CN103488814B

CN103488814B - Closed loop simulation system suitable for controlling attitude of reentry vehicle

Info

Publication number: CN103488814B
Application number: CN201310359916.1A
Authority: CN
Inventors: 李惠峰; 余光学; 李昭莹
Original assignee: Beihang University
Current assignee: Beihang University
Priority date: 2013-08-16
Filing date: 2013-08-16
Publication date: 2017-02-15
Anticipated expiration: 2033-08-16
Also published as: CN103488814A

Abstract

The invention discloses a closed loop simulation system suitable for controlling the attitude of a reentry vehicle. The system comprises a subsystem (100) for modeling the aerodynamic shape of the reentry vehicle, a subsystem (200) for building an attitude control model and a subsystem (300) of an air vehicle representation model. Parameter setting and simulation relation resolution of a simulation environment are realized by building the attitude control model subsystem (200), so that demonstration of simulation results under different tasks is realized. According to the closed loop simulation system suitable for controlling the attitude of the reentry vehicle disclosed by the invention, the control law design defect of a design phase can be picked via representation of input and output quantities and attitudes in different states, so that the safety of a control strategy under uncertain factors and external interference and the control ability of the air vehicle can be verified, the control characteristics of the air vehicle are analyzed, and optimization parameters are provided for an overall design.

Description

A kind of closed-loop simulation system being applied to reentry vehicle gesture stability

Technical field

The present invention relates to a kind of analogue system of gesture stability, more particularly, refer to that one kind is applied to reentry vehicle The closed-loop simulation system of gesture stability.

Background technology

Manufactaring is in national defence, the civilian new world opening application.It is little with risk, efficiency high, not climate Condition and the restriction in space, substantially reduce development cost and proving period and are used widely.

《Guided missile and carrying space technology》In the 5th phase in 1994, total 211st phase, disclose entitled " Spacecraft reentry track It is in progress with controlling ".Write in literary composition：Reentry vehicle can be divided into space flight（Empty sky）Aircraft,（Manned）Airship, satellite, bullet etc., also Can be divided into manned with not manned reentry vehicle, also can be divided into reusable with not reproducible use reentry vehicle etc..

Reentry vehicle needs, under the conditions of having and greatly initially reentering kinetic energy and potential energy, steadily safely to direct into set Touchdown area, again so that overload, dynamic pressure and Aerodynamic Heating is in allowed band, this is to aircraft manufacturing technology system simultaneously Design proposes high requirement.Attitude control system controls the actuator of aircraft, and then manipulates aircraft flight, in work Journey should be used, it is necessary to assure the safety that certain control redundancy just can complete aircraft under uncertain factor and external interference flies OK.

" world's guided missile and space flight " interim disclosed written by Lin Laixing at 1989 the 7th《Make the flight of god's space shuttle Control system》.Fig. 1 of the document describes in detail and makes god's track and gesture stability structural principle.

Using computer and simulation software（As Matlab R2008a-Simulink）Emulation is an important link, The engineering development cycle can not only be shortened, and greatly save checking flight cost.

Content of the invention

In order to realize the Simulation Control to reentry vehicle attitude, one aspect of the present invention adopts three-dimensional software to ablated configuration The aerodynamic configuration of device is modeled, and on the other hand constructs attitude control law model and aircraft characterization model.

A kind of closed-loop simulation system being applied to reentry vehicle gesture stability of the present invention, this system includes to reentering The subsystem that the aerodynamic configuration of aircraft is modeled（100）, build Attitude control model subsystem（200）With aircraft table Levy the subsystem of model（300）；Described structure Attitude control model subsystem（200）Include atmospheric environment model unit （201）, attitude command signal generating unit（202）, flight vehicle aerodynamic evaluation unit（203）, slow circuit controls restrain unit（204）, fast Circuit controls restrain unit（205）, complex controll allocation unit（206）, pneumatic rudder face control allocation unit（207）, RCS control point Join unit（208）, flight vehicle aerodynamic data base（209）With control integrated unit（210）；

The simulation flow that the closed-loop simulation system of described reentry vehicle gesture stability is carried out is：

（A）Analogue system initializes；Global variable is N=0, arranges simulation step length h, reenters initial time t₀=0, initially Attitude state y₀=[τ₀λ₀H₀V₀χ₀γ₀α₀β₀μ₀p₀q₀r₀]^T；And all control surface deflection angles of aircraft are zero （δ_e,δ_a,δ_g,δ_f）, RCS（16 thruster status commands are 0）Do not work；12 of initialization dummy vehicle unit 301 State, initialization RCS controls the RCS control of allocation unit 208 to be assigned as zero；Initialize pneumatic rudder face and control allocation unit 207 Pneumatic rudder control and be assigned as zero；It is zero that initialization controls the pneumatic control surface deflection of integrated unit 210；By dummy vehicle unit 12 status informations input Atmospheric models units 201 of 301 initialization, attitude command unit 202, pneumatic estimation unit 203, Slow circuit controls rule unit 204, fast loop control law unit 205, complex controll allocation unit 206, pneumatic rudder face controls distribution Unit 207, aerodynamic data library unit 209；

（B）According in the elevation information input atmospheric environment model unit 201 of dummy vehicle unit 301；Air Environmental model unit 201 by output information input to attitude command unit 202, pneumatic estimation unit 203, complex controll distributes Unit 206, pneumatic rudder face controls allocation unit 207, aerodynamic data library unit 209, controls integrated unit 210；

（C）Aerodynamic data library unit 209 is believed according to the atmospheric environment of the atmospheric environment model unit 201 of current input Breath, dummy vehicle unit 301 attitude information and the control inclined deflection information of integrated unit 210 rudder calculate output aerodynamic data information To control integrated unit 210；

（D）Pneumatic estimation unit 203 according to the atmospheric environment information of the atmospheric environment model unit 201 of current input, Dummy vehicle unit 301 attitude information and control integrated unit 210 control surface deflection information by calculate export pneumatic force information to Slow circuit controls rule unit 204 and aerodynamic moment coefficient estimation control allocation unit 207 and control integrated unit to pneumatic rudder face 210；

（E）Attitude command unit 202 is believed according to Atmospheric models unit 201 information of input and dummy vehicle unit 301 Breath calculates output and restrains unit 204 to slow circuit controls；

（F）Slow circuit controls are restrained unit 204 and are believed according to attitude command unit 202 information of input, pneumatic estimation unit 203 Breath and dummy vehicle unit 301 information calculate output to fast loop control law unit 205；

（G）Fast loop control law unit 205 restrains unit 204 information and dummy vehicle list according to the slow circuit controls of input Unit 301 information calculates output to complex controll allocation unit 206；

（H）Complex controll allocation unit 206 is according to Atmospheric models unit 201 information inputting, fast loop control law unit 205 information and dummy vehicle unit 301 information calculate output and control allocation unit 207 and RCS to control distribution to pneumatic rudder face Unit 208；

（I）Pneumatic rudder face controls allocation unit 207 according to Atmospheric models unit 201 information inputting, dummy vehicle list First 301 partial informations and the integrated unit 210 control surface deflection information that controls, comprehensive it is derived from the pneumatic of complex controll allocation unit 206 The aerodynamic moment coefficient estimation information of torque command information and pneumatic estimation unit 203 calculates output to control integrated unit 210；

（J）RCS controls allocation unit 208 according to complex controll allocation unit 206 information of input and to control integrated unit 210 partial informations calculate output to control integrated unit 210；

（K）Control integrated unit 210 according to Atmospheric models unit 201 information inputting, pneumatic estimation unit 203 aerodynamic force Moments estimation information, dummy vehicle unit 301 partial information and pneumatic rudder face control allocation unit 207, RCS to control allocation unit 208 information, aerodynamic data library unit 209 information calculates output information to dummy vehicle unit 301；And the rudder by real-time measurement Deflecting facet information output controls allocation unit 207 and pneumatic Database Unit 209 to pneumatic estimation unit 203, pneumatic rudder face, will RCS control and compensation moment information is passed to RCS and is controlled allocation unit 208；

（L）Dummy vehicle unit 301 is according to the control integrated unit 210 information updating state of flight of input, and exports , to Atmospheric models unit 201, output speed status information is to attitude command unit 202 and complex controll allocation unit for elevation information 206, export attitude angle information to slow circuit controls rule unit 204 and fast loop control law unit 205, output par, c information is to gas Dynamic estimation unit 203, pneumatic rudder face controls allocation unit 207 and pneumatic Database Unit 209；Export height and velocity information to Emulation whether end unit 302；

（M）According to input, from emulation, whether the height of end unit 302 and velocity information judge to emulate and terminate. If emulation terminates, show simulation result, and ends with system emulation；If emulation does not terminate, proceed to step（B）In repeat to imitate Very, until it reaches emulation termination condition.

The advantage that the present invention is applied to the closed-loop simulation system of reentry vehicle gesture stability is：

1. gesture stability modeling of the present invention establishes the ablated configuration attitude controller verification platform of reentry vehicle With flight validation platform, can be used in the ablated configuration process simulation of bullet, airship or space shuttle etc., there is versatility；

2. analogue system of the present invention can be as needed in the unit signal input output relation defining, by concrete Aircraft generate required corresponding function, each submodule has extensibility；

3. each submodule in analogue system of the present invention can be directly changed into corresponding algorithm, and submodule can be used for Other control system, has portability.

Brief description

Fig. 1 is the structured flowchart of the closed-loop simulation system of the reentry vehicle gesture stability of the present invention；Fig. 2A is pneumatic outer The front view of shape；Fig. 2 B is the top view of aerodynamic configuration；Fig. 2 C is the upward view of aerodynamic configuration；Fig. 2 D is that the right side of aerodynamic configuration regards Figure；Fig. 2 E is the left view of aerodynamic configuration；Fig. 2 H is the coordinate schematic diagram of reentry vehicle and geographic coordinate system；Fig. 2 I is to reenter Flight state variable defines schematic diagram with corresponding coordinate system；Fig. 3 is the closed loop of the reentry vehicle gesture stability of the present invention The control flow chart of analogue system；Fig. 3 A is the control flow chart that the present invention controls fusion part；Fig. 4 is the 16 of reentry vehicle The distribution schematic diagram of individual thruster；Fig. 4 A is the control mode schematic diagram of the split axle control of thruster；Fig. 4 B is dividing of thruster The control mode schematic diagram that shelves control.

Specific embodiment

Below in conjunction with accompanying drawing, the present invention is described in further detail.Shown in Figure 1, one kind of the present invention is suitable for In the closed-loop simulation system of reentry vehicle gesture stability, this system includes the aerodynamic configuration to reentry vehicle and is modeled Subsystem 100（Referred to as aerodynamic configuration builds subsystem）, build the subsystem 200 of Attitude control model and aircraft characterizes The subsystem 300 of model.

（One）Aerodynamic configuration builds subsystem 100

In the present invention, according to three-dimensional software（As Solidworks2006）Draw out the aerodynamic configuration of aircraft, such as scheme Shown in 2A～Fig. 2 G.For convenience attitude explanation is carried out to reentry vehicle, quote multiple coordinate systems（As shown in Fig. 2 H, Fig. 2 I） It is described in detail.

