CN109344449A - The spacecraft moon ground transfer orbit Reverse Design - Google Patents
The spacecraft moon ground transfer orbit Reverse Design Download PDFInfo
- Publication number
- CN109344449A CN109344449A CN201811041353.0A CN201811041353A CN109344449A CN 109344449 A CN109344449 A CN 109344449A CN 201811041353 A CN201811041353 A CN 201811041353A CN 109344449 A CN109344449 A CN 109344449A
- Authority
- CN
- China
- Prior art keywords
- moon
- orbit
- transfer orbit
- earth
- search
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Classifications
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
Landscapes
- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Theoretical Computer Science (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- Geometry (AREA)
- General Engineering & Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
- Navigation (AREA)
Abstract
The present invention relates to a kind of spacecraft moon transfer orbit Reverse Design, comprising steps of a. with establishing moon transfer orbit backstepping flight process, chooses variable control variable；B. alignment moon search is carried out, with determining the moon transfer orbit near end orbital plane；C. it carries out lunar orbit to set out restriction on the parameters search, determining the moon the nearly moon end orbital plane of transfer orbit；D. with carrying out moon transfer orbit search, meets that landing field is coplanar and reentry point constraint；E. track emulates, with obtaining the moon transfer orbit.The spacecraft moon according to the present invention transfer orbit Reverse Design use track backstepping method, solution procedure is simple, and search speed is fast, and convergence is good.
Description
Technical field
The present invention relates to the spacecraft moon transfer orbit technical field is related to a kind of spacecraft from the moon and returns to earth mistake
The moon is carried out in journey the method more particularly to a kind of ground transfer orbit Reverse Design of spacecraft moon of transfer orbit design.
Background technique
Month ground transfer orbit refer to that spacecraft shifts acceleration with carrying out the moon on lunar orbit after, escape the moon simultaneously return
The track of the earth.Return task is sampled for manned moon landing's task or the moon, lunar surface task is completed and is returning to lunar orbit
Afterwards, with needing to design suitable moon transfer orbit makes spacecraft apply pulse in given time and is transferred to the earth from lunar orbit
Boundary point is reentered, and returns to landing field safely using atmospheric breaking.If moon ground transfer orbit design mistake, spacecraft without
Method meets reentry condition, also just can not safe landing ground.Therefore, the moon transfer orbit design for manned moon landing and the moon
Sampling return task is extremely important, is personnel and the basis that secure sample returns.
Moon ground transfer orbit design relates generally to reentry point constraint, including reenters point height, reentry angle, reentry point track
Inclination angle and with landing field coplanar constraint.Month ground transfer orbit initial point be its perilune, the orbit altitude of perilune and
Orbit inclination angle is identical as the orbit altitude of initial lunar orbit and orbit inclination angle.Currently, with the development of moon exploration task,
Transfer orbit design achieves considerable progress between the ground moon, but the research based on Earth-moon transfer orbit is in the majority.For the moon shift
The research of track, domestic scholars inherit the design method of Earth-moon transfer orbit mostly, are derived just based on double two body Models
The relationship of beginning condition and end conswtraint adjusts control variable constantly by alternative manner to meet final constraint condition.Into
And using the moon under double two body Models transfer orbit then is modified iteration and shifts with obtaining the moon under accurate model as initial value
Track.There are two deficiencies for this method: 1) solution procedure is complicated, and needs in double two body Models and accurate model respectively
Iterative solution；It 2) is that control variable carries out the design of track forwards, and the selection of moon heart section parameter is without rule with moon heart section parameter
Property, with can not reflecting the moon geometrical property of transfer orbit, and variable can not be controlled moon heart section carry out simple initial value and estimate
Meter.
Summary of the invention
The purpose of the present invention is to solve the above problem, and transfer orbit is inversely set with providing a kind of high-precision spacecraft moon
Meter method.
For achieving the above object, the present invention a kind of spacecraft moon is provided transfer orbit Reverse Design, including
Following steps:
A. with establishing moon transfer orbit backstepping flight process chooses variable control variable；
B. alignment moon search is carried out, with determining the moon transfer orbit near end orbital plane；
C. it carries out lunar orbit to set out restriction on the parameters search, determining the moon the nearly moon end orbital plane of transfer orbit；
D. with carrying out moon transfer orbit search, meets that landing field is coplanar and reentry point constraint；
E. track emulates, with obtaining the moon transfer orbit.
According to an aspect of the present invention, in the b step, alignment moon search is carried out, differential revised law tune is passed through
Whole near-earth berth circular orbit right ascension of ascending node, the re-entry deceleration moment moon rotation be right ascension and the moon ground geard-down speed increment, alignment
The goal constraint of moon search is the right ascension declination difference at perilune moment, perilune and the moon.
