CN105930582A - Spacecraft orbit maneuver motor installation parameter optimization method - Google Patents

Abstract

The present invention provides a spacecraft orbit maneuver motor installation parameter optimization method. The method comprises the steps of: firstly, defining each coordinate system, and establishing a relation among a motor thrust vector, an acting point position vector and an undetermined coefficient; then calculating a relation between a disturbance torque value during each ignition; and with an optimization goal of making a maximum and a minimum of a disturbance torque component during each ignition equal in numerical value but opposite in sign, solving the undetermined coefficient, so as to obtain an installation angle of a spacecraft orbit maneuver motor and an installation position of a motor nozzle. According to the method provided by the present invention, an optimally designed mathematical equation is constructed, and the demand for optimization of the installation parameter of the spacecraft orbit maneuver motor is satisfied to the largest extent.

Description

A kind of spacecraft orbit maneuver motor installation parameter optimization method

Technical field

The present invention relates to a kind of spacecraft orbit maneuver motor installation parameter optimization method, belong to spacecraft overall Design field.

Background technology

The result of calculation of spacecraft orbit maneuver motor disturbance torque, is to determine that the important of spacecraft layout depends on According to, also it is to check whether meet delivery and control the foundation of Subsystem Design index.Due to spacecraft centroid Deviation, motor power vector deviation and general assembly deviation, can produce perturbed force during causing engine ignition Square.

In the spacecraft stage, the heat mark data that spacecraft totally provides according to reseach of engine unit, The installation requirement of electromotor need to be proposed.Disturbance torque during generally becoming rail is the smaller the better.

Prior art provides as follows: in order to ensure adjust after 490N electromotor thrust vectoring with Satellite machinery coordinate system-Z axis less than 0.1 °, adjust the relation of angle, θ and thrust vectoring skew angle such as Under: when 0.1 ° < during α≤0.12 °, θ=0.5 α (readjustment half)；As α > 0.12 ° time, θ=α-0.06. The shortcoming of prior art is, design barycenter bias is the most serious, and required counterweight is the most.Design matter at present Heart bias generally compares relatively big, and satellite layout is restricted by many factors, is difficult to be greatly decreased design barycenter, Determine prior art counterweight the most higher.

Conventional engine parameter choosing method, is a kind of fairly simple compromise algorithm.Send out by calculating Existing, according to previous methods, the disturbance torque of conceptual design self is relatively big sometimes, the most even close to wanting Seek the upper limit of scope.It is necessary to carry out the research of spacecraft orbit maneuver motor installation parameter optimization method.

Summary of the invention

The technology of the present invention solves problem: overcome the deficiencies in the prior art, it is provided that a kind of spacecraft becomes The optimization method of rail electromotor installation parameter, constructs the mathematical equation of optimized design, solves full On the premise of foot constraints, the optimal value of electromotor installation parameter (makes perturbed force during each igniting Square component maximum and minima numerical value are equal, and symbol is contrary) so that object function is optimum, maximum The demand that spacecraft orbit maneuver motor installation parameter optimizes is met in degree.

The technical solution of the present invention is: the optimization side of a kind of spacecraft orbit maneuver motor installation parameter Method, step is as follows:

(1) spacecraft mechanical coordinate system C is set up_S, spacecraft translational coordination system C_S", spacecraft centroid sit Mark system C_C, engine body coordinate system C_EB, engine target coordinate system C_TWith electromotor mounting coordinate It is C_EI；

Described spacecraft mechanical coordinate system C_SInitial point O_SIt is positioned at the mechanical separation of satellite and carrier rocket In face, and overlap with the center of circle of origin reference location pin institute formational theory circle in mechanical separation face, X_SAxle is square Spacecraft east plate, Y is pointed to from zero_SAxle positive direction points to spacecraft south plate from zero, Z_SAxle meets the right-hand rule；

Described spacecraft translational coordination system C_S"By spacecraft mechanical coordinate system C_STranslation obtains, and spacecraft is put down Move coordinate system C_S"Zero be motor mounting flange theory center of circle A；

Described spacecraft centroid coordinate system C_CBy spacecraft mechanical coordinate system C_STranslation obtains, spacecraft matter Heart coordinate system C_CZero be positioned at spacecraft centroid；