Characterize to realize the attitude to aircraft, quote body axis system O-x_by_bz_b, velocity coordinate system O-x_ay_az_a, boat Mark coordinate system O-x_ky_kz_k, local vertical coordinate system O-x_gy_gz_g, the earth's core equator inertial coodinate system S_i-Ex_iy_iz_iWith the rotation of the earth's core equator Turn coordinate system S_e-Ex_ey_ez_e.Define attitude angle and the flight-path angle of aircraft by four described coordinate systems.

In the present invention, the definition of coordinate system is：

(1) the earth's core equator inertial coodinate system S_i-Ex_iy_iz_i：The earth's core equator inertial coodinate system S_i-Ex_iy_iz_iAs inertia base Standard, E is in earth center point（The earth's core）.Ez_iAxle, perpendicular to earth equatorial plane, points to the arctic；x_iAxle and y_iAxle plane under the line Interior, wherein x_iAlong the direction pointing to the first point of Aries, this point is earth equatorial plane and infinite point on the intersecting lens of ecliptic plane to axle, that is, The direction of moment in Spring Equinox Earth-Sun line.The earth's core equator inertial coodinate system S_i-Ex_iy_iz_iForm right hand rectangular coordinate system.

(2) the earth's core equator rotating coordinate system S_e-Ex_ey_ez_e：The earth's core equator rotating coordinate system S_e-Ex_ey_ez_eConnect firmly with the earth, E In earth center point（The earth's core）.z_eAxle, perpendicular to earth equatorial plane, points to the arctic；x_eAxle and y_eAxle is under the line in plane, wherein x_eAxle is along the intersecting lens of equatorial plane and Greenwich meridian plane.The earth's core equator rotating coordinate system S_e-Ex_ey_ez_eForm the right hand Rectangular coordinate system.

(3) local vertical coordinate system O-x_gy_gz_g：Local vertical coordinate system O-x_gy_gz_gIt is connected with body, O is in aircraft At barycenter（Barycenter）；Plane Ox_gz_gIt is local vertical plane, x_gAxle points to direct north；z_gAxle is perpendicular to local level vertical Downwardly directed the earth's core；y_gAxle points to east in the horizontal plane.Local vertical coordinate system O-x_gy_gz_gForm right hand rectangular coordinate system.

(4) body axis system O-x_by_bz_b：Body axis system O-x_by_bz_bIt is connected with body, O is at the barycenter of aircraft（Matter The heart）；x_bAxle points to head in aircraft symmetrical plane and parallel to the design axis of aircraft；y_bAxle is perpendicular to aircraft pair Plane is claimed to point to body right；z_bAxle in aircraft symmetrical plane, with x_bAxle vertically and points to below body.Body axis system O-x_by_bz_bForm right hand rectangular coordinate system.

(5) velocity coordinate system O-x_ay_az_a：Velocity coordinate system O-x_ay_az_aIt is connected with body, O is at the barycenter of aircraft（Matter The heart）；x_aAxle is consistent with the coincidence of flight speed V；z_aAxle in aircraft symmetrical plane, with x_aAxle vertically and points to below body；y_a Axle is perpendicular to Ox_az_aPlane simultaneously points to body right.Velocity coordinate system O-x_ay_az_aForm right hand rectangular coordinate system.

(6) flight path axis system O-x_ky_kz_k：Flight path axis system O-x_ky_kz_kIt is connected with body, O is at the barycenter of aircraft（Matter The heart）；x_kAxle is consistent with the coincidence of flight speed V；z_kAxle position in the vertical guide comprising flight speed V, with x_kAxle vertically and refers to Downwards；y_kAxle is perpendicular to Ox_kz_kPlane, it points to and determines according to the right-hand rule.Flight path axis system O-x_ky_kz_kForm the right hand straight Angular coordinate system.

The earth's core E and aircraft barycenter O can determine away from R in geographic logitude τ of earth's surface intersection point and latitude λ, the earth's core in fig. 2 The position that aircraft is located, is defined as follows：

The earth's core is away from R：Aircraft barycenter O is apart from the distance of the earth's core E.

Azimuth τ：The angle that projection EO ' from the first point of Aries is eastwards to radius vector EO under the line plane is turned over, referred to as Azimuth, this is azimuthal to be entered as longitude.

Angle of site λ：Projection EO ' on R under the line plane for the earth's core northwards goes to the angle of radius vector EO from equatorial plane, letter The referred to as angle of site, this angle of site be entered as latitude.

Intersection point O '：The intersection point away from the projection in earth equatorial plane for the R and earth surface for the earth's core.

Described attitude angle is by body axis system O-x in fig. 2_by_bz_bWith velocity coordinate system O-x_ay_az_aBetween angle Transformational relation obtains.Attitude angle includes angle of attack and sideslip angle beta.

Angle of attack refers to speed V of the aircraft projection x ' on the longitudinally asymmetric face ABCD of aircraft_aWith body axis system O-x_by_bz_bX_bAngle between axle.As this projection x '_aIn x_bDuring the downside of axle, angle of attack is designated as positive-angle；As this projection x '_a In x_bDuring the upside of axle, angle of attack is designated as negative angle.

Sideslip angle beta refers to the angle between speed V of aircraft and the longitudinally asymmetric face ABCD of aircraft.When speed V When the right side of longitudinally asymmetric plane ABCD of aircraft, β yaw angle is designated as positive-angle in direction；When the direction of speed V is in flight During the left side of longitudinally asymmetric plane ABCD of device, β yaw angle is designated as negative angle.

Described flight-path angle is flight path axis system O-x in fig. 2_ky_kz_kWith local vertical coordinate system O-x_gy_gz_gBetween angle Degree transformational relation obtains.Flight-path angle includes flight path inclination angle γ, flight path azimuthangle χ and angle of heel μ.

Flight path inclination angle γ refers to speed V and the planar delta Ox of aircraft_gy_gBetween angle.When speed V is put down in triangle When above face, flight path inclination angle γ is designated as positive-angle；When speed V is below planar delta, flight path inclination angle γ is designated as negative angle Degree.Described planar delta Ox_gy_gRefer to local vertical coordinate system O-x_gy_gz_gMiddle connection Ox_g,Oy_g,x_gy_gThe plane being formed.

Flight path azimuthangle χ refers to speed V of aircraft in planar delta Ox_gy_gOn projection x '_kWith local vertical coordinate system O-x_gy_gz_gX_gAngle between axle.As projection x '_kIn x_gDuring the right side of axle, flight path azimuthangle χ is designated as positive-angle；Work as projection x′_kIn x_gDuring the left side of axle, flight path azimuthangle χ is designated as negative angle.

Angle of heel μ refers to aircraft around flight path axis system O-x_ky_kz_kX_kThe angle that axle is rotated.Wing when aircraft Occur left wing upwards, right flank downward rolling when, angle of heel μ is designated as positive-angle；When the wing of aircraft occurs Left Wind Down, the right side During rolling in flapwise, angle of heel μ is designated as negative angle.

In the present invention, application aerodynamic configuration structure subsystem 100 to carry out configuration to reentry vehicle, illustrates that Different attitudes（I.e. input and output instruction）Lower real time demonstration goes out the scene simulation of reentry vehicle.

（Two）Build the subsystem of Attitude control model

Shown in Figure 3, in the present invention, build Attitude control model subsystem 200 and include atmospheric environment model list Unit 201, attitude command signal generating unit 202, flight vehicle aerodynamic evaluation unit 203, slow circuit controls rule unit 204, the control of fast loop System rule unit 205, complex controll allocation unit 206, pneumatic rudder face control allocation unit 207, RCS to control allocation unit 208, fly Row device aerodynamic database 209 and control integrated unit 210.In the present invention, by building attitude controller model subsystem 200 Realize parameter setting, the emulation relation decomposing of simulated environment, thus the simulation result demonstration under obtaining different task.

（1）Atmospheric models unit 201

In the present invention, flying height according to residing for aircraft flight for the Atmospheric models unit 201, can obtain aircraft Temperature under being located highly, the velocity of sound, pressure and density.For aircraft dynamic pressure, aerodynamic force, aerodynamic moment resolving.

It is with reference to United States standard atmosphere for the output variable number in Atmospheric models unit 201（1976）Obtained by.The U.S. Standard atmosphere（1976）It is in United States standard atmosphere by U.S.National Oceanic Atmospheric Administration, National Aeronautics and Space Administration, air force Formulate on the basis of versions in 1962 and United States standard atmosphere version in 1966.United States standard atmosphere（1976）Carried using chart Supply following parameter with the profile of height：Temperature, pressure, density, the velocity of sound, dynamic and kinematic viscosity, pyroconductivity etc..This air Model is not modeled to wind field, and what the temperature that is given, density, pressure data represented is meansigma methodss.Here, being according to height In United States standard atmosphere（1976）Enter row interpolation in table, obtain temperature, pressure, density and the velocity of sound of residing height and position.

（2）Attitude command unit 202

In the present invention, attitude command unit 202 is according to aerial mission, reentry vehicle return be subject to severe overload, Rate of heat flow, the restriction of dynamic pressure, attitude command be given will in conjunction with aerial mission, the characteristic of aircraft and control ability come comprehensive to Go out attitude command, finally the tracking to attitude command is completed by attitude control system, so, aircraft could safety return.? This, by attitude command unit 202 integrated flight device speed and the highly lower velocity of sound that is located, to provide attitude angle and to instruct.

Attitude command unit 202 1 aspect receives velocity information V and air mould from dummy vehicle unit 301 output Velocity of sound information a of type unit 201 output_{The velocity of sound}；On the other hand, to V and a receiving_{The velocity of sound}Calculate Mach number Ma, thus exporting 3 Attitude command information.

In the present invention, attitude command unit 202 according to aircraft speed and be located highly under the velocity of sound, obtain horse Conspicuous numberAccording to Mach number Ma and flight speed V, angle of attack is provided by logical judgment_c, sideslip angle beta_c=0 and tilt Angle μ_cInstruction.

Aircraft, during reentering, instructs α to the angle of attack_cFollow the tracks of for controlling the Aerodynamic Heating that reenters and energy management. Meanwhile, in order to limit the heat flux of body surface, calm yaw angle, therefore yaw angle instruction β are needed_cIt is zero it is ensured that flight is pacified Entirely.μ is instructed to angle of heel_cFollow the tracks of for adjustment flight vertical journey, horizontal journey so that aircraft enters into predetermined energy management window Mouthful.Attitude angle instruction is given online by Guidance Law, and Guidance Law is not belonging to attitude control system, and here is refused to design.Here The attitude command logic being given is that attitude command will be protected simultaneously in order to the design of access control device can follow the tracks of upper attitude command Demonstrate,prove aircraft according to during attitude command flight, in the reentry corridor that dynamic pressure, overload and rate of heat flow are allowed.Different are reentered Aircraft, the numerical value of attitude command can be different, and attitude command given herein disclosure satisfy that designed high lift-drag ratio again Enter the demand of aircraft manufacturing technology system, the specific numerical value in attitude command will be according to the pneumatic of reentry vehicle and process Constraint and reentry trajectory are adjusting.