According to an aspect of the present invention, in the step c, the near-earth obtained in the b step is berthed circle first
Ascending node of orbit right ascension, the re-entry deceleration moment the moon rotation be right ascension and the moon geard-down speed increment replacement be originally worth；
Then carry out lunar orbit set out restriction on the parameters search, by differential correct adjustment near-earth berth circular orbit rise hand over
Point right ascension, the re-entry deceleration moment moon rotation be right ascension and the moon ground geard-down speed increment, lunar orbit sets out restriction on the parameters search
Goal constraint be perilune moment, perilune height and perilune track moon heart inclination angle.
According to an aspect of the present invention, in the Step d, the near-earth obtained in the step c is berthed circle first
Ascending node of orbit right ascension, the re-entry deceleration moment the moon rotation be right ascension and the moon geard-down speed increment replacement be originally worth；
Then with carrying out the accurate moon transfer orbit search for, by differential correct adjustment near-earth berth circular orbit terminal juncture,
Near-earth berth circular orbit semi-major axis, near-earth berth circular orbit right ascension of ascending node, the re-entry deceleration moment the moon rotation be right ascension and the moon
Ground geard-down speed increment, the accurate moon transfer orbit search goal constraint be perilune moment, perilune height, perilune
Track moon heart inclination angle, landing field coplanar constraint and reenter point height.
According to an aspect of the present invention, in the Step d, the near-earth obtained in step 4 is berthed circular orbit first
Terminal juncture, near-earth berth circular orbit semi-major axis, near-earth berth circular orbit right ascension of ascending node, the re-entry deceleration moment the moon rotation
Be right ascension and the moon geard-down speed increment replacement be originally worth；
Then carry out track emulation and extract perilune to reentry point section orbital data, with obtaining moon transfer orbit
Parameter.
The spacecraft moon according to the present invention transfer orbit Reverse Design the moon is utilized using the method for track backstepping
The geometrical property of ground transfer orbit, using the earth's core section parameter as starting point, to the moon transfer orbit carries out reverse orbit integration, passes through
With rationally designing moon transfer orbit search routine, transfer orbit controls variable with gradually determining the moon, utilizes Track desigh personnel
Simple differential corrections can directly in high-precision kinetic model fast search to the moon transfer orbit.This hair
Bright method solution procedure is simple, and search speed is fast, and convergence is good.
The spacecraft moon according to the present invention ground transfer orbit Reverse Design, solution procedure is simple, does not need double two
It is iteratively solved respectively in body Model and accurate model, and with being able to reflect out the moon geometrical property of transfer orbit, it can be to the moon
Heart section controls variable and carries out simple initial estimate.
Detailed description of the invention
It in order to more clearly explain the embodiment of the invention or the technical proposal in the existing technology, below will be in embodiment
Required attached drawing is briefly described, it should be apparent that, the accompanying drawings in the following description is only some realities of the invention
Example is applied, it for those of ordinary skill in the art, without creative efforts, can also be attached according to these
Figure obtains other attached drawings.
The flow chart of Fig. 1 with schematically showing the spacecraft moon according to the present invention transfer orbit Reverse Design；
The track backstepping procedure chart of Fig. 2 with schematically showing the moon according to the present invention transfer orbit；
Fig. 3 schematically shows that ground moon rotation is right ascension figure；
Fig. 4 is with schematically showing the earth's core J2000 coordinate system moon that b step searches in one embodiment of the present invention
Transfer orbit figure；
Fig. 5 schematically show that step c in one embodiment of the present invention searches the moon heart inertial coodinate system turn the moon
Move trajectory diagram；
Fig. 6 A~6B with schematically showing the moon that Step d searches in one embodiment of the present invention transfer orbit figure；
Fig. 7 A~7C is shifted with schematically showing the moon that step e track emulates in one embodiment of the present invention
Trajectory diagram.