Described engine body coordinate system C_EBZero be positioned at the theoretical circle of motor mounting flange Heart A, X_EBAxle positive direction and spacecraft mechanical coordinate system Z_SAxle positive direction is consistent, Y_EBAxle positive direction with Spacecraft mechanical coordinate system Y_SAxle negative direction is consistent, Z_EBAxle positive direction and spacecraft mechanical coordinate system X_S Axle positive direction is consistent；

Described engine target coordinate system C_TBy spacecraft translational coordination system C_S" rotate and obtain, electromotor Target-based coordinate system C_TZ_TAxle negative direction is to point to spout direction along electromotor theory geometrical axis；

Described electromotor mounting coordinate system C_EIFor engine target coordinate system C_TAround+X_TAxle rotates 180 ° And obtain, Z_EIThe positive direction of axle is to point to spout direction along electromotor theory geometrical axis；

(2) according to previously given thrust vectoring skew angle, thrust vectoringTraversing position angle beta, push away Force vectorDeflected position angle γ and traversing amount δ of thrust vectoring application point P, at engine body coordinate It is C_EBUnder, calculate motor power vector F_EBWith position of action point vector

(3) according to the position of action point vector in step (2)Motor power vector F_EB In conjunction with engine body coordinate system C_EBTo electromotor mounting coordinate system C_EITransformation matrix of coordinates Obtain at electromotor mounting coordinate system C_EIUnder position of action point vectorWith motor power vector F_EI；

(4) order is from satellite translational coordination system C_S"Transform to engine target coordinate system C_TProcess be: By satellite translational coordination system O_S"X_S"Y_S"Z_S"First around+X_S"Axle anglec of rotation α_T(°), further around Y_TAxle rotation β_T(°), α_T(°) and β_T(°) is undetermined coefficient；Determine electromotor mounting coordinate system C_EITo setting up spacecraft mechanical coordinate It is C_SSpin matrixWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；And at spacecraft machine Tool coordinate system C_SUnder, the theoretical center of circle A setting up motor mounting flange issues to engine tilts state The vector of motivation mounting flange center of circle CWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；

(5) vector obtained in step (4) is utilizedWith undetermined coefficient α_T(°) and β_TPass between (°) It is formula, determines at spacecraft mechanical coordinate system C_SUnder, O_STo motor mounting flange theory center of circle A's Vector

(6) spin matrix in step (4) is utilizedWith undetermined coefficient α_T(°) and β_TRelation between (°) Motor power vector F in formula integrating step (3)_EI, position of action point vectorAnd step (5) vector inObtain spacecraft mechanical coordinate system C_SLower electromotor thrust vectoring F_S、 Position of action point vectorWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；

(7) utilize two groups of relational expressions in step (6) calculate during each igniting disturbance torque value M and Undetermined coefficient α_T(°) and β_TRelational expression between (°)；

(8) during each igniting, disturbance torque component maximum and minima numerical value are equal, and symbol Contrary is optimization aim, and the relational expression in solution procedure (7) obtains undetermined coefficient α_T(°) and β_T(°), And then obtain spin matrix

(9) according to spin matrixObtain spacecraft orbit maneuver motor setting angle and engine nozzle Installation site B.

Position vector in described step (3)With motor power vector F_EIComputing formula is as follows:

$F_{EI}=R_{EB}^M{F}_{EB}$

Motor power vector F in described step (6)_S, position of action point vectorWith undetermined Factor alpha_T(°) and β_TRelational expression between (°) is as follows:

$F_S=R_{EI}^S{F}_{EI}$

In formula,

Relational expression in step (7) is as follows:

In satellite machinery coordinate system C_SUnder, the position of centroid of satellite during precise tracking igniting Vector.

According to the result in step (7), disturbance torque component maximum and minima during each igniting Numerical value is equal, and symbol is optimization aim on the contrary, is embodied as equation below:

max(M_x,M_y,M_Z)_{I=1, n}+min(M_x,M_y,M_Z)_{I=1, n}=0

Wherein, M_x,M_y,M_ZFor the disturbance torque three axle components under spacecraft centroid coordinate system.