（3）Pneumatic estimation unit 203

Pneumatic estimation unit 203 is used for estimating aerodynamic force suffered by aircraft, aerodynamic moment coefficient, pneumatic by resolve Power be supplied to attitude control law module 204 for resolve compensate the nonlinear terms controlling, carry out the rule resolving of slow circuit controls.To estimate The aerodynamic moment coefficient of meter is supplied to pneumatic rudder face and controls allocation unit 207 and control integrated unit 210.Unit 207 are used In the control allocation matrix resolving pneumatic rudder face.It is used for the resolving of RCS Torque Control compensation vector in Unit 210.Return slow During road control law 204 resolves, need to calculate nonlinear compensation item according to the flight vehicle aerodynamic power of estimation, control in pneumatic rudder face In the design of distribution 207, the flight vehicle aerodynamic moment coefficient according to estimation is needed to estimate to calculate pneumatic control surface deflection instruction.? Controlling in the calculating of RCS aerodynamic moment compensation vector merge 210 needs to resolve RCS torque compensation according to aerodynamic moment coefficient Amount.

In the present invention, pneumatic estimation unit 203 1 aspect receive from 201 place highly under velocity of sound information, institute Atmospheric temperature information under highly, the atmospheric pressure information under being located highly, the atmospheric density information under being located highly, 301 Aircraft altitude information, velocity information, angle of attack information, sideslip angle information, rolling angular rate information, pitch rate information, Rolling angular rate information.210 aircraft left elevon rudder believes one side only breath, and right elevon rudder believes one side only breath, rudder information, body Flap-type rudder believes one side only breath；On the other hand, the lift suffered by aircraft, resistance are estimated after being resolved, side force information output gives 204 lists Unit；And the pneumatic rolling moment coefficient (6) that will estimate, pneumatic pitching moment coefficient (5), pneumatic yawing moment coefficient (6 ) amount to 17 export to Unit 207 and Unit 210.

Dense aerodynamic is from continuous media supposition, although gas is by individual other molecular, But in general, the mean free path of moleculeCompare much smaller with characteristic length Le of the problem studied, therefore, It is approximate well that gas is assumed to continuous media.But, thin out with air, when mean free path and the institute of molecule When the ratio of characteristic length Le of problem of research is close to and above unit value, the feature that gas is made up of individual molecules will manifest Out, the hypothesis of continuous media is just no longer applicable, now just need to be solved with superaerodynamics.The division in flowing field Be according to Kn number Lai.

Define aerodynamic characteristics parameter Michel Knuysen（Knudsen）Number Kn is to flow molecule mean free pathWith studied a question Characteristic length Le ratio For the molecule mean free path of air, characteristic length Le is taken as aircraft Total length.

By Knudsen number Kn, flow region is divided into continuous stream area, transition flow region and free-molecule regime.When Kn >=10 When, flow region is free-molecule regime；When Kn≤0.001, flow region is continuous stream area；As 0.001 ＜ Kn ＜ 10, Flow region is transition flow region, all inapplicable in the hypothesis of this regional air continuum Model and free molecule flow.With gram The amplitude of Knudsen number Kn change, in order to take into full account the impact information to flight vehicle aerodynamic for the air rarefied content, the pneumatic system of matching Choose during numberIt is characterized one of parameter, for the aerodynamic data of various height, according to Mach number, the angle of attack, break away Angle, the drift angle of pneumatic rudder face being fitted into regard to the angle of attack, yaw angle, Mach number, pneumatic rudder partially andFunction.Therefore Obtain the angle of attack C rising under force environment taking the logarithm in Knudsen number Kn_L,α, left rudder inclinedRight standard rudder is inclinedRudder is inclinedBody flap rudder is inclinedAngle of attack C under the resistance environment that Knudsen number Kn takes the logarithm_D,α, left rudder inclinedRight standard rudder is inclinedRudder is inclinedBody flap rudder is inclinedThe angle of attack under the Mach number environment that Knudsen number Kn takes the logarithmLeft Rudder is inclinedRight standard rudder is inclinedRudder is inclinedBody flap rudder is inclinedPitch rate C_{M, q}Take the logarithm in Knudsen number Kn Side force environment under yaw angle C_Y,β, left rudder inclinedRight standard rudder is inclinedRudderUnder the liter force environment that matching is estimated Yaw angle C_l,β, left rudder inclinedRight standard rudder is inclinedRudder is inclinedRolling angular speed C_l,p, yaw rate C_l,r；Matching Yaw angle C under the Mach number environment estimated_n,β, left rudder inclinedRight standard rudder is inclinedRudder is inclinedRolling angular speed C_n,p, yaw rate C_n,r.

Restraining unit 204 due to slow circuit controls needs to receive the lift estimatedThe resistance estimatedThe side force estimatedTherefore need to resolve the lift coefficient estimated in pneumatic estimation unit 203The resistance coefficient estimatedEstimate Lateral force coefficientPneumatic estimation unit 203 needs to export the correlative of estimation to Unit 207 and Unit 208.

According to the theory of molecular motion, the viscosity of gasDensity p, warm-up movement average speedAnd molecule is average certainly By journeyBetween exist relation beWherein For air constant, its value takesT is the temperature under being located highly；WhereinViscosityUsing Sa Seraing（Sutherland）Formula calculates：I.e.Wherein, temperature T when initial₀= 288.15K, gas viscosity coefficient when initialC_TemperatureFor thermal constant, value is 110.4K.Then There is Knudsen number

Aerodynamic database according to aircraft（Set up by aerodynamic data library unit 209）Data, gained is pneumatic Data carries out higher order polynomial and is fitted, and obtains linear expression-form, and the design for being controlled distributing is fast with pneumatic Speed resolves.Aerodynamic Coefficient C by aircraft matching_L,α,C_D,α,C_Y,β, C_l,p, C_l,r；C_m,α,C_m,q；C_n,β,C_{N, p}, C_{N, r}It is loaded into and fly control On computer, according to receiving Aircraft Angle of Attack information, sideslip angle information, the speed of aircraft and the velocity of sound under being located highly turn Mach number obtained by changing, aerodynamic characteristics parameterAnd aircraft left elevon rudder believes one side only breath, right elevon rudder believes one side only Breath, rudder information, body flap-type rudder believes one side only breath resolving aerodynamic force-lift, resistance, side force, with aerodynamic moment coefficient.

B is that the span of the wing of aircraft is long, and c is the mean aerodynamic chord of the wing of aircraft.

In the present invention, calculate dynamic pressure nowI.e.

In the present invention, carry out synthesis by pneumatic with atmospheric environment, resolve the control law of fast loop control unit.By big compression ring The atmosphere data of border input, according to current aircraft speed, obtains dynamic pressure.Then there is the aircraft that aerodynamic prediction module provides Pneumatic estimation and the parameter of aircraft itself, obtain the aerodynamic force suffered by aircraft and estimate, export Unit 204.According to work as Dynamic pressure residing for front aircraft and aircraft parameters, estimate the aerodynamic force suffered by aircraft,Finally, by obtained lift estimated informationResistance estimated informationAnd side Power estimated informationExport to Unit 204, by C_l,β,C_l,p, C_l,r；C_m,α, C_m,q；C_n,p, C_n,rExport to Unit 207 and Unit 210.

（4）Slow circuit controls rule unit 204

In the present invention, slow circuit controls restrain unit 204 according to current atmospheric environment and aerodynamic estimation value, and work as The attitude angle of front feedback, generates attitude angular rate instruction output.

In the present invention, slow circuit controls are restrained unit 204 and are received aircraft altitude information, the velocity information being derived from 301, boat Mark obliquity information and angle of attack information, sideslip angle information, tilt angle information, rolling angular rate information, pitch rate information, rolling Angular rate information.According to 202 angle of attack command information, yaw angle command information and angle of heel command information, comprehensively estimate from 203 Lift suffered by meter aircraft, resistance, side force information resolves slow circuit controls rule, by obtained attitude angular rate instruction output To Unit 205.For the elevation information of aircraft, rolling angular rate information, pitch rate information and yawrate information Do not deal with.

Slow circuit controls restrain unit 204 by from 301 attitude angle information and velocity information, substitute into from 203 pneumatic Lift, the side force estimated information nonlinear compensation item in control lawIn.

Wherein M is the quality of aircraft, and g is acceleration of gravity.

The control matrix of slow state

In order to realize attitude command Ω_cTracking, according to time-scale separation principle, first to required for attitude angle Ω dynamic design Attitude angular rate instruction ω_c, referred to as slow circuit controls rule；Control moment required for attitude angular rate dynamic design refers to again Make M_c, referred to as fast loop control law.The internal ring in fast loop Shi Man loop, when the two frequency band differs 3 to 5 times, in design and analysis Fast loop dynamic characteristic approximately can be ignored during slow loop.

Thus obtaining slow circuit controls rule ω_cFor：

Wherein, attitude angleAttitude angle instructsThe instruction reciprocal of attitude angle

Ω_c- Ω is attitude error,For the inverse of attitude error,For the inverse of the angle of attack,For the inverse of yaw angle,For the inverse of angle of heel, k_p,sFor attitude error ratio diagonal matrix, k_i,sFor attitude by mistake Difference-product divides diagonal matrix, k_d,sFor attitude error differential diagonal matrix, i.e. three-dimensional square formation, the unit on each diagonal is all higher than Zero, the unit on off-diagonal is equal to zero.

（5）Fast loop control law unit 205

In the present invention, fast loop control law 205 instructs ω according to the attitude angular rate that slow loop provides_cWith being used to of aircraft Property parameter ω, and feedback attitude angular rate state, generate control moment instruction Wherein, ω_c- ω is attitude angular rate error,For Integration from the attitude angular rate error of initial time to current time t.Fast loop control law 205 adopts LQR method, carries out fast The control law of road controller resolves.

In the present invention, fast loop control law 205 receives aircraft altitude information, the velocity information being derived from 301, and flight path inclines Angle information and angle of attack information, sideslip angle information, tilt angle information, rolling angular rate information, pitch rate information, roll angle speed Rate information.According to 204 roll angle rate instruction information, pitch rate command information and yawrate command information, comprehensive Close the aircraft rolling angular rate information from 301, pitch rate information, rolling angular rate information resolves fast circuit controls Rule, by obtained control moment instruction output to Unit 206.For the elevation information of aircraft, velocity information, flight path angle Information, angle of attack information, sideslip angle information and tilt angle information do not deal with.

Fast loop control law 205 is according to from 301 attitude angular rate information and the rotary inertia information of aircraft, construction Fast loop sytem matrix A and fast circuit controls matrix B：

It is around x_bThe rotary inertia of axle,It is around y_bThe rotary inertia of axle,It is around z_bThe rotary inertia of axle,For x_bAxle and z_bThe product of inertia of axle.

In the present invention, fast loop control law 205 restrains the attitude angular rate instruction of unit 204 output according to slow circuit controls ω_cDynamic with fast state Represent control moment, and control moment is obtained based on method for optimally controlling Command M_c.

Optimum control is a very active field of scientific study, has solid theoretical basiss and is widely applied the back of the body Scape, also achieves very good achievement in aerospace field.For linear quadratic regulator（LQR, Linear Quadratic Regulator）Problem, optimal state feed-back control device is state regulator, and the tracking that complete to instruct needs to move fast state StateCarry out augmentation, and dynamic to fast stateIntroduce the integrating state of error

In the present invention, the fast loop sytem matrix after augmentation isI₃Unit matrix for 3 × 3, augmentation Fast circuit controls matrix afterwards isFast state after augmentation is

In the present invention, carry matrix equation according to multitude's card and solve fast loop positive definite symmetric matrices G.

For reentry vehicle, fast state is dynamicHave different under different flying conditions Fast loop sytem matrix A.The dynamic consecutive variations of A in flight envelope, its corresponding feedback gain matrix K is also dynamically continuous Change, according to real-time attitude angular rate p, the different feedback gain matrix K of q, r design.According to M_c=-K η obtains controling power Square command information, exports to Unit 206

It is critical only that of LQR controller design chooses the nonnegative definite symmetrical matrix Q of suitably fast loop index and fast loop The parameter of the positive definite symmetric matrices R of index, the fast response characteristic of control to be improved, then can increase the weights of respective element in Q； Will the amplitude that controlled quentity controlled variable effectively suppressed and its energy expenditure causing, then can improve the weights of respective element in R；The choosing of Q and R Selecting is mutual restriction, needs compromise to consider.Optimum control can obtain linear state feedback control rule, constitutes closed loop optimum anti- Feedback controls it is easy to Project Realization, thus it is actual to be widely used in engineering.

（6）Complex controll allocation unit 206

Reentry vehicle returns from track, and dynamic pressure is gradually increased.When dynamic pressure is less, the execution efficiency of pneumatic rudder face is low, Control ability is not enough it is necessary to complete gesture stability using RCS.With the increase of dynamic pressure, pneumatic rudder intervenes control, RCS successively Progressively exit, gesture stability task is jointly completed with pneumatic rudder by RCS.When flying height reduces, after dynamic pressure increases further, gas Dynamic rudder completely intervenes, and last RCS exits control, only completes gesture stability by pneumatic rudder.Simultaneously participate in control in RCS with pneumatic rudder When, according to dynamic pressure border, control authority is transitted linearly to pneumatic rudder by dynamic pressure from RCS.Design RCS is distributed with the control of pneumatic rudder Strategy, thus realize control moment command M_cDistribute to pneumatic rudder face and RCS.

Complex controll allocation unit 206 receives aircraft altitude information, velocity information and the angle of attack information being derived from 301.Root According to rolling moment command information, pitching moment command information and the yawing command information of 205 inputs, comprehensively it is derived from 201 Atmospheric environment information, the complex controll distribution designing the pneumatic rudder face of RCS/ resolves, and obtains distributing to pneumatic rudder face control distribution list The pneumatic rudder control moment information of unit 207 and the RCS control moment command information of RCS control allocation unit 208.For being derived from 201 temperature information, velocity of sound information, atmospheric pressure information and the elevation information of aircraft from 301, angle of attack information are not made to locate Reason.

According to dynamic pressure border, press pressure linear transitions, realize the control allocation strategy of RCS and pneumatic manoeuvring systemWherein, k_{Flexible strategy}For weight coefficient, For pneumatic Rudder face starts to intervene dynamic pressure during gesture stability,Exit dynamic pressure during gesture stability, the control of RCS and pneumatic rudder face for RCS Authority is according to dynamic pressureIt is allocated and coordinate.AndRespectively distribute to the controling power of RCS and pneumatic manoeuvring system Square.IfIt is allocated to the control moment instruction of RCS rolling, pitching and jaw channel respectively, It is allocated to the control moment instruction of pneumatic rudder rolling, pitching and jaw channel, M respectively_R、M_AIt is respectively RCS thruster and pneumatic The manipulation of physical moment that rudder face produces.

In the present invention, the control moment distributing to RCS instructs and isThe RCS control moment producing in real time Instruct and beThe control moment distributing to pneumatic rudder face instructs and isThe pneumatic rudder face producing in real time Control moment instructs and is

（7）Pneumatic rudder face controls allocation unit 207

In the present invention, the control moment of pneumatic rudder system is instructedIt is mapped to angle of rudder reflection instruction δ_c, that is, complete gas The control distribution of dynamic steerable system.Wherein, For basic Control moment vector, M_A,δFor controlling allocation matrix.

Wherein,S is the area of reference of aircraft wing, and unit is m²；B is the wing of the wing of aircraft Length, unit is m；C is the mean aerodynamic chord of the wing of aircraft, and unit is m.

In the present invention, pneumatic rudder face control allocation unit 207 receive from 201 place highly under velocity of sound information, Atmospheric temperature information under being located highly, the atmospheric pressure information under being located highly, the atmospheric density information under being located highly.? This, the pressure under inputting the temperature under being located highly, the velocity of sound and being located highly does not deal with；The 210 left elevon of aircraft Rudder believes one side only breath, and right elevon rudder believes one side only breath, rudder information, and body flap-type rudder believes one side only breath；301 aircraft altitude information, speed Degree information, angle of attack information, sideslip angle information, rolling angular rate information, pitch rate information, rolling angular rate information, accept Aerodynamic Coefficient from 203 and 206 pneumatic control torque command, provide the left elevon rudder face of aircraft after being resolved Deflection command, right elevon control surface deflection instruction, rudder kick instruction and the instruction of body wing flap control surface deflection, output information All to Unit 210.

Left and right elevon is used for the pitching of aircraft and rolling controls, and provides aileron to control work(when elevon is differential Can, elevator function is provided when elevon is moved upwardly or downwardly simultaneously.

The configuration of body wing flap, in the back lower place of fuselage, is on the one hand used for protecting the orbit maneuver of afterbody configuration to start Machine, is on the other hand used for trim and the manipulation of aircraft pitch channel.

Reentry vehicle adopts Large attack angle reentry to fly, and rudder is in the lee face of fuselage, and it effectively controls at end End energy management section.Therefore, during reentering, course passage needs RCS to participate in controlling.

（8）RCS controls allocation unit 208

In the present invention, 16 thrusters refer to the numeral numbering with circle in Fig. 4, that is, 1., 2., 3., 4., 5., ⑥、⑦、⑧、⑨、⑩、1WithThe thruster representing respectively.

As shown in Fig. 2A～Fig. 2 G, Fig. 4, in order to complete orbit maneuver, accurate pointing and the tasks such as docking that cross, reenter winged Row device is anterior and afterbody all can install RCS.For needing in aircraft return course to complete to reenter the attitude at initial stage using afterbody RCS Control.8 thrusters are each configured with installed surface S1, installed surface S2, arrow has marked spray when each thruster is opened Flow direction, the geometric center of installed surface S1 is, center of gravity isLine is vertical with installed surface S1, installed surface S2.Thruster is pacified Holding position is symmetrical, and in installed surface S1,1. number thruster and the horizontal range of 2. number thruster are designated as L₁, 1. number thruster 6. the vertical dimension of number thruster is designated as L₂, installed surface S1 is from center of gravityDistance be designated as L₃, installed surface S1 and installed surface S2's Axial distance is designated as d.Normal thrust T_Thrust=(F₀-p_{Air static pressure}A_{Cross-sectional area}).

The configuration of thruster determines the moment size and Orientation producing when RCS respective thrust device is opened, and control system will Comprehensively complete gesture stability using all or part thruster.Coordinate the work of multiple thrusters, need to be set according to RCS layout Meter controls distribution, can be controlled and multiple-rank control using split axle.In designed RCS allocation plan, installed surface S1 and installation Thruster on the S2 of face all can complete the gesture stability task of pitching, rolling and driftage, and the redundancy of allocation plan is 2.Table 1 is given Go out the control effect producing when each thruster of reentry vehicle afterbody is opened, can have been realized by the combination of thruster multiple Control moment, and provide hardware guarantee for system margin.

The control moment that table 1RCS thruster produces

In the present invention, RCS controls allocation unit 208 to instruct the control moment distributing to RCS from 207 and be derived from 210 RCS control moment compensation vector, using pulse width-pulse frequency（Pulse-Width Pulse-Frequency, PWPF) manipulator, the control moment instruction to realize RCS is transformed into the switch order of RCS thruster（1 opens, 0 closes).To be derived from The control moment instruction of 207 RCSWith from 210Sum is designated asAnd willIt is mapped to 16 The switch order of thruster, that is, complete the control distribution of RCS.Various control moment can be realized by the combination of thruster, and Hardware is provided to ensure for system margin.The control of RCS needs to coordinate the work of multiple thrusters, needs research to control distribution, with And a control reconfiguration scheme in the case of Actuator failure occurs.

RCS controls allocation unit 208 to receive the RCS control moment instruction from 207 and the RCS compensating torque from 210 Vector information, according to RCS layout designs, completes the thruster switch order life of RCS roll channel, pitch channel and jaw channel Become, provide the switch order of 16 thrusters.And export the switch order of each thruster to Unit 210, directly control 16 The working condition of individual thruster.

Various control moment can be realized by the combination of thruster, and provide hardware to ensure for system margin.The control of RCS System needs to coordinate the work of multiple thrusters, needs research to control distribution, and occurs in the case of an Actuator failure Control reconfiguration scheme.The combination of RCS thruster can realize various control moment, and provides hardware to ensure for system margin, control System distribution then needs to coordinate the work of multiple thrusters.

208-1, split axle control

In the present invention 16 thrusters are distributed to body axis system O-x_by_bz_bThree axles, so that them is not coupled, Each axle only one of which thrust rank and only a kind of combination.Although not making full use of Maximum controlling moment, can be simultaneously Control three axles, it is to avoid coupling is produced to select logistical difficulties.The moment size of so each axle is fixing.Now 9., ⑩、WithNumber thruster is controlled to wobble shaft, using a 1., 2., 5. and 6. number thruster, pitch axis is controlled, ③、④、WithNumber thruster is controlled to yaw axis, and moment size and Orientation is true as requested Fixed respective opening and time.Concrete thruster is applied in combination method as shown in Figure 4 A.

In Figure 4 A, reentry vehicle is along x_bAxle rolling, and the aircraft left side occurs motion upwards, the right to occur to transport downwards Move as positive rolling；Reentry vehicle is along x_bAxle rolling, and the aircraft left side moves downward, the right occurs that to move upwards be negative Rolling.Reentry vehicle is along y_bAxle pitching, and Vehicle nose is just to bow upwards；Reentry vehicle is along y_bAxle pitching, and aircraft Head is downwards negative bowing.Reentry vehicle is along z_bAxle swings, and Vehicle nose navigates to the left side for negative bias；Reentry vehicle is along z_b Axle swings, and Vehicle nose is just to go off course to the right.

The advantage of this method is that logical course is simple, and the control of three axles is separate, does not affect each other, control The effect that effect processed can be given close to control law well；Shortcoming is each axle only one of which class, the power that can utilize Square is too little, and when there is thruster inefficacy, RCS will lose the control ability of thruster respective shaft, and system margin cannot be real Existing.

208-2, multiple-rank control

Multiple-rank control improves on the basis of split axle control, to improve the available moment class of each axle.Three axle controls The class rank of system is 5 grades, and corresponding class word is -2 grades（Maximum negative moment）, -1 grade（Negative moment）, 0 grade（Do not work）, 1 grade （Positive moment）With 2 grades（Maximum positive moment）.Concrete thruster is applied in combination method as shown in Figure 4 B.

Due in split axle control, three axles all distribute fixing thruster the need of controlling to it, so when Certain axle does not need moment and other axles need during larger torque it is impossible to rational distribute.Therefore make improvements, that is, first Judge the moment size required for each axle.When the moment of three axles is above particular value it is believed that three axles are required for carrying out Control, now then respectively three axles are controlled according to " split axle control "；When some or two axles do not need control moment When, then other axle can be provided with two grades of controls.When any thruster lost efficacy, each axle control moment is part thrust level Do not enable, RCS still keeps the completeness that three axles control.Can see from Fig. 4 B, in addition to side thruster, remaining Thruster can use cooperatively, and so can make full use of thruster and obtain larger control moment.

RCS is side-jet control, in order that the effect that control effect is given close to control law as far as possible, special to verify flight Property, here, the control distribution of RCS adopts split axle control mode.Design the RCS split axle control program under this layout, and be based on PWPF manipulator is realized.Allocation unit is controlled to be composed in series by adder 1 and PWPF manipulator.Adder is realized being derived from The control moment instruction of 207 RCSWith from 210It is added, be designated asManipulator willIt is mapped to In the switch order (0-1) of 16 thrusters.

As shown in Figure 4 A, 16 thrusters are allocated fixedly to rolling, pitching and jaw channel.So, thruster Call and there is not coupling, each axle only one of which thrust rank and only one kind combine.Although the moment size of each axle is solid Fixed, do not make full use of Maximum controlling moment, but 3 axles can be controlled simultaneously, it is to avoid coupling is produced to select logic to be stranded Difficult.Now 9., 10.,WithNumber thruster is controlled to wobble shaft, using 1., 2., 5. and 6. number thruster to pitch axis Be controlled, 3., 4.,⑦、⑧、WithNumber thruster is controlled to jaw channel, according to RCS control moment Size and Orientation determined the opening of each thruster by manipulator.

Control moment in order to realize RCS instructs the switch order being transformed into RCS thruster, can be based on schmidt trigger Device, using pulse width-pulse frequency modulated（PWPF）To realize.Analog circuit is capable of the modulation of PWPF, and here will PWPF manipulator to be realized as an independent link, is not linear relationship between the input of PWPFM control loop and output, And it is in switching mode, the thruster open command for pitching, driftage and roll channel generates.

（9）Aerodynamic data library unit 209

Aerodynamic data library unit 209 is used for providing aerodynamic force suffered by simulated flight device, aerodynamic moment desired parameters.Energy Enough play aerodynamic coefficient, the effect of aerodynamic moment coefficient that calculating aircraft in-flight measures in real time, pneumatic by input Control surface deflection angle, enters row interpolation according to Mach number, the angle of attack and yaw angle in aerodynamic data table, obtains aerodynamic force, aerodynamic force Moment coefficient.In aerocraft real flight course suffered pneumatic be to there is interference and uncertain, can be in this unit Introduce.

Aerodynamic data library unit 209 receives aircraft altitude information, the velocity information being derived from 301, angle of attack information, yaw angle Information, rolling angular rate information, pitch rate information, yawrate information, 201 place highly under velocity of sound information, Atmospheric temperature information under being located highly, the atmospheric pressure information under being located highly, the atmospheric density information under being located highly. 210 aircraft left elevon rudder believes one side only breath, and right elevon rudder believes one side only breath, rudder information, and body flap-type rudder believes one side only breath.Enter Row provides lift coefficient, resistance coefficient, lateral force coefficient and the pneumatic rolling moment coefficient suffered by aircraft, pneumatic pitching after resolving Moment coefficient, pneumatic yawing moment coefficient information, output information is all to Unit 210.

Reentry vehicle will realize reusable it is necessary to satisfaction reenters dynamic pressure, overload and rate of heat flow of return course etc. Limit, this requires that aircraft has higher lift-drag ratio, by lift control so that aircraft being capable of level after reentering air Land.Meanwhile, to complete to reenter the control of attitude and track, management system need to be configured.

Reenter as typical lift, reentry vehicle will meet low speed, transonic speed Low Angle Of Attack and the hypersonic big angle of attack Aerodynamic force, pneumatic heat request, meet in orbit and the required thermal protection requirement of endoatmosphere flight, and unpowered reenter Flight and the aeroperformance landing, stability and operability require.With reference to the aerodynamic arrangement of space shuttle and X-37B, design Reentry vehicle aerodynamic configuration, as shown in Fig. 2A～Fig. 2 G, employs multi-wall interference layout, and wing is single-blade under S-shaped, in machine The trailing edge of the wing is configured with left and right elevon；The fuselage back lower place is configured with body wing flap, and afterbody is configured with vertical fin, aircraft It is that left and right face is symmetrical.

The aerodynamic characteristic of hypersonic aircraft is closely related with aerodynamic configuration, flight path and flying quality, is also true Determine hypersonic aircraft load-up condition（Pressure distribution）And thermal environment（Skin friction stress is distributed）Important evidence.Due to height It is multifactor that the profile of supersonic aircraft depends on task mission and pneumatic, propulsion system, structure and material etc..

Definition S is aircraft area of reference, and b is wing length, and c is wing mean aerodynamic chord,For dynamic pressure, act on Carry-on aerodynamic liftResistanceSide forcePneumatic rolling momentPitching momentAnd yawingWherein, C_LIt is lift coefficient, C_D It is resistance coefficient, C_YIt is lateral force coefficient, C_lIt is rolling moment coefficient, C_mIt is pitching moment coefficient, C_nIt is yawing moment coefficient.

Reentry vehicle is the product that aerospace craft is organically blended with aviation aircraft technology, and existing a lot of technology can With direct with reference to continue to use, especially hypersonic aerodynamics, as one of the heavy Technology Difficulties of reentry vehicle, is ought One focus of modern aerodynamic studies and forward position.Three kinds of Main Means of hypersonic aerodynamics research are that theory is ground Study carefully, ground experiment is studied and Flight Test Research, these three means are complementary.Wherein, include work in theoretical research again Journey calculating, numerical computations, Fluid Mechanics Computation（CFD）.The engineering calculating method design cycle is short, and computational efficiency is high, and cost is relatively low； In reentry vehicle conceptual approach and concept phase, to predict that the pneumatic of aircraft is critically important using engineering calculating method Design tool.CFD approach then needs to carry out substantial amounts of calculating, spends the more calculating time, but but being capable of simulated flight The flow field of device, obtains more aerodynamic force and aerothermal information, with the development of aerodynamic and computer technology, CFD Obtain more and more and widely apply, become the important means of pneumatic design and research.

Obtain the aerodynamic data table of aircraft, C by theoretical research, ground experiment research and Flight Test Research_L, C_D, C_Y, C_l, C_m, C_n, the dynamic pressure according to residing for current aircraft and aircraft parameters, obtain the aerodynamic force suffered by aerocraft real, Aerodynamic moment coefficient, this module is pneumatic for the measurement in real time of simulated flight device, and aerocraft real is in-flight and air The physical process of relative motion, measurement is pneumatic relevant with flight specific environment in real time, and also various the air-dried of the external world disturb.

（10）Control integrated unit 210

Integrated unit 210 is controlled to be used for coordinating the complex controll of RCS and pneumatic rudder.Distribution 207 is controlled according to pneumatic rudder face The control surface deflection instruction that subelement is given and RCS control distribute the RCS thruster switch order that subelement 208 provides, and are controlling Merge the control coordination realizing RCS and pneumatic rudder in this unit 210, and consider the dynamic of RCS thruster and Rudder loop Dynamic, RCS control moment compensation dosage is passed to RCS control by the pneumatic estimation (17) in conjunction with 203,301 partial information Distribution 208, the control surface deflection of real-time measurement is fed back to 203,207 and 209.Control in merging, need to receive the letter from 201 The aerodynamic data of breath, 301 partial information and 209 subelements, the aerodynamic force suffered by estimation aircraft, aerodynamic moment, comprehensive Obtain aerodynamic force, the axial force transmission of total moment and RCS thruster gives Unit 301.

Referring to shown in Fig. 3 A, the deflection of pneumatic rudder face to be completed by Rudder loop, and it dynamically to be simulated by order transfer function. For left elevon Rudder loop it is dynamicallyWherein, s is aircraft area of reference, damping Ratio ζ=0.7, natural angular frequency ω_n=90rad/s；In rate versus position limiter 1, control surface deflection scope is limited to ± 30 °.Right In right elevon Rudder loop it is dynamicallyWherein, dampingratioζ=0.7, natural angular frequency ω_n=90rad/s；In rate versus position limiter 2, control surface deflection scope is limited to ± 30 °.

For rudder Rudder loop it is dynamicallyWherein, dampingratioζ=0.7, natural Angular frequency_n=70rad/s；In rate versus position limiter 3, control surface deflection scope is limited to ± 20 °.

For body flap-type rudder loop dynamic it isWherein, dampingratioζ=0.7, from So angular frequency_n=90rad/s；In rate versus position limiter 4, control surface deflection scope is limited to [- 20 ° ,+30 °].

In the present invention, the deflection command of 4 pneumatic rudders through respective Rudder loop is dynamic and rate versus position limiter after Obtain the real-time measurement deflection angle of rudder face, on the one hand pass to RCS control moment and compensate subelement, on the one hand rudder is wilfully turned Angle feed-back gives 203,207 and 209.Control in merging, here is used for designing the control coordination of the pneumatic rudder of RCS/ it is ensured that control law Effective realization.

β, p, q and the r control integrated unit 210 Aerodynamic Coefficient inputting according to 203, combining 301 inputs, velocity information V, Aircraft area of reference s, calculates dynamic pressureAnd define pneumatic control moment vector function

Then have RCS compensating torque vectorPneumatic control surface deflection instruction vectorPneumatic control surface deflection VectorRCS compensating torque vector

When pneumatic Rudder loop signal is within rate limit and position limitation, then error signal is that zero, RCS does not work. Only when pneumatic Rudder loop exceeds rate limit and position limitation scope, ε is not zero, now RCS work, in order to compensate pneumatic rudder face The deficiency of operating torque.Merged using controlling, be to coordinate RCS/ pneumatic rudder complex controll, merged by controlling, can subtract Little fuel consumption, it is to avoid RCS works long hours, extends its service life.

Reenter the starting stage, reentry vehicle only to complete gesture stability by RCS.And work as RCS and coordinate to control with pneumatic rudder When, with pneumatic rudder for main actuator, the full time service of main actuator, by the arrangement of chain type order, gradually it is incremented by；RCS For assisting actuator, auxiliary actuator just works when needed.So, could be to greatest extent using Conventional pneumatic rudder Face, reduces the working time of RCS, and then reduces the RCS fuel carrying needed for reentry vehicle.

RCS is made up of one group of thruster and igniter, fuel tank and supply pipeline etc..Thruster is pushed away using constant value Power, can only control being turned on and off of thruster.During RCS thruster is dynamic, normal thrust T_Thrust=(F₀-p_{Air static pressure}A_{Cross-sectional area}), its In, F₀For the normal thrust of thruster vacuum, p_{Air static pressure}For ambient atmosphere static pressure, p_{Air static pressure}=p (H) is the function of height, A_{Cross-sectional area} For thruster exit cross-sectional area.RCS thruster open and close dynamic process beWherein, the time is normal Number

0 pass of 16 shown in Fig. 4 thruster instructs, 1 opens instruction through RCS thruster dynamically, has obtained total RCS axle To thrust（Body is axiallyBody lateral RCS thrustBody normal direction RCS thrustThe moment that RCS thruster produces （l_R, m_R, n_R）.Aerodynamic force, aerodynamic moment computation subunit are according to the dynamic pressure information calculating in this unitWith flying of 209 inputs The lift coefficient C of row device measurement in real time_L, the resistance coefficient C of aircraft measurement in real time_D, the lateral force coefficient of aircraft measurement in real time C_Y, the pneumatic rolling moment coefficient C of aircraft measurement in real time_l, the pneumatic pitching moment coefficient C of aircraft measurement in real time_m, flight The pneumatic yawing moment coefficient C of device measurement in real time_n, and the parameter of aircraft, come aerodynamic force suffered by calculating aircraft, gas Kinetic moment acts on.Obtain the aerodynamic force suffered by aerocraft real, aerodynamic moment：

By aerodynamic moment l_A, m_A, n_AMoment l producing with RCS thruster_R, m_R, n_RRespectively through addition process, flown Total moment suffered by row device is：Total rolling moment l=l_R+l_A；Total pitching moment m=m_R+m_A；Total yawing n =n_R+n_A；Finally, by aerodynamic force L, D, the Y suffered by the aerocraft real obtaining and total moment l, m, n, total RCS axially pushes away Power（Body is axiallyBody lateral RCS thrustBody normal direction RCS thrustExport to Unit 301.

（Three）Aircraft characterization model 300

Aircraft characterization model 300 includes dummy vehicle unit 301 and emulation terminates judging unit 302；

In the present invention, six degree of freedom is designated as the first degree of freedom, the second degree of freedom, Three Degree Of Freedom, the 4th freedom respectively Degree, five degree of freedom and six degree of freedom；First degree of freedom：Refer to along Ox_gThe translation in direction.Second degree of freedom：Refer to along Oy_g The translation in direction.Three Degree Of Freedom：Refer to along Oz_gThe translation in direction.Four-degree-of-freedom：Refer to around Ox_bThe rotation of axle.5th certainly By spending：Refer to around Oy_bThe rotation of axle.Six degree of freedom：Refer to around Oz_bThe rotation of axle.

In the present invention, ten two-states are designated as first state, the second state, the third state, the 4th state, the 5th shape respectively State, the 6th state, the 7th state, the 8th state, the 9th state, the tenth state, the 11st state, the tenth two-state；

First state：Refer to the longitude information of reentry vehicle present position；Second state：Refer to residing for reentry vehicle The latitude information of position；The third state：Refer to the elevation information of reentry vehicle present position；4th state：Refer to reenter winged The airspeed information of row device；5th state：Refer to the trace drift angle information of reentry vehicle；6th state：Refer to reenter winged The trace obliquity information of row device；7th state：Refer to the angle of attack attitude information of reentry vehicle；8th state：Refer to reenter winged The yaw angle attitude information of row device；9th state：Refer to the angle of heel attitude information of reentry vehicle；Tenth state：Refer to again Enter the rolling angular rate information of aircraft；11st state：Refer to the pitch rate information of reentry vehicle；12nd shape State：Refer to the yawrate information of reentry vehicle.

Dummy vehicle unit 301

In the present invention, dummy vehicle unit 301 obtains the state of subsequent time according to current state, completes to fly The renewal of device six degree of freedom 12 state.Vehicle dynamics/kinematics model 301 unit accepts single from control fusion 210 The resistance suffered by the aircraft of output information of unit, the lift suffered by aircraft, the side force suffered by aircraft, aircraft Suffered total rolling moment, the total pitching moment suffered by aircraft, the total yawing suffered by aircraft, body axial direction RCS pushes away Power, body lateral RCS thrust, body normal direction RCS thrust.The information of 301 unit outputs is the aircraft six degree of freedom after updating Ten two-states, i.e. the longitude information of reentry vehicle present position；The latitude information of present position；The height letter of present position Breath；The airspeed information of reentry vehicle；Flight path drift angle information；Flight path angle information；Angle of attack attitude information；Yaw angle attitude Information；Angle of heel attitude information；Rolling angular rate information；Pitch rate information；Yawrate information.

Described dummy vehicle unit 301 first aspect is used for for the elevation information H after updating feeding back to 201；Second party Face is for giving 202 and 206 by the velocity information V feedback after renewal；The third aspect is used for 12 status informations after updating Give 203,204,205,209 respectively；Fourth aspect exports elevation information H and velocity information V to emulation judging unit 302.

In the present invention, dummy vehicle unit 301 realizes state more using 4 rank Runge-Kutta algorithms of standard Newly.Make current flight state vector y_N=[τ λ H V χ γ α β μ p q r]^T, t_NFor current time.Control input vector

By vehicle dynamics/kinematical equation, according to 4 rank Rung-Kutta methods, by current time t_NFlight shape State y_NWith control input vector u, obtain subsequent time t_N+1State of flight y_N+1.Here, vehicle dynamics/kinesiology six is certainly By spending ten two-state equations it is：

The derivative of longitudeR be the earth's core away from,

The derivative of latitude

The derivative of height

The derivative of speed

The derivative of flight path angle

The derivative of flight path drift angle

The derivative of the angle of attack

The derivative of yaw angle

The derivative of angle of heel

The derivative of rolling angular speed

The derivative of pitch rate

The derivative of yawrate

4 rank Runge-Kutta algorithms are realized as follows：

The derivative of state of flight vector

Time interval h=t_N+1-t_N；

Status variation rate first estimator k of current time₁=f (t_N,y_N,u)；

Status variation rate second estimator of current time

Status variation rate the 3rd estimator of current time

Status variation rate the 4th estimator of current time

Judge emulation whether end unit 302

Emulation judging unit 302 first aspect is to the elevation information H receiving and given thresholdCarry out size judgement；The Two aspects are to velocity information V receiving and given thresholdCarry out size judgement；IfOrThen terminate Emulation；IfThen proceed to emulate, until meetingOrAfter terminate emulate.Emulation Result carries out real time demonstration.

Need to apply the closed-loop simulation system of the present invention to proceed emulation, execution global variable N cumulative 1（Program is compiled Code form is N=N+1）, then it is transferred to attitude command signal generating unit 202 and run again according to the closed loop flow process of control system, Obtain new simulation time and 12 status informations of aircraft, until meeting the condition that emulation terminates.At the end of emulation, system Closed loop completes ablated configuration emulation, the control effect of attitude control law and the method for designing of modules of reentry vehicle Embody in input-output information all after emulation terminates, and realize respective function, whole closed-loop simulation result is able to verify that The function of modules, and modules of can upgrading, or write new method in modules.Attitude control system Circled simulation flat has versatility, expansible new and transplantability.

The simulation flow that carries out of closed-loop simulation system applying reentry vehicle gesture stability of the present invention is：

（A）Analogue system initializes；Global variable is N=0, arranges simulation step length h, the initial time t of reentry vehicle₀ =0, initial attitude state y₀=[τ₀λ₀H₀V₀χ₀γ₀α₀β₀μ₀p₀q₀r₀]^T；τ₀For initial orientation angle, λ₀For first The beginning angle of site, H₀Elemental height, V₀Initial velocity, χ₀Initial flight path azimuthangle, γ₀Initial trajectory inclination angle, α₀The initial angle of attack, β₀Just Beginning yaw angle, μ₀Initial angle of heel, p₀Initial roll angle, q₀Initial pitch angle, r₀Initial yaw angle；Initial value is classified as zero, and flies All control surface deflection angles of row device are zero（δ_e,δ_a,δ_g,δ_f）, RCS（16 thruster status commands are 0）Do not work；Initially Change 12 states of dummy vehicle unit 301, initialization RCS controls the RCS control of allocation unit 208 to be assigned as zero；Just Beginning activating QI moves rudder face and controls the pneumatic rudder control of allocation unit 207 to be assigned as zero；Initialization controls the pneumatic rudder of integrated unit 210 Deflecting facet is zero；12 status informations of initialization of dummy vehicle unit 301 are inputted Atmospheric models unit 201, attitude Command unit 202, pneumatic estimation unit 203, slow circuit controls restrain unit 204, fast loop control law unit 205, complex controll Allocation unit 206, pneumatic rudder face controls allocation unit 207, aerodynamic data library unit 209；

（J）RCS controls allocation unit 208 according to complex controll allocation unit 206 information of input and to control integrated unit 210 information calculate output to control integrated unit 210；

（K）Control integrated unit 210 according to Atmospheric models unit 201 information inputting, pneumatic estimation unit 203 aerodynamic force Moments estimation information, dummy vehicle unit 301 information and pneumatic rudder face control allocation unit 207, RCS to control allocation unit 208 Information, aerodynamic data library unit 209 information calculates output information to dummy vehicle unit 301；And the rudder face by real-time measurement Deflection information exports and controls allocation unit 207 and pneumatic Database Unit 209 to pneumatic estimation unit 203, pneumatic rudder face, will RCS control and compensation moment information is passed to RCS and is controlled allocation unit 208；

（L）Dummy vehicle unit 301 is according to the control integrated unit 210 information updating state of flight of input, and exports , to Atmospheric models unit 201, output speed status information is to attitude command unit 202 and complex controll allocation unit for elevation information 206, export attitude angle, attitude angular rate information to slow circuit controls rule unit 204 and fast loop control law unit 205, output Highly, speed, the angle of attack, yaw angle, attitude angular rate information give pneumatic estimation unit 203, and pneumatic rudder face controls allocation unit 207 With pneumatic Database Unit 209；Export height and velocity information to emulation whether end unit 302；

For the aerial mission returning from track and attitude control system, the present invention proposes a kind of reentry vehicle attitude The closed-loop simulation system controlling, flight function and assembly is carried out modularized design, establishes general module and system emulation Framework.Closed-loop simulation system can be used as an emulation platform, and the detailed design standard of modules has autgmentability and rule Plasticity.The closed-loop simulation system modular structure of the present invention, including initialization program, the submodule of multiple function opposite independent, And data storage file.Here is it is proposed that the control system closed-loop simulation framework of entirety, each module and signal stream have been carried out in detail Thin design, and the function to module and flow process give specific description.Designed control system closed-loop simulation, for the greatest extent The early defect finding design of control law, the safety of control strategy, the control of aircraft under checking uncertain factor and external interference Ability processed, the control characteristic of analysis aircraft, and provide with reference to significant for master-plan, research and development can not only be shortened Cycle, and greatly save development cost and flight test cost.Can be applied to execute ablated configuration task sky and space plane, Space station occupant's recoverable capsule, the Attitude Control System Design of Control System for Reusable Launch Vehicle.

The closed-loop simulation system of the present invention, has implemented the general modeling of modules, is the one of attitude control system Plant implementation, can effectively reduce design cycle and the workload of simulated program layout, be also convenient for simulation management software in addition By the different module of different rates scheduling, reduce calculated load, meet the requirement of real-time simulation.In analogue system modules it Between carry out the transmission of signal, each module realizes corresponding function and algorithm, and overall emulation platform is to attitude control system Simulation and the simulation modeling of algorithm.

Access control strategy and the design of control law, the control ability of comprehensive analysis aircraft and characteristic, design reenters winged The aerial mission of row device, this is the final purpose of control system closed-loop simulation.Closed-loop simulation system designed by patent of the present invention SIMULINK environment is built, can quickly and easily be converted into c program or directly call in MATLAB environment, be used for testing The design of card attitude controller, control strategy and control allocation algorithm.Verify such new flying vehicles control performance and ability, control The reasonability of scheme processed, the scope of application of control and analysis and the improvement direction needing, are controlled assessment and the test restrained, Only corresponding algorithm corresponding functional module need to be loaded into, the aircraft related to attitude control system can be rapidly verified Can index and characteristic.

Claims

1. a kind of closed-loop simulation system being applied to reentry vehicle gesture stability, this system includes the gas to reentry vehicle Subsystem (100), the structure subsystem (200) of Attitude control model and the son of aircraft characterization model that dynamic profile is modeled System (300)；It is characterized in that：The described subsystem (200) building Attitude control model includes atmospheric environment model unit (201), attitude command signal generating unit (202), flight vehicle aerodynamic evaluation unit (203), slow circuit controls rule unit (204), fast Circuit controls rule unit (205), complex controll allocation unit (206), pneumatic rudder face control allocation unit (207), RCS control point Join unit (208), flight vehicle aerodynamic data base (209) and control integrated unit (210)；

(A) analogue system initialization；Global variable is N=0, arranges simulation step length h, reenters initial time t₀=0, initial attitude State y₀=[τ₀λ₀H₀V₀χ₀γ₀α₀β₀μ₀p₀q₀r₀]^T；And all control surface deflection angles of aircraft are zero, that is, δ_e,δ_a,δ_g,δ_fIt is set to zero, RCS not work, that is, 16 thruster status commands are 0；Initialization dummy vehicle unit (301) 12 states, initialization RCS controls the RCS control of allocation unit (208) to be assigned as zero；Initialize pneumatic rudder face The pneumatic rudder control controlling allocation unit (207) is assigned as zero；Initialization control integrated unit (210) pneumatic control surface deflection be Zero；By 12 status informations input atmospheric environment model unit (201) of initialization of dummy vehicle unit (301), attitude Instruction generation unit (202), pneumatic estimation unit (203), slow circuit controls rule unit (204), fast loop control law unit (205), complex controll allocation unit (206), pneumatic rudder face controls allocation unit (207), flight vehicle aerodynamic data base (209)；

τ₀For initial orientation angle；

λ₀For the initial angle of site；

H₀For elemental height；

V₀For initial velocity；

χ₀For initial flight path azimuthangle；

γ₀For initial trajectory inclination angle；

α₀For the initial angle of attack；

β₀For initial side-slip angle；

μ₀For initial angle of heel；

p₀For initial roll angle；

q₀For initial pitch angle；

r₀For initial yaw angle；

δ_eFor left elevon rudder partially, unit is degree；

δ_aFor right elevon rudder partially, unit is degree；

δ_gFor rudder rudder partially, unit is degree；

δ_fFor body flap-type rudder partially, unit is degree；

(B) according in elevation information input atmospheric environment model unit (201) of dummy vehicle unit (301)；Air Environmental model unit (201) by output information input to attitude command signal generating unit (202), pneumatic estimation unit (203), multiple Close and control allocation unit (206), pneumatic rudder face controls allocation unit (207), flight vehicle aerodynamic data base (209), controls and merges Unit (210)；

(C) flight vehicle aerodynamic data base (209) believes according to the atmospheric environment of the atmospheric environment model unit (201) of current input Breath, dummy vehicle unit (301) attitude information and the control inclined deflection information of integrated unit (210) rudder calculate output aerodynamic data Information is given and is controlled integrated unit (210)；

(D) pneumatic estimation unit (203) according to the atmospheric environment information of the atmospheric environment model unit (201) of current input, Dummy vehicle unit (301) attitude information and control integrated unit (210) control surface deflection information will calculate output aerodynamic force letter Cease and control allocation unit (207) to slow circuit controls rule unit (204) and aerodynamic moment coefficient estimation to pneumatic rudder face and control Integrated unit (210)；

(E) attitude command signal generating unit (202) is according to atmospheric environment model unit (201) information inputting and dummy vehicle list First (301) information calculates output to slow circuit controls rule unit (204)；

(F) slow circuit controls restrain unit (204) according to attitude command signal generating unit (202) information inputting, pneumatic estimation unit (203) information and dummy vehicle unit (301) information calculate output to fast loop control law unit (205)；

(G) unit (204) information and dummy vehicle list are restrained according to the slow circuit controls of input in fast loop control law unit (205) First (301) information calculates output to complex controll allocation unit (206)；

(H) complex controll allocation unit (206) is according to atmospheric environment model unit (201) information inputting, fast loop control law Unit (205) information and dummy vehicle unit (301) information calculate output to pneumatic rudder face control allocation unit (207) and RCS controls allocation unit (208)；

(I) pneumatic rudder face controls allocation unit (207) according to atmospheric environment model unit (201) information inputting, aircraft mould Type unit (301) partial information and integrated unit (210) the control surface deflection information that controls, synthesis are derived from complex controll allocation unit (206) aerodynamic moment command information and the aerodynamic moment coefficient estimation information of pneumatic estimation unit (203) calculate output to control Integrated unit (210) processed；

(J) RCS controls allocation unit (208) according to complex controll allocation unit (206) information of input and to control integrated unit (210) partial information calculates output to control integrated unit (210)；

(K) control integrated unit (210) according to atmospheric environment model unit (201) information inputting, pneumatic estimation unit (203) Aerodynamic moment estimated information, dummy vehicle unit (301) partial information and pneumatic rudder face control allocation unit (207), RCS control Allocation unit (208) information processed, flight vehicle aerodynamic data base (209) information calculates output information to dummy vehicle unit (301)；And real control surface deflection information output controls allocation unit (207) to pneumatic estimation unit (203), pneumatic rudder face With flight vehicle aerodynamic data base (209), RCS control and compensation moment information is passed to RCS and controls allocation unit (208)；

(L) dummy vehicle unit (301), according to control integrated unit (210) the information updating state of flight of input, and exports Elevation information and is combined to attitude command signal generating unit (202) to atmospheric environment model unit (201), output speed status information Control allocation unit (206), export attitude angle information to slow circuit controls rule unit (204) and fast loop control law unit (205), output par, c information gives pneumatic estimation unit (203), and pneumatic rudder face controls allocation unit (207) and flight vehicle aerodynamic number According to storehouse (209)；Export height and velocity information to emulation whether end unit (302)；

(M) according to input, from emulation, whether the height of end unit (302) and velocity information judge to emulate and terminate；If Emulation terminates, and shows simulation result, and ends with system emulation；If emulation does not terminate, proceed in step (B) and repeat to imitate Very, until it reaches emulation termination condition.

2. the closed-loop simulation system being applied to reentry vehicle gesture stability according to claim 1 it is characterised in that：Institute State complex controll allocation unit (206) receive from the aircraft altitude information of dummy vehicle unit (301), velocity information and Angle of attack information；The rolling moment command information being inputted according to fast loop control law unit (205), pitching moment command information and partially Boat torque command information, the comprehensive atmospheric environment information being derived from atmospheric environment model unit (201), design RCS and pneumatic rudder face Complex controll distribution resolve, obtain distributing to pneumatic rudder face control allocation unit (207) pneumatic rudder control moment information and RCS controls the RCS control moment command information of allocation unit (208)；

According to dynamic pressure border, press pressure linear transitions, realize the control allocation strategy of RCS and pneumatic manoeuvring systemWherein,Control moment for distributing to RCS instructs, and k is weight coefficient, Control moment for distributing to pneumatic rudder system instructs, andM_cFor controling power Square instructs,For the rolling moment instruction of pneumatic rudder, unit is Nm,For the pitching moment instruction of pneumatic rudder, unit is Nm,For the yawing instruction of pneumatic rudder, unit is Nm,Start to intervene dynamic pressure during gesture stability for pneumatic rudder face,For RCS exits dynamic pressure during gesture stability, and the control authority of RCS and pneumatic rudder face is according to dynamic pressureBe allocated with Coordinate, q be pitch rate, unit be rad/s, ρ be aircraft be located highly under density, unit be kg/m³, V is flight The speed of device.

3. the closed-loop simulation system being applied to reentry vehicle gesture stability according to claim 1 it is characterised in that：Gas Dynamic rudder face controls allocation unit (207) to instruct the control moment of pneumatic rudder systemIt is mapped to angle of rudder reflection instruction δ_c, that is, complete The control distribution of pneumatic manoeuvring system；Wherein, For base This control moment vector, M_A,δFor controlling allocation matrix；

For the rolling moment instruction of pneumatic rudder, unit is Nm；

For the pitching moment instruction of pneumatic rudder, unit is Nm；

For the yawing instruction of pneumatic rudder, unit is Nm；

For the instruction of left elevon, unit is degree；

For the instruction of right elevon, unit is degree；

For rudder instruction, unit is degree；

Partially instruct for body flap-type rudder, unit is degree；

M_{A}^{b} = [\begin{matrix} \overset{&OverBar;}{q} {SbC}_{l, β} β + \overset{&OverBar;}{q} {Sb}^{2} C_{l, p} \frac{p}{2 V} + \overset{&OverBar;}{q} {Sb}^{2} C_{l, r} \frac{r}{2 V} \\ \overset{&OverBar;}{q} {ScC}_{m, α} + \overset{&OverBar;}{q} {Sc}^{2} C_{m, q} \frac{q}{2 V} \\ \overset{&OverBar;}{q} {SbC}_{n, β} β + \overset{&OverBar;}{q} {Sb}^{2} C_{n, p} \frac{p}{2 V} + \overset{&OverBar;}{q} {Sb}^{2} C_{n, r} \frac{r}{2 V} \end{matrix}];

M_{A, δ} = [\begin{matrix} \overset{&OverBar;}{q} {SbC}_{l, δ_{e}} & \overset{&OverBar;}{q} {SbC}_{l, δ_{a}} & \overset{&OverBar;}{q} {SbC}_{l, δ_{r}} & 0 \\ \overset{&OverBar;}{q} {ScC}_{m, δ_{e}} & \overset{&OverBar;}{q} {ScC}_{m, δ_{a}} & 0 & \overset{&OverBar;}{q} {ScC}_{m, δ_{f}} \\ \overset{&OverBar;}{q} {SbC}_{n, δ_{e}} & \overset{&OverBar;}{q} {SbC}_{n, δ_{a}} & \overset{&OverBar;}{q} {SbC}_{n, δ_{r}} & 0 \end{matrix}];

Wherein, dynamic pressureQ be pitch rate, unit be rad/s, ρ be aircraft place highly under density, Unit is kg/m³, V is the speed of aircraft；

S is the area of reference of aircraft wing, and unit is m²；

B is that the span of the wing of aircraft is long, and unit is m；

C is the mean aerodynamic chord of the wing of aircraft, and unit is m；

The angle of attack C rising under force environment taking the logarithm in Knudsen number Kn_L,α, left rudder inclinedRight standard rudder is inclinedRudder is inclinedBody flap rudder is inclined

Angle of attack C under the resistance environment that Knudsen number Kn takes the logarithm_D,α, left rudder inclinedRight standard rudder is inclinedRudder is inclinedBody flap rudder is inclinedAngle of attack C under the Mach number environment that Knudsen number Kn takes the logarithm_m,α, left rudder inclinedRight standard rudder is inclinedRudder is inclinedBody flap rudder is inclinedPitch rate C_m,q；

Yaw angle C under the side force environment that Knudsen number Kn takes the logarithm_Y,β, left rudder inclinedRight standard rudder is inclinedRudderThe yaw angle C rising under force environment that matching is estimated_l,β, left rudder inclinedRight standard rudder is inclinedRudder is inclinedRoll angle Speed C_l,p, yaw rate C_l,r；

Yaw angle C under the Mach number environment that matching is estimated_n,β, left rudder inclinedRight standard rudder is inclinedRudder is inclinedRoll angle Speed C_n,p, yaw rate C_n,r；

α is the angle of attack, and unit is rad；

β is yaw angle, and unit is rad；

P is rolling angular speed, and unit is rad/s；

Q is pitch rate, and unit is rad/s；

R is yawrate, and unit is rad/s.

4. the closed-loop simulation system being applied to reentry vehicle gesture stability according to claim 1 it is characterised in that：Institute State dummy vehicle unit (301) according to 4 rank Rung-Kutta methods, by current time t_NState of flight y_NWith controlled quentity controlled variable u, Obtain subsequent time t_N+1State of flight y_N+1；Here, vehicle dynamics with kinesiology six degree of freedom ten two-state equation is：

The derivative of longitudeV be aircraft speed, R be the earth's core away from；

The derivative of latitude

The derivative of height

The derivative of speed

M is aircraft Quality, D is the resistance of aircraft measurement in real time, and g is acceleration of gravity, ω_nFor natural angular frequency,Push away for body axial direction RCS Power,For body lateral RCS thrust,For body normal direction RCS thrust；

The derivative of flight path angle

L is to act on flight Aerodynamic lift on device, Y is to act on carry-on side force；

The derivative of flight path drift angle

\begin{matrix} \overset{\cdot}{χ} = \frac{1}{M V \cos γ} (Y \cos μ + L \sin μ) + \frac{V}{R \cos γ} {tanλcos}^{2} γ \sin χ + \\ + \frac{2 ω_{n}}{\cos γ} (- \cos λ \sin γ \cos χ + \sin λ \cos γ) + \frac{{Rω}_{n}^{2}}{V \cos γ} \sin λ \cos λ \sin χ \\ + \frac{1}{M V \cos γ} (T_{x}^{R} (\sin α \sin μ - \cos α \sin β \cos μ) \\ + T_{y}^{R} \cos β \cos μ - T_{z}^{R} (\cos α \sin μ + \sin α \sin β \cos μ)) \end{matrix};

The derivative of the angle of attackP is rolling angular speed_,Q is angle of pitch speed Rate, r is yawrate；

The derivative of yaw angle

The derivative of angle of heel

\begin{matrix} \overset{\cdot}{μ} = \frac{1}{M V} (L \tan β + L \tan γ \sin μ + Y \tan γ \cos μ \\ - M g \tan β \cos γ \cos μ) + p \cos α \sec β + r \sin α \sec β \\ + \frac{1}{M V} (T_{x}^{R} (\sin α \sin μ \tan γ - \cos α \sin β \cos μ \tan γ \\ + \sin α \tan β) + T_{y}^{R} \cos β \cos μ \tan γ \\ - T_{z}^{R} (\cos α \sin μ \tan γ + \sin α \sin β \cos μ \tan γ + \cos α \tan β)) \end{matrix};

The derivative of rolling angular speed

\overset{\cdot}{p} = \frac{I_{z_{b}} l + I_{z_{b} x_{b}} n}{I_{x_{b}} I_{z_{b}} - I_{z_{b} x_{b}}^{2}} + \frac{I_{y_{b}} I_{z_{b}} - I_{z_{b}}^{2} - I_{z_{b} x_{b}}^{2}}{I_{x_{b}} I_{z_{b}} - I_{z_{b} x_{b}}^{2}} q r + \frac{I_{z_{b} x_{b}} (I_{x_{b}} - I_{y_{b}} + I_{z_{b}})}{I_{x_{b}} I_{z_{b}} - I_{z_{b} x_{b}}^{2}} p q;

The derivative of pitch rate

The derivative of yawrate

\overset{\cdot}{r} = \frac{I_{z_{b} x_{b}} l + I_{x_{b}} n}{I_{x_{b}} I_{z_{b}} - I_{z_{b} x_{b}}^{2}} + \frac{I_{x_{b}}^{2} - I_{x_{b}} I_{y_{b}} + I_{z_{b} x_{b}}^{2}}{I_{x_{b}} I_{z_{b}} - I_{z_{b} x_{b}}^{2}} p q + \frac{I_{z_{b} x_{b}} (I_{y_{b}} - I_{z_{b}} - I_{x_{b}})}{I_{x_{b}} I_{z_{b}} - I_{z_{b} x_{b}}^{2}} q r;

L is total rolling moment of aircraft measurement in real time；

M is total pitching moment of aircraft measurement in real time；

N is total yawing of aircraft measurement in real time；

It is around x_bThe rotary inertia of axle；

It is around y_bThe rotary inertia of axle；

It is around z_bThe rotary inertia of axle；

For x_bAxle and z_bThe product of inertia of axle；

γ is trajectory tilt angle, and unit is rad；

Sin γ is the sine of trajectory tilt angle；

Cos γ is the cosine of trajectory tilt angle；

Tan γ is the tangent of trajectory tilt angle；

χ is trajectory deflection angle, and unit is rad；

Sin χ is the sine of trajectory deflection angle；

Cos χ is the cosine of trajectory deflection angle

Tan χ is the tangent of trajectory deflection angle

λ is highly lower latitude that aircraft is located, and unit is rad；

Sin λ for aircraft be located highly under latitude sine；

Cos λ for aircraft be located highly under latitude cosine；

Tan λ for aircraft be located highly under latitude tangent；

α is the angle of attack, and unit is rad；

Sin α is the sine of the angle of attack；

Cos α is to attack cosine of an angle；

Tan α is to attack tangent of an angle；

β is yaw angle, and unit is rad；

Sin β is the sine of yaw angle；

Cos β is sideslip cosine of an angle；

Tan β is sideslip tangent of an angle；

μ is angle of heel, and unit is rad；

Sin μ is the sine of angle of heel；

Cos μ is tilt cosine of an angle；

Tan μ is to tilt tangent of an angle.

5. the closed-loop simulation system being applied to reentry vehicle gesture stability according to claim 1 it is characterised in that：Greatly Height according to residing for aircraft flight for compression ring border model unit (201), can obtain residing for aircraft temperature under highly, sound Speed, pressure and density；For aircraft dynamic pressure, aerodynamic force, aerodynamic moment resolving；Attitude command signal generating unit (202) foundation The speed of aircraft and be located highly under the velocity of sound, obtain Mach numberAccording to Mach number Ma and flight speed V, Angle of attack is provided by logical judgment_c, sideslip angle beta_c=0 and angle of heel μ_cInstruction；

6. the closed-loop simulation system being applied to reentry vehicle gesture stability according to claim 1 it is characterised in that：Fly Row device aerodynamic prediction unit (203) is used for estimating aerodynamic coefficient suffered by aircraft, aerodynamic moment coefficient, the gas that will resolve Power is supplied to slow circuit controls and restrains the nonlinear terms that unit (204) is used for resolving compensation control, carries out slow circuit controls rule solution Calculate；The aerodynamic moment coefficient of estimation is supplied to pneumatic rudder face control allocation unit (207) and control integrated unit (210)；? Pneumatic rudder face controls in allocation unit (207) unit for resolving the control allocation matrix of pneumatic rudder face；Controlling integrated unit (210) it is used for the resolving of RCS Torque Control compensation vector in unit；In slow circuit controls rule unit (204) resolve, need root It is estimated that flight vehicle aerodynamic power calculating nonlinear compensation item, control in the design of allocation unit (207) in pneumatic rudder face, need To estimate to calculate pneumatic control surface deflection instruction according to the flight vehicle aerodynamic moment coefficient of estimation；Controlling integrated unit (210) The calculating of RCS aerodynamic moment compensation vector in need to resolve RCS torque compensation vector according to aerodynamic moment coefficient.

7. the closed-loop simulation system being applied to reentry vehicle gesture stability according to claim 1 it is characterised in that： RCS controls allocation unit (208) to receive from pneumatic rudder face and controls the RCS control moment instruction of allocation unit (207) and be derived from Control the RCS compensating torque vector information of integrated unit (210), according to RCS layout designs, complete RCS roll channel, pitching is led to The thruster switch order of road and jaw channel generates, and provides the switch order of 16 thrusters；And opening each thruster Close instruction output to controlling integrated unit (210), directly control the working condition of 16 thrusters；

Wherein, split axle controls and is assigned as：Reentry vehicle is along x_bAxle rolling, and the aircraft left side occurs motion upwards, the right to occur Move downward as positive rolling；Reentry vehicle is along x_bAxle rolling, and the aircraft left side moves downward, the right occurs to transport upwards Move as bearing rolling；Reentry vehicle is along y_bAxle pitching, and Vehicle nose is just to bow upwards；Reentry vehicle is along y_bAxle pitching, and Vehicle nose is downwards negative bowing；Reentry vehicle is along z_bAxle swings, and Vehicle nose navigates to the left side for negative bias；Ablated configuration Device is along z_bAxle swings, and Vehicle nose is just to go off course to the right；

Wherein, multiple-rank control is assigned as：The class rank of three axle controls is 5 grades, and corresponding class word is -2, -1,0,1 and 2 grade.

8. the closed-loop simulation system being applied to reentry vehicle gesture stability according to claim 1 it is characterised in that：Imitative True judging unit (302) first aspect is to the elevation information H receiving and given thresholdCarry out size judgement；Second aspect pair Velocity information V receiving and given thresholdCarry out size judgement；IfOrThen terminate to emulate；IfAndThen proceed to emulate, until meetingOrAfter terminate emulate；Simulation result enters Row real time demonstration.