Specific embodiment
It, below will be to embodiment party in order to illustrate more clearly of embodiment of the present invention or technical solution in the prior art
Attached drawing needed in formula is briefly described.It should be evident that the accompanying drawings in the following description is only of the invention one
A little embodiments for those of ordinary skills without creative efforts, can also basis
These attached drawings obtain other attached drawings.
When being described for embodiments of the present invention, term " longitudinal direction ", " transverse direction ", "upper", "lower", " preceding ",
" rear ", "left", "right", "vertical", "horizontal", "top", "bottom" "inner", orientation or positional relationship is to be based on expressed by "outside"
Orientation or positional relationship shown in relevant drawings, is merely for convenience of description of the present invention and simplification of the description, rather than indicate or
Imply that signified device or element must have a particular orientation, be constructed and operated in a specific orientation, therefore above-mentioned term is not
It can be interpreted as limitation of the present invention.
The present invention is described in detail with reference to the accompanying drawings and detailed description, embodiment cannot herein one by one
It repeats, but therefore embodiments of the present invention are not defined in following implementation.
The flow chart of Fig. 1 with schematically showing the spacecraft moon according to the present invention transfer orbit Reverse Design.Such as figure
Shown in 1, the spacecraft moon according to the present invention transfer orbit Reverse Design the following steps are included:
A. with establishing moon transfer orbit backstepping flight process chooses variable control variable；
B. alignment moon search is carried out, with determining the moon transfer orbit near end orbital plane；
C. it carries out lunar orbit to set out restriction on the parameters search, determining the moon the nearly moon end orbital plane of transfer orbit；
D. with carrying out moon transfer orbit search, meets that landing field is coplanar and reentry point constraint；
E. track emulates, with obtaining the moon transfer orbit.
The track backstepping procedure chart of Fig. 2 with schematically showing the moon according to the present invention transfer orbit；Fig. 3 is schematically shown
Moon rotation in ground is right ascension figure.
As shown in Fig. 2, in above-mentioned a step of the invention, by the moon shift flight course near end to nearly moon end into
Row backstepping, and it is divided into 4 stages.1) near-earth parking orbit section, setting near-earth berth circular orbit LEO last current state [tf2,aLEO,
eLEO,iLEO,ΩLEO,ωLEO,θLEO], track backstepping to tf1Moment.Wherein, it enableseLEO=0, iLEO=
iEI, ωLEO=θLEO=0,For the moon transfer orbit perigee altitude, REarthFor earth radius.2) moon ground braking section, mould
The quasi- moon transfer orbit enter near-earth using 1 tangential deceleration pulse and berth circular orbit, if the moon geard-down speed increment beWith obtaining moon transfer orbit near-earth dotted state3) moon ground transfer orbit
Rear inflight phase is reentered, with the moon transfer orbit near-earth dotted state is end value, track backstepping to reentry point tfMoment, with obtaining the moon
Transfer orbit reenters dotted state4) moon transfer orbit reenter preceding inflight phase,
With the moon transfer orbit reenters dotted state as last value, track backstepping to perilune t0Moment, with obtaining the moon transfer orbit perilune
StateIts perilune orbit parameter meets lunar orbit restriction on the parameters,t0=tTEI。
Wherein near-earth berths circular orbit last current state moment tf2, last current state semi-major axis aLEO, last current state right ascension of ascending node ΩLEO、
Rotating the moon to the re-entry deceleration moment is right ascension(as shown in Figure 3), the moon ground geard-down speed incrementIt is unknown, as variable
Control variable, and initialization.
In above-mentioned b step of the invention, for the convergence for improving search routine, first fine tuning near end orbital plane, make
Month ground transfer orbit nearly moon end be aligned the moon.Near-earth in a step is berthed circular orbit last current state right ascension of ascending node ΩLEO、
Rotating the moon to the re-entry deceleration moment is right ascensionMonth ground geard-down speed incrementAs control variable, goal constraint is set as
Perilune moment t0=tTEI, perilune and moon ascensional difference Δ φ≤0.2 °, perilune and moon declination differenceIt is logical
It crosses differential corrections and constantly corrects ΩLEO、WithNear end orbital plane is adjusted, until meeting goal constraint item
Part, to be allowed to be directed at the moon.