The present invention compared with prior art provides the benefit that:

The present invention is directed to orbit maneuver motor installation parameter On The Choice, propose a kind of electromotor installation parameter Optimization method.The method with each time igniting during disturbance torque component maximum and minima numerical value is equal and Symbol is optimization aim on the contrary, uses multiple optimized algorithm to solve, it was demonstrated that the method achieves change rail and starts The purpose of the optimized design of machine installation parameter, overcomes the deficiencies in the prior art, with current electromotor peace Dress parameter confirmation method is compared, and can reach disturbance torque and minimize, thus decrease entrained by satellite Counterweight, saves spacecraft fuel, improves the spacecraft life-span in-orbit.

Accompanying drawing explanation

Fig. 1 is the flow chart of method involved in the present invention；

Fig. 2 is engine body coordinate system and spacecraft mechanical coordinate system schematic diagram；

Fig. 3 is electromotor scheme of installation on engine support；

Fig. 4 is securing member scheme of installation between electromotor and engine support；

Fig. 5 is engine thermal mark parameter space schematic diagram under electromotor coordinate system；

Fig. 6 is the relativeness schematic diagram of each coordinate system；

Fig. 7 is orbit maneuver motor coordinate system schematic diagram；

Fig. 8 is the disturbance torque cartogram during each change rail；

Fig. 9 is Different Optimization Algorithm for Solving precision block diagram；

Figure 10 is MATLAB result of calculation in embodiment.

Detailed description of the invention

Below in conjunction with the accompanying drawings the detailed description of the invention of the present invention is further described in detail.

It is illustrated in figure 2 engine body coordinate system and spacecraft mechanical coordinate system schematic diagram；Satellite machinery Coordinate system is defined as follows:

Coordinate origin Osc is positioned at satellite lower end frame and carrier rocket mechanical separation face, with satellite On interface, the center of circle of the theoretical circle that pin is formed overlaps；

OscXsc axle is consistent with satellite east plate theory normal direction, positive direction and east plate exterior normal Direction is consistent；

OscYsc axle is consistent with satellite north plate theory normal direction, positive direction and south plate exterior normal Direction is consistent；

OscZsc axle is perpendicular to the connection parting surface of satellite and carrier rocket, and its positive direction is from initial point Point to floor；

OscXscYscZsc coordinate system meets right-hand rule.

Engine body coordinate system is defined as follows:

Electromotor self also has a coordinate system, its initial point be positioned on the Zsc axle of satellite machinery coordinate system away from Being away from H from its initial point, the Zsc axle of the X-axis forward of electromotor coordinate system and satellite machinery coordinate system is just To identical, the Y-axis forward of electromotor coordinate system is identical with the Ysc axle negative sense of satellite machinery coordinate system, The Z axis of electromotor coordinate system and X-axis, Y-axis meet right-hand rule.

It is illustrated in figure 3 electromotor scheme of installation on engine support.As can be known from Fig. 3, send out Motivation 1 is by motor mounting flange 3 (top surface edge 31, lower surface edge 32), installation spiral shell Nail 4 is fixedly mounted on the engine support ring flange 2 of engine support 5, under original state, sends out The axis of motivation is vertical with the plane at motor mounting flange place；

Fig. 4 is securing member scheme of installation between electromotor and engine support, as can be seen from Figure 4, electromotor Insulation thermal pad is installed between 1 and engine support 5 and adjusts pad 6；

Fig. 5 show engine thermal mark parameter space schematic diagram under electromotor coordinate system, such as Fig. 5 institute Show, for engine thermal mark parameter space schematic diagram under electromotor coordinate system, wherein X, Y, Z generation The coordinate axes of table electromotor coordinate system, other meaning of parameters is as follows:

α thrust vectoring angle of deviation (on the basis of X-axis), unit degree；

The traversing position angle of β thrust vectoring is (on the basis of Y-axis, by the electromotor top view direction inverse time Pin is just), unit degree；

γ thrust vectoring deflected position angle is (on the basis of Y-axis, counterclockwise by electromotor top view direction For just), unit degree；

The traversing amount of δ thrust vectoring (away from the distance of zero), unit mm.

P is motor power application point, and F represents thrust vectoring.

According to the requirement of spacecraft master-plan, build the object function of electromotor installation parameter optimized design. Herein propose, choose disturbance torque component maximum and minima numerical value during each igniting equal, and symbol Number be on the contrary optimization aim, employing simulated annealing, genetic algorithm, minima method, pattern search algorithm Four kinds of optimized algorithms.Contrast effect of optimization, uses best optimized algorithm to solve this optimization problem.

