN104007665A (en) | Flight simulation test system for solid-liquid power aircraft | |
| CN108763747A (en) | Carrier rocket dimensional parameters evaluation method | |
| CN108763746B (en) | Carrier rocket mass parameter estimation method | |
| Zheng et al. | Multidisciplinary integrated design of long-range ballistic missile using PSO algorithm | |
| CN114384799A (en) | Boosting and core-level engine combined thrust vector control method | |
| Olds | Multidisciplinary design techniques applied to conceptual aerospace vehicle design | |
| Chen et al. | Development of a small launch vehicle with hybrid rocket propulsion | |
| CN116382124A (en) | Carrier rocket attitude control simulation method and system | |
| Gupta et al. | Evolution of Indian launch vehicle technologies | |
| Choi et al. | Optimal air-launching rocket design using system trades and a multi-disciplinary optimization approach | |
| Lee | Mission and trajectory optimization of the air-launching rocket system using MDO techniques | |
| Paulino et al. | Fault Tolerant Control for a Cluster of Rocket Engines–Methods and outcomes for guidance and control recovery strategies in launchers | |
| Tanil et al. | External configuration optimization of missiles in conceptual design | |
| CN109917801B (en) | Load balance control distribution method of MEMS micro-thrust array | |
| CN113110535A (en) | Spacecraft attitude control method under multi-constraint condition | |
| Leverone | Performance modelling and simulation of a 100km hybrid sounding rocket. | |
| CN113485395A (en) | Fixed time relative attitude and orbit tracking control method under error constraint | |
| CN112393648A (en) | Balance flight theoretical method for autonomous control under rocket thrust failure mode | |
| Messinger | Conceptual Design and Optimization of Hybrid Rockets | |
| CN109976362B (en) | Rocket upper-level trajectory re-planning method using differential evolution algorithm | |
| Sagliano | Apollo 11 reloaded: Optimization-based trajectory reconstruction | |
| 김영훈 | Investigation of the Aerodynamic Characteristics for the Low Aspect Ratio Launch Vehicle with the External Structure |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |