CN108984907A - A kind of interative guidance method based on yaw corner condition - Google Patents

A kind of interative guidance method based on yaw corner condition Download PDF

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CN108984907A
CN108984907A CN201810789782.XA CN201810789782A CN108984907A CN 108984907 A CN108984907 A CN 108984907A CN 201810789782 A CN201810789782 A CN 201810789782A CN 108984907 A CN108984907 A CN 108984907A
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韦常柱
崔乃刚
琚啸哲
李源
刁尹
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Harbin Institute of Technology
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Abstract

The invention proposes a kind of interative guidance methods based on yaw corner condition, calculate and are converted into the calculating of vector under guidance coordinate system including the vector under launching inertial system, first and second integral calculation for estimating residual non-uniformity, thrust item and gravitation item, convert to the program angle under the solution of program ascent and two coordinate systems.This interative guidance method can adapt to initial large deviation or other it is many due to and guidance problems under the conditions of the big yaw angle of bring, it is simple to guidance command structure, engineering can practicality it is strong, interative guidance adaptation of methods can be promoted, and the advanced delivery technology including entering the orbit including VTOL, adaptively etc. for China's future development provides technical support.

Description

A kind of interative guidance method based on yaw corner condition
Technical field
The invention belongs to guide and control technology field, more particularly to a kind of iteration system based on yaw corner condition Guiding method.
Background technique
Conventional iterative method of guidance is high with its guidance precision, task compatibility is strong, flight software is simple on arrow, the offline set of data The relatively low advantage of preparation requirement is widely applied in vehicle injection guidance problem.This method is derived from optimal control System is theoretical, using optimal-fuel as performance indicator, according to the real-time status online resolution posture of vehicle itself navigation system offer Angle instruction, the i.e. optimal thrust vectoring direction of fuel needed for completion aerial mission, it is final to guarantee shutdown moment terminal velocity, position It is satisfied Deng five item constraints in six state constraints.This method flies in " Saturn 5 " heavy launcher in the U.S., space flight Machine, Arian's rocket series of European Space Agency, " energy number " heavy launcher of Russia etc. are applied.
And the continuous development of application model and delivery technology with vehicle, more and more classes problem of entering the orbit emerge Come, the application field of interative guidance method is also constantly being expanded accordingly.To have/the VTVL of vertical landing ability that takes off vertically For (Vertical Takeoff Vertical Landing, VTOL) Control System for Reusable Launch Vehicle, the type vehicle In the case where returning to whole process and need to undergo pose adjustment section, boosting return phase, high-altitude descending section, the dynamic braking section in high-altitude, endoatmosphere Section and vertical landing section drop.Wherein boosting return phase is to return to the whole inflight phase controlled for the first time flight path, for Voyage amendment, terminal location are adjusted and the decomposition of subsequent each section of precision chain is most important.And boosting return phase guidance problems are substantial It is that class is entered the orbit problem, this section of Celestial Guidance Scheme design need to guarantee that vehicle grade accurately reaches subsequent high-altitude descending section, and Burnout Speed state is identical as the speed state of same point on nominal trajectory, that is, is successfully entered the high-altitude descending Duan Weiyi with nominal trajectory Section virtual rail, therefore can be designed based on the method for interative guidance.
However, big initial deviation or other it is many due to (such as VTVL Control System for Reusable Launch Vehicle boosting return phase Guidance target with earth rotation), generally require to fly under the conditions of big course angle in practical projects, at this time conventional iterative system The small yaw angle used in guiding method derivation process is assumed no longer to be applicable in, and the remaining working time of real-time estimation is difficult to ensure presentation Convergent tendency can even dissipate, and then lead to the failure of guidance task, therefore the adaptability of conventional iterative method of guidance drops significantly It is low.
Based on above-mentioned application background, a kind of interative guidance method based on yaw corner condition is proposed, to improve conventional iterative The adaptability of method of guidance, and successfully realize that VTOL, launch mission such as adaptively adjust at a variety of advanced fortune for China's future Load technology provides certain technical support.
Summary of the invention
Object of the present invention is to for enter the orbit problem under the conditions of yaw angle or class enter the orbit problem provide it is a kind of based on yaw corner condition Interative guidance method.This method can be widely used in include VTOL Control System for Reusable Launch Vehicle boosting return phase guidance, Class including spacecraft orbit reentry guidance, carrier rocket powered phase guidance etc. is entered the orbit problem or problem of entering the orbit.
The purpose of the present invention is achieved through the following technical solutions: a kind of interative guidance method based on yaw corner condition, packet Include following steps:
Step 1: the position vector and velocity vector that GPS/INS measurement obtains vehicle under launching inertial system are carried by arrow；
Step 2: in conjunction with the expectation target point position R of taskocffWith speed Vocff, guidance coordinate system is established, and pass through appearance Real time position under launching inertial system and velocity vector are converted position vector R under guidance coordinate system by state transition matrixocf0And speed Spend vector Vocf0；
Step 3: estimation residual non-uniformity tg；
Step 4: residual non-uniformity and yaw angle based on estimation are assumed to carry out boosting return phase thrust item one, two First and second integral calculation of secondary integral and gravitation item；
Step 5: control program ascent is solved；
Step 6: it will be guided under coordinate system using pose transformation matrix and resolve the program angle obtained and be converted into launching inertial system Lower pitch program angleWith yaw program angle ψT, then this, which guides vehicle in period and flies according to launching inertial system program angle, is It can.
Further, the step 3 specifically:
Step 1: setting residual non-uniformity as tg, resolve the speed increment generated by engine are as follows:
Wherein Vxocff、Vyocff、VzocffFor three axis components of target spot speed under guidance coordinate system, Vxocf0、Vyocf0、Vzocf0 For three axis components of real-time speed under guidance coordinate system, gxocf、gyocf、gzocfIt is three of average gravitational acceleration in the case where guidance is Axis component；
Step 2: can speed increment known to Paderewski formula and residual non-uniformity relational expression based on Qi Aoer are as follows:
Wherein m0Real-time quality for vehicle in each guidance period, m indicate the gross mass of vehicle, IspTo start Machine specific impulse,For engine second consumption, F indicates motor power, and t is the time；
Step 3: deforming to previous step relational expression, residual non-uniformity t is estimatedg1Are as follows:
Wherein
Step 4: if the residual non-uniformity t of estimationg1With tgMeet | tg-tg1| < ε, wherein ε is that given precision is wanted It asks, then residual non-uniformity is tg1, otherwise, continue in next step；
Step 5: by tg1Assignment gives tg, and return to the first step.
Further, the step 4 specifically:
Step 1: considering optimal control theory and replacing accurate optimal solution, setting guidance using near-optimization parsing form It is lower posture program angle form are as follows:
Wherein,And ψocfIndicate the lower pitch program angle of guidance system and yaw program angle,WithAs program angle A part is used to carry out terminal velocity state constraint, andAnd Kψ2t-Kψ1Then it is set to meet terminal for controlling vehicle 2 location status constraints,WithIndicate that pitch program angle changes slope, Kψ1And Kψ2Indicate that yaw program angle changes slope；
Step 2: the primary and secondary integral for carrying out thrust item derives based on yaw corner condition, can obtain:
Step 3: above formula is unfolded, then can obtain:
Wherein:
F0(tg)=Ispln(th/(th-tg))
F1(tg)=thF0(tg)-Isptg
F2(tg)=F0(tg)tg-F1(tg)
F3(tg)=F2(tg)th-(tg)2Isp/2
F4(tg)=Isp(th)2ln(th/(th-tg))-Isp(tg)2/2-Ispthtg
F5(tg)=Isp(th)2tgln(th/(th-tg))-Isp(th)3ln(th/(th-tg))-Isp(tg)3/6-Isp(tg)2th/2
Meanwhile
Step 4: use is averaged, gravitational method carries out gravitation integral calculation, then the primary integral and quadratic integral of gravitation item are as follows:
Further, the step 5 specifically:
Step 1: considering constraint of velocity, solved using speed increment Δ VWithAre as follows:
Step 2: considering position constraint, resolved using target point terminal location, velocity information and real time position velocity information It obtains:
Wherein, ZocffIndicate that guidance is the Z-direction position of lower target point, Zocf0Indicate real-time Z-direction position under guidance system, Yocff Indicate that guidance is the Y-direction position of lower target point, Yocf0Indicate real-time Y-direction position under guidance system, A indicates intermediate variable；
Step 3: pitch program angle and yaw program angle instruction under guidance coordinate system are as follows:
WhereinWithFor meeting terminal velocity state constraint, and And Kψ2t-Kψ1Then for meeting position constraint.
Further, the step 6 specifically:
Program angle under guidance coordinate system is converted with guidance coordinate system pose transforming relationship using launching inertial system:
IB_1=M_GltogT*IB
ψT=-arcsin (IB_1 (3))
Wherein M_Gltog is the pose transformation matrix that coordinate system is led in the launching inertial system transformation of ownership,WithAs emit inertia It is lower pitch program angle and yaw program angle.
The invention has the advantages that:
The invention proposes a kind of interative guidance methods based on yaw corner condition.It is to hold that this method, which is devised with engine, Row mechanism carries the vehicle state of navigation system output and the target three-point state information of bookbinding as input quantity, without small partially using arrow Boat angle is assumed to have derived the interative guidance expression formula for considering yaw corner condition, and then has obtained zero-miss guidance instruction.
This interative guidance method can adapt to initial large deviation or other it is many due to (such as target point with the earth turn It is dynamic) and guidance problems under the conditions of bring yaw angle, it is simple to guidance command structure, engineering can practicality it is strong, iteration can be promoted The adaptability of method of guidance, and for China's future development including VTOL, adaptively enter the orbit etc. including advanced delivery technology Technical support is provided.
Detailed description of the invention
Fig. 1 is that the present invention is based on the flow charts of the interative guidance method of yaw angle.
Specific embodiment
Technical solution in the embodiment of the present invention that following will be combined with the drawings in the embodiments of the present invention carries out clear, complete Ground description, it is clear that described embodiments are only a part of the embodiments of the present invention, instead of all the embodiments.Based on this Embodiment in invention, every other reality obtained by those of ordinary skill in the art without making creative efforts Example is applied, shall fall within the protection scope of the present invention.
In conjunction with Fig. 1, the present invention proposes a kind of interative guidance method based on yaw corner condition, comprising the following steps:
Step 1: the position vector and velocity vector that GPS/INS measurement obtains vehicle under launching inertial system are carried by arrow；
Step 2: in conjunction with the expectation target point position R of taskocffWith speed Vocff, guidance coordinate system is established, and pass through appearance Real time position under launching inertial system and velocity vector are converted position vector R under guidance coordinate system by state transition matrixocf0And speed Spend vector Vocf0；
Step 3: estimation residual non-uniformity tg；
Step 4: residual non-uniformity and yaw angle based on estimation are assumed to carry out boosting return phase thrust item one, two First and second integral calculation of secondary integral and gravitation item；
Step 5: control program ascent is solved；
Step 6: it will be guided under coordinate system using pose transformation matrix and resolve the program angle obtained and be converted into launching inertial system Lower pitch program angleWith yaw program angleThen vehicle flies according to launching inertial system program angle in this guidance period ?.
Described program angle is pitch program angle and yaw program angle.Ocf indicates the definition under guidance coordinate system, and 0 indicates just Initial value, f indicate terminal value.
The step 3 specifically:
Step 1: setting residual non-uniformity as tg, resolve the speed increment generated by engine are as follows:
Wherein Vxocff、Vyocff、VzocffFor three axis components of target spot speed under guidance coordinate system, Vxocf0、Vyocf0、Vzocf0 For three axis components of real-time speed under guidance coordinate system, gxocf、gyocf、gzocfIt is three of average gravitational acceleration in the case where guidance is Axis component；
Step 2: can speed increment known to Paderewski formula and residual non-uniformity relational expression based on Qi Aoer are as follows:
Wherein m0Real-time quality for vehicle in each guidance period, m indicate the gross mass of vehicle, IspTo start Machine specific impulse,For engine second consumption, F indicates motor power, and t is the time；The gross mass of vehicle herein is not spy Refer to the quality at a certain moment, and refers to from 0 moment to tgIn the generation that vehicle gross mass is inscribed when all in moment integral process, claims.
Step 3: deforming to previous step relational expression, residual non-uniformity t is estimatedg1Are as follows:
Wherein
Step 4: if the residual non-uniformity t of estimationg1With tgMeet | tg-tg1| < ε, wherein ε is that given precision is wanted It asks, then residual non-uniformity is tg1, otherwise, continue in next step；
Step 5: by tg1Assignment gives tg, and return to the first step.
The step 4 specifically:
Step 1: considering optimal control theory and replacing accurate optimal solution, setting guidance using near-optimization parsing form It is lower posture program angle form are as follows:
Wherein,And ψocfIndicate the lower pitch program angle of guidance system and yaw program angle,WithAs program angle A part is used to carry out terminal velocity state constraint, andAnd Kψ2t-Kψ1Then it is set to meet terminal for controlling vehicle (for the general lower terminal X of guidance system that decontrols to position constraint, selection meets the lower terminal Y-direction of guidance system and Z-direction for 2 location status constraints Position constraint),WithIndicate that pitch program angle changes slope, Kψ1And Kψ2Indicate that yaw program angle changes slope；
Step 2: the primary and secondary integral for carrying out thrust item derives based on yaw corner condition, can obtain:
Step 3: above formula is unfolded, then can obtain:
Wherein:
F0(tg)=Ispln(th/(th-tg))
F1(tg)=thF0(tg)-Isptg
F2(tg)=F0(tg)tg-F1(tg)
F3(tg)=F2(tg)th-(tg)2Isp/2
F4(tg)=Isp(th)2ln(th/(th-tg))-Isp(tg)2/2-Ispthtg
F5(tg)=Isp(th)2tgln(th/(th-tg))-Isp(th)3ln(th/(th-tg))-Isp(tg)3/6-Isp(tg)2th/2
Meanwhile
Step 4: use is averaged, gravitational method carries out gravitation integral calculation, then the primary integral and quadratic integral of gravitation item are as follows:
The step 5 specifically:
Step 1: considering constraint of velocity, solved using speed increment Δ VWithAre as follows:
Step 2: considering position constraint, resolved using target point terminal location, velocity information and real time position velocity information It obtains:
Wherein, ZocffIndicate that guidance is the Z-direction position of lower target point, Zocf0Indicate real-time Z-direction position under guidance system, Yocff Indicate that guidance is the Y-direction position of lower target point, Yocf0Indicate real-time Y-direction position under guidance system, A indicates intermediate variable；
Step 3: pitch program angle and yaw program angle instruction under guidance coordinate system are as follows:
WhereinWithFor meeting terminal velocity state constraint, and WithThen for meeting position constraint.
The step 6 specifically:
Program angle under guidance coordinate system is converted with guidance coordinate system pose transforming relationship using launching inertial system:
IB_1=M_GltogT*IB
ψT=-arcsin (IB_1 (3))
Wherein M_Gltog is the pose transformation matrix that coordinate system is led in the launching inertial system transformation of ownership,And ψTAs emit inertia It is lower pitch program angle and yaw program angle.
Above to a kind of interative guidance method based on yaw corner condition provided by the present invention, it is described in detail, Used herein a specific example illustrates the principle and implementation of the invention, and the explanation of above embodiments is only used In facilitating the understanding of the method and its core concept of the invention；At the same time, for those skilled in the art, according to the present invention Thought, there will be changes in the specific implementation manner and application range, in conclusion the content of the present specification should not be construed as Limitation of the present invention.

Claims (5)

1. a kind of interative guidance method based on yaw corner condition, which comprises the following steps:
Step 1: the position vector and velocity vector that GPS/INS measurement obtains vehicle under launching inertial system are carried by arrow；
Step 2: in conjunction with the expectation target point position R of taskocffWith speed Vocff, guidance coordinate system is established, and turn by posture It changes matrix and converts position vector R under guidance coordinate system for real time position under launching inertial system and velocity vectorocf0It is sweared with speed Measure Vocf0；
Step 3: estimation residual non-uniformity tg；
Step 4: residual non-uniformity and yaw angle based on estimation are assumed to carry out the product of boosting return phase thrust item first and second Divide first and second integral calculation with gravitation item；
Step 5: control program ascent is solved；
Step 6: it will be guided under coordinate system using pose transformation matrix and resolve the program angle obtained and be converted into launching inertial system nutation Face upward program angleWith yaw program angle ψT, then vehicle flies according to launching inertial system program angle in this guidance period.
2. the method according to claim 1, wherein the step 3 specifically:
Step 1: setting residual non-uniformity as tg, resolve the speed increment generated by engine are as follows:
Wherein Vxocff、Vyocff、VzocffFor three axis components of target spot speed under guidance coordinate system, Vxocf0、Vyocf0、Vzocf0For system Lead three axis components of real-time speed under coordinate system, gxocf、gyocf、gzocfFor three axis point of the average gravitational acceleration under guidance system Amount；
Step 2: can speed increment known to Paderewski formula and residual non-uniformity relational expression based on Qi Aoer are as follows:
Wherein m0Real-time quality for vehicle in each guidance period, m indicate the gross mass of vehicle, IspFor engine ratio Punching,For engine second consumption, F indicates motor power, and t is the time；
Step 3: deforming to previous step relational expression, residual non-uniformity t is estimatedg1Are as follows:
Wherein
Step 4: if the residual non-uniformity t of estimationg1With tgMeet | tg-tg1| < ε, wherein ε is given required precision, then Residual non-uniformity is tg1, otherwise, continue in next step；
Step 5: by tg1Assignment gives tg, and return to the first step.
3. according to the method described in claim 2, it is characterized in that, the step 4 specifically:
Step 1: considering optimal control theory and replacing accurate optimal solution using near-optimization parsing form, set under guidance system Posture program angle form are as follows:
Wherein,And ψocfIndicate the lower pitch program angle of guidance system and yaw program angle,WithA part as program angle For carrying out terminal velocity state constraint, andAnd Kψ2t-Kψ1Then it is set to meet 2 positions of terminal for controlling vehicle State constraint is set,WithIndicate that pitch program angle changes slope, Kψ1And Kψ2Indicate that yaw program angle changes slope；
Step 2: the primary and secondary integral for carrying out thrust item derives based on yaw corner condition, can obtain:
Step 3: above formula is unfolded, then can obtain:
Wherein:
F0(tg)=Ispln(th/(th-tg))
F1(tg)=thF0(tg)-Isptg
F2(tg)=F0(tg)tg-F1(tg)
F3(tg)=F2(tg)th-(tg)2Isp/2
F4(tg)=Isp(th)2ln(th/(th-tg))-Isp(tg)2/2-Ispthtg
F5(tg)=Isp(th)2tgln(th/(th-tg))-Isp(th)3ln(th/(th-tg))-Isp(tg)3/6-Isp(tg)2th/2
Meanwhile
Step 4: use is averaged, gravitational method carries out gravitation integral calculation, then the primary integral and quadratic integral of gravitation item are as follows:
4. according to the method described in claim 3, it is characterized in that, the step 5 specifically:
Step 1: considering constraint of velocity, solved using speed increment Δ VWithAre as follows:
Step 2: considering position constraint, resolved using target point terminal location, velocity information and real time position velocity information Out:
Wherein, ZocffIndicate that guidance is the Z-direction position of lower target point, Zocf0Indicate real-time Z-direction position under guidance system, YocffIt indicates Guidance is the Y-direction position of lower target point, Yocf0Indicate real-time Y-direction position under guidance system, A indicates intermediate variable；
Step 3: pitch program angle and yaw program angle instruction under guidance coordinate system are as follows:
WhereinWithFor meeting terminal velocity state constraint, andWith Kψ2t-Kψ1Then for meeting position constraint.
5. according to the method described in claim 4, it is characterized in that, the step 6 specifically:
Program angle under guidance coordinate system is converted with guidance coordinate system pose transforming relationship using launching inertial system:
IB_1=M_GltogT*IB
ψT=-arcsin (IB_1 (3))
Wherein M_Gltog is the pose transformation matrix that coordinate system is led in the launching inertial system transformation of ownership,And ψTAs launching inertial system nutation Face upward program angle and yaw program angle.
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CN109798902A (en) * 2019-03-11 2019-05-24 北京星际荣耀空间科技有限公司 One kind being suitable for carrier rocket and enters the orbit modified interative guidance method
CN109798902B (en) * 2019-03-11 2020-09-22 北京星际荣耀空间科技有限公司 Iterative guidance method suitable for carrier rocket in-orbit correction
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CN110908396A (en) * 2019-11-28 2020-03-24 天津大学 Full-stage reentry return guidance method for reusable vehicle
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CN110775301A (en) * 2019-12-04 2020-02-11 哈尔滨工业大学 Aircraft with high rail-entering efficiency and strong maneuvering capability and rail-entering method thereof
CN111272173A (en) * 2020-02-20 2020-06-12 哈尔滨工业大学 Gradient solving iterative guidance method considering earth rotation and large yaw angle
CN112306075A (en) * 2020-10-20 2021-02-02 中国运载火箭技术研究院 High-precision off-orbit reverse iterative guidance method
CN112306075B (en) * 2020-10-20 2023-08-29 中国运载火箭技术研究院 High-precision off-track reverse iteration guidance method
CN112034703B (en) * 2020-11-03 2021-03-19 蓝箭航天空间科技股份有限公司 Self-adaptive iterative guidance method and device for spacecraft
WO2022095643A1 (en) * 2020-11-03 2022-05-12 蓝箭航天空间科技股份有限公司 Self-adaptive iterative guidance method and device for aerospace vehicle
CN112034703A (en) * 2020-11-03 2020-12-04 蓝箭航天空间科技股份有限公司 Self-adaptive iterative guidance method and device for spacecraft
GB2616530A (en) * 2020-11-03 2023-09-13 Landspace Science & Tech Co Ltd Self-adaptive iterative guidance method and device for aerospace vehicle
CN113022893A (en) * 2021-02-26 2021-06-25 北京控制工程研究所 Space rendezvous interception autonomous self-adaptive remote guidance method and system
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