About spacecraft translational coordination system C_S"(spacecraft mechanical coordinate system C_S) and engine target coordinate system C_TThe relation of coordinate system, specify herein, from spacecraft translational coordination system C_S"(spacecraft mechanical coordinate system C_S) to engine target coordinate system C_TTransformation matrix of coordinates, it is assumed that by spacecraft translational coordination system O_S"X_S"Y_S"Z_S"(C_S") first around+X_S"Axle anglec of rotation α_T(°), becomes O_S"X_S"Y_S""Z_S"", i.e. O_S"X_S"Y_TZ_S"", Then further around Y_TAxle rotation β_T(°), becomes coordinate system O_TX_TY_TZ_T(C_T), order can not be exchanged.Order is not Can exchange.Therefore α_T(°) and β_T(°) is undetermined coefficient.

Due to engine support mounting flange (plane) on spacecraft fixing, it is contemplated that electromotor is installed The impact of flange, in rotary course, motor mounting flange will be along spacecraft mechanical coordinate system C_S+ Z_STo Lower movement.

For the ease of installing, it is stipulated that in installation process, electromotor Method for Installation under the non-heeling condition of electromotor The vector of blue theoretical center of circle C to motor mounting flange theory center of circle AWith spacecraft mechanical coordinate system C_S+ Z_SParallel, i.e. without moving horizontally.

It is illustrated in figure 1 the method flow diagram of the present invention, from fig. 1, it can be seen that a kind of boat that the present invention proposes It device orbit maneuver motor installation parameter optimization method, step is as follows:

(1) spacecraft mechanical coordinate system C is set up_S, spacecraft translational coordination system C_S", spacecraft centroid sit Mark system C_C, engine body coordinate system C_EB, engine target coordinate system C_TWith electromotor mounting coordinate It is C_EI；The mutual relation of each coordinate system is as shown in Figure 6:

Spacecraft mechanical coordinate system C_SInitial point be positioned at the mechanical separation face of satellite and carrier rocket, and Overlap with the center of circle of origin reference location pin institute formational theory circle in mechanical separation face, X_SAxle positive direction is from coordinate Initial point points to spacecraft east plate, Y_SAxle positive direction points to spacecraft south plate, Z from zero_SAxle meets The right-hand rule；

Spacecraft translational coordination system C_S"By spacecraft mechanical coordinate system C_STranslation obtains, and spacecraft translation is sat Mark system C_S"Zero be motor mounting flange theory center of circle A；

Spacecraft centroid coordinate system C_CBy spacecraft mechanical coordinate system C_STranslation obtains, and spacecraft centroid is sat Mark system C_CZero be positioned at spacecraft centroid；

Engine body coordinate system C_EBZero be positioned at the theoretical center of circle A of motor mounting flange, X_EBAxle positive direction and spacecraft mechanical coordinate system Z_SAxle positive direction is consistent, Y_EBAxle positive direction and spacecraft Mechanical coordinate system Y_SAxle negative direction is consistent, Z_EBAxle positive direction and spacecraft mechanical coordinate system X_SAxle is square To unanimously；

Engine target coordinate system C_TBy spacecraft translational coordination system C_S" rotate and obtain, engine target Coordinate system C_TZ_TAxle negative direction is to point to spout direction along electromotor theory geometrical axis；Electromotor mesh Mark coordinate system C_TCoordinate axes X_T、Y_T、Z_TRespectively with spacecraft translational coordination system C_S"Coordinate axes X_S"、Y_S"、 Z_S"Angle is acute angle.

Electromotor mounting coordinate system C_EIFor engine target coordinate system C_TAround+X_TAxle rotates 180 ° and obtains Arrive, Z_EIThe positive direction of axle is to point to spout direction along electromotor theory geometrical axis；

The geometrical axis taking special frock to measure electromotor, tooling shaft is needed when installing orbit maneuver motor LinePositive direction is defined as along engine head to nozzle exit direction.Frock axis direction can be considered as Electromotor theory geometrical axis direction, when adjusting orbit maneuver motor with electromotor frock axis and spacecraft Mechanical coordinate system coordinate axes X_S、Y_S、Z_SAngle determine.

(2) as it is shown in fig. 7, according to previously given thrust vectoring skew angle, thrust vectoringTraversing position Angle setting β, thrust vectoringDeflected position angle γ and traversing amount δ of thrust vectoring application point P, at electromotor Body coordinate system C_EBUnder, calculate motor power vector F_EBWith position of action point vector

Motor power vector F_EBWith position of action point vectorComputing formula is as follows:

F_EB=F (cos α sin α cos γ sin α sin γ)^T

In formula, F is motor power, and thrust vectoring skew angle is electromotor X_EBAxle positive direction with push away Force vectorBetween acute angle；Thrust vectoring traversing position angle beta is Y_EBAxle positive direction and thrust vectoring At Y_EBO_EBZ_EBAngle between plane projection, thrust vectoring deflected position angle γ is Y_EBAxle withBetween angle, traversing amount δ of thrust vectoring is that thrust point is away from zero O_EBAway from From.

(3) according to the position of action point vector in step (2)Motor power vector F_EB In conjunction with engine body coordinate system C_EBTo electromotor mounting coordinate system C_EITransformation matrix of coordinates Obtain at electromotor mounting coordinate system C_EIUnder position of action point vectorWith motor power vector F_EI；

Position vectorWith motor power vector F_EIComputing formula is as follows:

$F_{EI}=R_{EB}^M{F}_{EB}$

(4) order is from satellite translational coordination system C_S"Transform to engine target coordinate system C_TProcess be: will defend Star translational coordination system O_S"X_S"Y_S"Z_S"First around+X_S"Axle anglec of rotation α_T(°), further around Y_TAxle rotation β_T(°), α_T(°) and β_T(°) is undetermined coefficient；Determine electromotor mounting coordinate system C_EITo setting up spacecraft mechanical coordinate system C_SRotation Torque battle arrayWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；And at spacecraft mechanical coordinate system C_S Under, set up the theoretical center of circle A of motor mounting flange to motor mounting flange circle under engine tilts state The vector of heart CWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；

Electromotor mounting coordinate system C_EITo setting up spacecraft mechanical coordinate system C_SSpin matrixSpecifically By formula:

Be given,

Wherein:

$R_T^S={\left(R_S^T\right)}^T$

$R_S^T=R_y\left({\mathrm{\&beta;}}_T\right)R_x\left({\mathrm{\&alpha;}}_T\right)$

$R_x\left({\mathrm{\&alpha;}}_T\right)=\left[\begin{array}{ccc}1& 0& 0\\ 0& {\mathrm{cos\&alpha;}}_T& {\mathrm{sin\&alpha;}}_T\\ 0& -{\mathrm{sin\&alpha;}}_T& {\mathrm{cos\&alpha;}}_T\end{array}\right]$

$R_y\left({\mathrm{\&beta;}}_T\right)=\left[\begin{array}{ccc}{\mathrm{cos\&beta;}}_T& 0& -{\mathrm{sin\&beta;}}_T\\ 0& 1& 0\\ {\mathrm{sin\&beta;}}_T& 0& {\mathrm{cos\&beta;}}_T\end{array}\right]$

The theoretical center of circle A of motor mounting flange to motor mounting flange center of circle C under engine tilts state VectorComputing formula is as follows:

Assume installing in end face edge along+Z_SThe vector of the point translating up maximum isConcrete formula:

Wherein, r_EfRadius for motor mounting flange.

(5) vector obtained in step (4) is utilizedWith undetermined coefficient α_T(°) and β_TPass between (°) It is formula, determines at spacecraft mechanical coordinate system C_SUnder, O_STo motor mounting flange theory center of circle A's Vector

O_SVector to motor mounting flange theory center of circle AComputing formula is as follows:

Be given,For at spacecraft mechanical coordinate system C_SUnder, spacecraft mechanical coordinate system C_SSeat Mark initial point O_SVector to motor mounting flange theory center of circle A；For sitting at spacecraft machinery Mark system C_SUnder, spacecraft mechanical coordinate system C_SZero O_SIssue to the non-heeling condition of electromotor The vector of motivation mounting flange theory center of circle C.

(6) spin matrix in step (4) is utilizedWith undetermined coefficient α_T(°) and β_TRelation between (°) Motor power vector F in formula integrating step (3)_EI, position of action point vectorAnd step (5) vector inObtain spacecraft mechanical coordinate system C_SLower electromotor thrust vectoring F_S、 Position of action point vectorWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；

Spacecraft mechanical coordinate system C_SLower electromotor thrust vectoring F_S, position of action point vectorWith Undetermined coefficient α_T(°) and β_TComputing formula between (°) is as follows:

Specifically by formula:

$F_S=R_{EI}^S{F}_{EI}$

Be given, in formula,In satellite machinery coordinate system C_SUnder, send out The vector of motivation mounting flange theory center of circle A to motor power vector application point P.

(7) utilize two groups of relational expressions in step (6) calculate during each igniting disturbance torque value M and Undetermined coefficient α_T(°) and β_TRelational expression between (°)；

Disturbance torque value M and undetermined coefficient α during each igniting_T(°) and β_TComputing formula between (°) is such as Under:

Define according to coordinate system, F_C=F_S,Therefore:

Wherein:

F_EI: at electromotor mounting coordinate system C_EIUnder, motor power vector；

In satellite machinery coordinate system C_SUnder, the position vector of motor power vector application point P；

In satellite machinery coordinate system C_SUnder, the position of centroid of satellite during precise tracking igniting Vector.

(8) during each igniting, disturbance torque component maximum and minima numerical value are equal, and symbol Contrary is optimization aim, and the relational expression in solution procedure (7) obtains undetermined coefficient α_T(°) and β_T(°), And then obtain spin matrix

During each igniting, disturbance torque component maximum and minima numerical value are equal, and symbol is on the contrary Optimization aim, specifically by formula:

max(M_x,M_y,M_Z)_{I=1, n}+min(M_x,M_y,M_Z)_{I=1, n}=0

Obtain undetermined coefficient α_T(°) and β_T(°), and then obtain spin matrix

Wherein, M_x,M_y,M_ZFor the disturbance torque three axle components under spacecraft centroid coordinate system.

About solving, in order to improve solving precision, simulated annealing, genetic algorithm, minimum are attempted Value method, four kinds of optimized algorithms of pattern search algorithm.By the above four kinds of algorithms of contrast, choose precision the highest Solution.Contrast display, simulated annealing precision is the highest, and computational accuracy is up to 10^-12.Other three kinds Arithmetic accuracy is relatively low, and only 10^-5～10^-8(because having random factor, precision slightly to float).Simulation Annealing algorithm is a kind of searching algorithm of overall importance, even if given initial value is undesirable, also can obtain better result. Genetic algorithm is the most sensitive to initial value, when initial value is preferable, it is possible to obtain better result, otherwise result is difficult To ensure.According to solving precision height, finally use simulated annealing, meet boat to the full extent The demand that it device orbit maneuver motor installation parameter optimizes.

(9) according to spin matrixObtain spacecraft orbit maneuver motor setting angle and engine nozzle Installation site B.

Solve spin matrix in spacecraft orbit maneuver motor setting angle, i.e. solution procedure (8)Each unit The inverse cosine value of element；It is referred to as by C_EITo C_SThe transformation matrix of coordinates of coordinate system, the element of square formation is exactly Direction cosines between corresponding coordinate axle.Setting angle is C_EIAnd C_SAngle (scope between coordinate axes Belong to [0, π]), i.e.Each element inverse cosine value.

Installation site B of engine nozzle is specifically by formula:

Be given, whereinBy formula:

Be given,By formula:

Being given, | AB | is previously given engine nozzle theory center of circle B to mounting flange theory center of circle A Distance.

Embodiment

Initial conditions (known conditions)

Engine thermal mark data (thrust vectoring and position of action point vector), the most as shown in table 1:

Table 1

Electromotor mechanical dimension (the engine nozzle theory center of circle to mounting flange theory center of circle distance) is such as table 2 Shown in:

Table 2

Engine support position is as shown in table 3:

Table 3

Spacecraft centroid coordinate dataAs shown in table 4:

Table 4

(2) specifically solve:

A () has write solver in Matlab software, including mastery routine and post processor.Main Program comprises parameter input subroutine and interative routine.In order to improve solving precision, program provides Simulated annealing, genetic algorithm, minima method, four kinds of optimized algorithms of pattern search algorithm.Different excellent Change the solving precision of algorithm respectively as shown in table 5 and Fig. 9.

Table 5

By the above four kinds of algorithms of contrast, choose the solution that precision is the highest.Contrast display, simulated annealing Precision is the highest, and computational accuracy is up to 10^-12.Other three kinds of arithmetic accuracy are relatively low, and only 10^-5～10^-8 (because having random factor, precision slightly to float).Simulated annealing is a kind of searching algorithm of overall importance, Even if given initial value is undesirable, better result also can be obtained.Genetic algorithm is the most sensitive to initial value, originally When being worth preferable, it is possible to obtain better result, otherwise result is difficult to ensure that.According to solving precision height, Use simulated annealing eventually, meet what spacecraft orbit maneuver motor installation parameter optimized to the full extent Demand.

In terms of search speed, simulated annealing is the longest, each run the most about 2 minutes, its Time-consuming about 1 minute of his algorithm each run, is satisfied by requirement of engineering.

B () is according to step (1), concrete such as Figure 10 by the output result of Matlab solver Shown in:

C (), according to step (1), the theoretical center of circle A of calculated motor mounting flange is to sending out The vector of motor mounting flange theory center of circle C under the non-heeling condition of motivationConcrete as shown in table 6:

Table 6

(d) undetermined coefficient α_T(°) and β_T(°) is as shown in table 7:

Table 7

The anglec of rotation (undetermined coefficient) (°)	Numerical value
		αT	-0.0754
βT	-0.3987

(e)C_EBWith C_SCoordinate axes angle is as shown in table 8:

Table 8

F () setting angle is as shown in table 9:

Table 9

G () installation site is as shown in table 10:

Table 10

The optimum results that optimization method according to a kind of spacecraft orbit maneuver motor installation parameter obtains, utilizes This group installation parameter calculates following result, is shown in Table the disturbance torque during 11, each change rail, sees Fig. 8.

Table 11

As can be seen from Table 11, the disturbance torque problem produced when lighting a fire for orbit maneuver motor, this The bright novel design method proposed, component maximum and minima be respectively 0.79N.m and-0.79N.m, Both even can reach-8.88e-16 at sum, shows that solving precision can control at 1E-12Nm model In enclosing.

By can be seen that so that disturbance torque component maximum and minima numerical value phase during each igniting Deng, symbol is contrary, illustrates that electromotor installation parameter design comparison is reasonable, it is ensured that during each change rail The optimized design object of disturbance torque, from holocyclic system perspective, has reached to optimize the purpose of design.

The content not being described in detail in description of the invention belongs to the known of professional and technical personnel in the field Technology.

Claims (5)

1. a spacecraft orbit maneuver motor installation parameter optimization method, it is characterised in that step is as follows:

(1) spacecraft mechanical coordinate system C is set up_S, spacecraft translational coordination system C_S", spacecraft centroid sit Mark system C_C, engine body coordinate system C_EB, engine target coordinate system C_TWith electromotor mounting coordinate It is C_EI；

Described spacecraft mechanical coordinate system C_SInitial point O_SIt is positioned at the mechanical separation of satellite and carrier rocket In face, and overlap with the center of circle of origin reference location pin institute formational theory circle in mechanical separation face, X_SAxle is square Spacecraft east plate, Y is pointed to from zero_SAxle positive direction points to spacecraft south plate from zero, Z_SAxle meets the right-hand rule；

Described spacecraft translational coordination system C_S"By spacecraft mechanical coordinate system C_STranslation obtains, and spacecraft is put down Move coordinate system C_S"Zero be motor mounting flange theory center of circle A；

Described spacecraft centroid coordinate system C_CBy spacecraft mechanical coordinate system C_STranslation obtains, spacecraft matter Heart coordinate system C_CZero be positioned at spacecraft centroid；

Described engine body coordinate system C_EBZero be positioned at the theoretical circle of motor mounting flange Heart A, X_EBAxle positive direction and spacecraft mechanical coordinate system Z_SAxle positive direction is consistent, Y_EBAxle positive direction with Spacecraft mechanical coordinate system Y_SAxle negative direction is consistent, Z_EBAxle positive direction and spacecraft mechanical coordinate system X_S Axle positive direction is consistent；

Described engine target coordinate system C_TBy spacecraft translational coordination system C_S" rotate and obtain, electromotor Target-based coordinate system C_TZ_TAxle negative direction is to point to spout direction along electromotor theory geometrical axis；

Described electromotor mounting coordinate system C_EIFor engine target coordinate system C_TAround+X_TAxle rotates 180 ° And obtain, Z_EIThe positive direction of axle is to point to spout direction along electromotor theory geometrical axis；

(2) according to previously given thrust vectoring skew angle, thrust vectoringTraversing position angle beta, push away Force vectorDeflected position angle γ and traversing amount δ of thrust vectoring application point P, at engine body coordinate It is C_EBUnder, calculate motor power vector F_EBWith position of action point vector

(3) according to the position of action point vector in step (2)Motor power vector F_EB In conjunction with engine body coordinate system C_EBTo electromotor mounting coordinate system C_EITransformation matrix of coordinates Obtain at electromotor mounting coordinate system C_EIUnder position of action point vectorWith motor power vector F_EI；

(4) order is from satellite translational coordination system C_S"Transform to engine target coordinate system C_TProcess be: By satellite translational coordination system O_S"X_S"Y_S"Z_S"First around+X_S"Axle anglec of rotation α_T(°), further around Y_TAxle rotation β_T(°), α_T(°) and β_T(°) is undetermined coefficient；Determine electromotor mounting coordinate system C_EITo setting up spacecraft mechanical coordinate It is C_SSpin matrixWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；And at spacecraft machine Tool coordinate system C_SUnder, the theoretical center of circle A setting up motor mounting flange issues to engine tilts state The vector of motivation mounting flange center of circle CWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；

(5) vector obtained in step (4) is utilizedWith undetermined coefficient α_T(°) and β_TPass between (°) It is formula, determines at spacecraft mechanical coordinate system C_SUnder, O_STo motor mounting flange theory center of circle A's Vector

(6) spin matrix in step (4) is utilizedWith undetermined coefficient α_T(°) and β_TRelation between (°) Motor power vector F in formula integrating step (3)_EI, position of action point vectorAnd step (5) vector inObtain spacecraft mechanical coordinate system C_SLower electromotor thrust vectoring F_S、 Position of action point vectorWith undetermined coefficient α_T(°) and β_TRelational expression between (°)；

(7) utilize two groups of relational expressions in step (6) calculate during each igniting disturbance torque value M and Undetermined coefficient α_T(°) and β_TRelational expression between (°)；

(8) during each igniting, disturbance torque component maximum and minima numerical value are equal, and symbol Contrary is optimization aim, and the relational expression in solution procedure (7) obtains undetermined coefficient α_T(°) and β_T(°), And then obtain spin matrix

(9) according to spin matrixObtain spacecraft orbit maneuver motor setting angle and engine nozzle Installation site B.

A kind of spacecraft orbit maneuver motor installation parameter optimization method the most according to claim 1, It is characterized in that: the position vector in described step (3)With motor power vector F_EIMeter Calculation formula is as follows:

$F_{EI}=R_{EB}^M{F}_{EB}$

A kind of spacecraft orbit maneuver motor installation parameter optimization method the most according to claim 1, It is characterized in that: the motor power vector F in described step (6)_S, position of action point vector With undetermined coefficient α_T(°) and β_TRelational expression between (°) is as follows:

$F_S=R_{E1}^S{F}_{EI}$

In formula,

A kind of spacecraft orbit maneuver motor installation parameter optimization side the most according to claim 1 and 2 Method, it is characterised in that: the relational expression in step (7) is as follows:

In satellite machinery coordinate system C_SUnder, the position of centroid of satellite during precise tracking igniting Vector.

A kind of spacecraft orbit maneuver motor installation parameter optimization method the most according to claim 1, its Be characterised by: according to the result in step (7), during each igniting disturbance torque component maximum and Minima numerical value is equal, and symbol is optimization aim on the contrary, is embodied as equation below:

max(M_x,M_y,M_Z)_{I=1, n}+min(M_x,M_y,M_Z)_{I=1, n}=0

Wherein, M_x,M_y,M_ZFor the disturbance torque three axle components under spacecraft centroid coordinate system.