In above-mentioned step c of the invention, the near-earth obtained in b step is berthed circular orbit last current state right ascension of ascending node
ΩLEO, the re-entry deceleration moment the moon rotation be right ascensionMonth ground geard-down speed incrementReplacement is originally worth, with guaranteeing the moon
Transfer orbit is inversely directed at the moon.It is red that the near-earth circular orbit last current state ascending node that berths further is finely tuned by differential corrections
Through ΩLEO, the re-entry deceleration moment the moon rotation be right ascensionMonth ground geard-down speed incrementTransfer orbit is close with making the moon
Put the perilune moment constraint t for meeting lunar orbit the moon0=tTEI, perilune it is highly constrainedWith the nearly moon
Point moon heart orbit inclination angle constraint
In above-mentioned Step d of the invention, completion lunar orbit sets out after restriction on the parameters search, then carries out near end again
Access point restriction on the parameters search.The near-earth obtained in step c is berthed circular orbit last current state right ascension of ascending node ΩLEO, re-entry deceleration
Rotating the moon to moment is right ascensionMonth ground geard-down speed incrementReplacement is originally worth, and with guaranteeing moon transfer orbit meets
Nearly moon end constraint, while control variable near-earth is added and berths circular orbit last current state moment tf2With last current state semi-major axis aLEO.It is comprehensive
Amendment near-earth berths circular orbit last current state moment tf2, last current state right ascension of ascending node ΩLEO, last current state semi-major axis aLEO, re-entry deceleration
Rotating the moon to moment is right ascensionWith the moon geard-down speed incrementTransfer orbit meets the perilune moment about with making the moon
Beam t0=tTEI, perilune it is highly constrainedIt is constrained with perilune moon heart orbit inclination angleReenter moment and landing field coplanar constraint η (tf)=pi/2 and the highly constrained h of reentry pointE(tf)=hEI。
Wherein η (tf) it is the angle for reentering moment track normal vector and landing field position vector.Shift rail to the moon that iterative search obtains
Road had not only met the lunar orbit primary condition for bringing out hair the nearly moon, but meet near end arrival to reenter point height and landing field total
Area Objects constraint.
In above-mentioned step e of the invention, near-earth will be obtained in Step d and berthed circular orbit last current state moment tf2, last current state
Right ascension of ascending node ΩLEO, last current state semi-major axis aLEO, the re-entry deceleration moment the moon rotation be right ascensionWith the moon geard-down speed
IncrementReplacement is originally worth, with carrying out the moon track emulation of transfer process, thus the moon under obtaining high-precision model turn
Orbit parameter is moved, the moon can be drawn transfer orbit flight path and acquisition key point orbit parameter.
It is as follows to provide a specific embodiment for above scheme according to the present invention:
Above-mentioned a step according to the present invention, with establishing the moon track backstepping flight process of transfer orbit.Enable initial parameter
Are as follows: set out the ring moon moment tTEI=29Jun 2014 02:00:00.000UTCG, orbit altitude hLLO=100km, the ring moon inclination angle
iLLO=160 °, terminal condition is to reenter point height hEI=120km, reentry angle γEI=-6.0 °, reentry point geocentric orbit inclination angle
iEI=20 °, landing field longitude and latitude (α=110 ° E, δ=20 ° N), the moon transfer time be about 3 days.Control initial guess difference
It is taken as tf2=2Jul 2014 02:00:00.000UTCG, aLEO=6428.137km, ΩLEO=90 °,
Above-mentioned b step according to the present invention, goal constraint are set as perilune moment t0=tTEI=29Jun 2,014 02:
00:00.000UTCG, perilune and moon ascensional difference Δ φ≤0.2 °, perilune and moon declination differenceBy micro-
Divide revised law, only 4 step of iteration just restrains, and obtains control variable ΩLEO=20.5 °,Flight path under the J2000 coordinate system of its earth's core is as shown in Figure 4.
Above-mentioned step c according to the present invention, by ΩLEO=20.5 °,Substitution
ΩLEO=90 °,Goal constraint is set as perilune moment t0=tTEI=29Jun
2014 02:00:00.000UTCG, perilune are highly constrainedIncline with perilune moon heart track
Angle constraintBy differential revised law, only 2 step of iteration just restrains, and obtains control variable ΩLEO=
20.5°、Flight path under the J2000 coordinate system of its earth's core is as shown in Figure 5.
Above-mentioned Step d according to the present invention, by ΩLEO=20.5 °,Substitution
ΩLEO=20.5 °,Goal constraint is set as perilune moment t0=tTEI=
2014 02:00:00.000UTCG of 29Jun, perilune are highly constrainedThe perilune moon heart
Orbit inclination angle constraintLanding field coplanar constraint η (tfThe highly constrained h of)=pi/2, reentry pointE
(tf)=hEI=120km.By differential revised law, only 3 step of iteration just restrains, and obtains control variable ΩLEO=20.5 °,aLEO=6430.685km, tf22014 03:36 of=13Jun:
32.591UTCG.Flight path under the J2000 coordinate system of its earth's core is as shown in Fig. 6 A~6B.By the substar rail of Fig. 6 A~6B
Mark can be seen that reentry point is coplanar with landing field.
Above-mentioned step e according to the present invention carries out the flight that track emulates with can obtaining different coordinates next month transfer orbit
Track and its earth sub-satellite track, as figs. 7 a to 7 c.
It can be seen from the above result that can be directly in the high-precision perturbation model of business software STK using method of the invention
Search for moon ground transfer orbit, and fast convergence rate in CisLunar, one month ground transfer orbit of average search only used time 30 seconds.
The above-mentioned spacecraft moon according to the present invention ground transfer orbit Reverse Design, using the method for track backstepping, benefit
With the moon the geometrical property of transfer orbit, using the earth's core section parameter as starting point, to the moon transfer orbit carries out reverse orbit integration,
By with rationally designing moon transfer orbit search routine, transfer orbit controls variable with gradually determining the moon, makes Track desigh personnel
Using simple differential corrections can directly in high-precision kinetic model fast search to the moon transfer orbit.
Method solution procedure of the invention is simple, and search speed is fast, and convergence is good.
The above-mentioned spacecraft moon according to the present invention ground transfer orbit Reverse Design, solution procedure is simple, does not need
It is iteratively solved respectively in double two body Models and accurate model, and with being able to reflect out the moon geometrical property of transfer orbit, it can
Simple initial estimate is carried out to moon heart section control variable.
The foregoing is merely an embodiment of the invention, are not intended to restrict the invention, for this field
For technical staff, the invention may be variously modified and varied.All within the spirits and principles of the present invention, made
What modification, equivalent replacement, improvement etc., should all be included in the protection scope of the present invention.
Claims (5)
1. a kind of ground transfer orbit Reverse Design of spacecraft moon, comprising the following steps:
A. with establishing moon transfer orbit backstepping flight process chooses variable control variable；
B. alignment moon search is carried out, with determining the moon transfer orbit near end orbital plane；
C. it carries out lunar orbit to set out restriction on the parameters search, determining the moon the nearly moon end orbital plane of transfer orbit；
D. with carrying out moon transfer orbit search, meets that landing field is coplanar and reentry point constraint；
E. track emulates, with obtaining the moon transfer orbit.
2. the spacecraft moon according to claim 1 ground transfer orbit Reverse Design, which is characterized in that walked in the b
In rapid, alignment moon search is carried out, near-earth is adjusted by differential revised law and is berthed circular orbit right ascension of ascending node, re-entry deceleration moment
Moon rotation in ground is right ascension and the moon ground geard-down speed increment, the goal constraint of alignment moon search be the perilune moment, perilune with
The right ascension declination difference of the moon.
3. the spacecraft moon according to claim 1 ground transfer orbit Reverse Design, which is characterized in that walked in the c
In rapid, first by the near-earth obtained in the b step berth circular orbit right ascension of ascending node, the re-entry deceleration moment moon rotation be red
Originally it is worth through the replacement of with the moon geard-down speed increment；
Then it carries out lunar orbit to set out restriction on the parameters search, it is red that the adjustment near-earth circular orbit ascending node that berths is corrected by differential
Through, re-entry deceleration moment moon rotation is right ascension and the moon ground geard-down speed increment, and lunar orbit sets out the mesh of restriction on the parameters search
Mark is constrained to perilune moment, perilune height and perilune track moon heart inclination angle.
4. the spacecraft moon according to claim 1 ground transfer orbit Reverse Design, which is characterized in that walked in the d
In rapid, first by the near-earth obtained in the step c berth circular orbit right ascension of ascending node, the re-entry deceleration moment moon rotation be red
Originally it is worth through the replacement of with the moon geard-down speed increment；
Then transfer orbit is searched for carrying out the accurate moon, is corrected adjustment near-earth by differential and is berthed circular orbit terminal juncture, near-earth
Berth circular orbit semi-major axis, near-earth berth circular orbit right ascension of ascending node, the re-entry deceleration moment the moon rotation be right ascension and the moon subtract
Fast speed increment, the accurate moon transfer orbit search goal constraint be perilune moment, perilune height, the perilune track moon
Heart inclination angle, landing field coplanar constraint and reenter point height.
5. the spacecraft moon according to claim 1 ground transfer orbit Reverse Design, which is characterized in that walked in the d
In rapid, the near-earth obtained in step 4 berth circular orbit semi-major axis, the near-earth of circular orbit terminal juncture, near-earth that berth is berthed first
Circular orbit right ascension of ascending node, the re-entry deceleration moment the moon rotation be right ascension and the moon geard-down speed increment replacement be originally worth；
Then carry out track emulation and extract perilune to reentry point section orbital data, with obtaining moon transfer orbit parameter.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811041353.0A CN109344449B (en) | 2018-09-07 | 2018-09-07 | Spacecraft monthly transfer orbit reverse design method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811041353.0A CN109344449B (en) | 2018-09-07 | 2018-09-07 | Spacecraft monthly transfer orbit reverse design method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109344449A true CN109344449A (en) | 2019-02-15 |
| CN109344449B CN109344449B (en) | 2022-02-11 |
Family
ID=65304888
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201811041353.0A Active CN109344449B (en) | 2018-09-07 | 2018-09-07 | Spacecraft monthly transfer orbit reverse design method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN109344449B (en) |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110104219A (en) * | 2019-04-24 | 2019-08-09 | 中国人民解放军63920部队 | A kind of method and device controlling detector landing objects outside Earth |
| CN110736469A (en) * | 2019-11-06 | 2020-01-31 | 北京理工大学 | Asteroid detection accurate orbit transfer method based on sun-ground rotation coordinate system |
| CN110765504A (en) * | 2019-10-29 | 2020-02-07 | 北京空间技术研制试验中心 | Orbit design method for rendezvous and docking of spacecraft orbits around the moon |
| CN112009727A (en) * | 2020-08-21 | 2020-12-01 | 北京空间技术研制试验中心 | Optimal low-thrust transfer sectional design method for translation point orbit |
| CN113086250A (en) * | 2021-03-12 | 2021-07-09 | 北京空间飞行器总体设计部 | Monthly transfer track correction method based on engineering constraints |
| CN113093776A (en) * | 2021-03-04 | 2021-07-09 | 北京航天飞行控制中心 | Method and device for determining off-orbit parameters of spacecraft |
| CN113128032A (en) * | 2021-04-01 | 2021-07-16 | 北京航空航天大学 | Intersection point time and position solving algorithm based on orbit analysis perturbation solution |
| CN113310496A (en) * | 2021-05-08 | 2021-08-27 | 北京航天飞行控制中心 | Method and device for determining lunar-ground transfer orbit |
| CN113311854A (en) * | 2021-05-19 | 2021-08-27 | 北京空间飞行器总体设计部 | Fixed-point landing orbit design method in lunar sampling return task |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103984236A (en) * | 2014-05-30 | 2014-08-13 | 哈尔滨工业大学 | Space-based dispenser different-plane orbit dispersion control method |
| WO2016126503A1 (en) * | 2015-02-08 | 2016-08-11 | Hyperloop Technologies, Inc | Transportation system |
-
2018
- 2018-09-07 CN CN201811041353.0A patent/CN109344449B/en active Active
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103984236A (en) * | 2014-05-30 | 2014-08-13 | 哈尔滨工业大学 | Space-based dispenser different-plane orbit dispersion control method |
| WO2016126503A1 (en) * | 2015-02-08 | 2016-08-11 | Hyperloop Technologies, Inc | Transportation system |
Non-Patent Citations (1)
| Title |
|---|
| 彭坤等: "基于弹道逃逸和小推力捕获的地月转移轨道设计", 《航空学报》 * |
Cited By (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110104219A (en) * | 2019-04-24 | 2019-08-09 | 中国人民解放军63920部队 | A kind of method and device controlling detector landing objects outside Earth |
| CN110765504A (en) * | 2019-10-29 | 2020-02-07 | 北京空间技术研制试验中心 | Orbit design method for rendezvous and docking of spacecraft orbits around the moon |
| CN110736469A (en) * | 2019-11-06 | 2020-01-31 | 北京理工大学 | Asteroid detection accurate orbit transfer method based on sun-ground rotation coordinate system |
| CN112009727A (en) * | 2020-08-21 | 2020-12-01 | 北京空间技术研制试验中心 | Optimal low-thrust transfer sectional design method for translation point orbit |
| CN113093776A (en) * | 2021-03-04 | 2021-07-09 | 北京航天飞行控制中心 | Method and device for determining off-orbit parameters of spacecraft |
| CN113093776B (en) * | 2021-03-04 | 2024-02-02 | 北京航天飞行控制中心 | Off-orbit parameter determination method and device for spacecraft |
| CN113086250A (en) * | 2021-03-12 | 2021-07-09 | 北京空间飞行器总体设计部 | Monthly transfer track correction method based on engineering constraints |
| CN113128032A (en) * | 2021-04-01 | 2021-07-16 | 北京航空航天大学 | Intersection point time and position solving algorithm based on orbit analysis perturbation solution |
| CN113128032B (en) * | 2021-04-01 | 2022-09-16 | 北京航空航天大学 | Intersection point time and position solving algorithm based on orbit analysis perturbation solution |
| CN113310496A (en) * | 2021-05-08 | 2021-08-27 | 北京航天飞行控制中心 | Method and device for determining lunar-ground transfer orbit |
| CN113310496B (en) * | 2021-05-08 | 2024-01-09 | 北京航天飞行控制中心 | Method and device for determining moon-earth transfer track |
| CN113311854A (en) * | 2021-05-19 | 2021-08-27 | 北京空间飞行器总体设计部 | Fixed-point landing orbit design method in lunar sampling return task |
Also Published As
| Publication number | Publication date |
|---|---|
| CN109344449B (en) | 2022-02-11 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN109344449A (en) | The spacecraft moon ground transfer orbit Reverse Design | |
| Canuto et al. | Spacecraft dynamics and control: the embedded model control approach | |
| Schmidt | Application of state-space methods to navigation problems | |
| CN103063217B (en) | Deep space detector astronomy/radio combination navigation method based on ephemeris correction | |
| CN109592079A (en) | A kind of spacecraft coplanar encounter of limiting time becomes rail strategy and determines method | |
| CN102607564A (en) | Small satellite autonomous navigation system based on starlight/ geomagnetism integrated information and navigation method thereof | |
| CN104332707A (en) | Method for tracking ground station through low earth orbit space-borne antenna | |
| CN110104219A (en) | A kind of method and device controlling detector landing objects outside Earth | |
| CN111156986B (en) | Spectrum red shift autonomous integrated navigation method based on robust adaptive UKF | |
| Yoshimitsu et al. | Hayabusa-final autonomous descent and landing based on target marker tracking | |
| CN112629543A (en) | Orbit planning method for large elliptical orbit and small-inclination-angle circular orbit | |
| Wang et al. | Absolute navigation for Mars final approach using relative measurements of X-ray pulsars and Mars orbiter | |
| CN100442015C (en) | Astronomical/doppler combined navigation method for spacecraft | |
| CN112649006A (en) | Orbit planning method for sun synchronous circular orbit | |
| Ono et al. | GNC design and evaluation of Hayabusa2 descent operations | |
| CN111427004A (en) | Coordinate conversion method suitable for pointing of ground survey station antenna to satellite | |
| CN107966149A (en) | A kind of program angle and optimizing design method of multiple constraint automated spacecraft | |
| Antreasian et al. | OSIRIS-REx Proximity Operations and Navigation Performance at Bennu | |
| CN109117543A (en) | The rail design method that manned spacecraft detects near-Earth asteroid and returns | |
| Cui et al. | Real-time navigation for Mars final approach using X-ray pulsars | |
| CN106777580B (en) | Method for rapidly designing emission window of near-earth inclined orbit | |
| Wescott et al. | L= 1.24 conjugate magnetic field line tracing experiments with barium shaped charges | |
| Muñoz et al. | Preparations and strategy for navigation during Rosetta comet phase | |
| Wood | The Evolution of Deep Space Navigation: 2006–2009 | |
| Li et al. | Launch window for manned Moon‐to‐Earth trajectories |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |