CN101354251A - Method for determining deep space detector equivalent transfer orbit - Google Patents

Abstract

The invention discloses a method for determining an equivalent transfer orbit of a deep space probe, which comprises the steps: (1) the equivalent transfer orbit of the deep space probe is designed according to an orbit dynamic model, and the orbit parameter of the designed transfer orbit is extracted; (2) the near-earth equivalent parameter of the transfer orbit of the deep space probe is solved according to the orbit parameter extracted in the step (1) and the spatial geometric relationship. By the method of the invention, when a carrier rocket directly sends a probe into the transfer orbit, the carrying party does not need to establish a complex dynamic model for the deep space probe, and the existing model used for estimating the launching precision can estimate the transmission precision more correctly, thereby solving the problem that the current carrier rocket in China does not have an upper stage and can not send the probe into the transfer orbit by using the upper stage for firing. The method of the invention adopts the concept of equivalence to solve the problem and is simple and reliable.

Description

A kind of deep space detector equivalent transfer orbit is determined method

Technical field

The present invention relates to a kind of equivalent parameters method for designing of deep space probe transfer orbit, particularly adopt carrier rocket directly to launch the equivalent parameters method for designing of the deep space probe transfer orbit that enters transfer orbit.

Background technology

Though in the world " deep space " and " survey of deep space " also do not had uniform definition at present, but in encyclopedia by American-European countries and the Aero-Space of dictionary and the Encyclopadia Sinica volume to the definition of deep space, and International Telecommunication Association is when communication frequency carries out regulation in the usage space detected event, to defining that " deep space " and " non-deep space " carries out, all can find: the distance of the detection of a target of detector and the earth is more than or equal to ground moon distance in the survey of deep space activity.

In the world to the defining of " deep space ", the detection of a target of deep space probe is remote apart from the earth by above-mentioned.Arrive the final detection of a target, its transfer orbit must be to start near the earth, and near the remote finally detection of a target.This shows that detector need carry out for a long time, flight at a distance on transfer orbit.Detector not only is subjected to the terrestrial gravitation effect on transfer orbit, also be subjected to the influence of other gravitation such as the bigger moon and the sun, the space mechanics circumstance complication.In the design process of transfer orbit, the designer can not simply adopt the disome model, and must adopt complicated trisome model even limbs model.So the long duration of action of complicated mechanical environment makes the transfer orbit shape become complicated space curve, and no longer is the simple plane ellipse, so its track can not simply be expressed by altitude of the apogee and perigee altitude.

Mission need at different deep space probes, the radiation pattern of Russia and U.S.'s employing multistep rocket level above different, for example Muscovite alliance rocket cooperates with level above the Fei Leigaite, once finished the emission of deep space probes such as Mars Express, its emission process had a detailed description in " European Space Agency carry out express detector " literary composition on 2006 the 3rd phases in Venus Express.the Maximum Power Point Tracker and the InitialCommissioning one literary composition neutralization " China Aerospace " people such as Ferdinando Tonicello; And the U.S. adopts fourstage rocket emissions such as Delta, Atlas, Titan IV/SRMU more, referring to David W.Dunham, people's such as people's such as Robert W " Libration PointMission, 1978-2002 ", Troy D.Goodson " Cassini Maneuver Experience:Launch and Early Cruise ".Above-mentioned rocket is a fourstage rocket, and wherein the fourth stage is the top level of rocket.After carrier rocket is sent detector into parking orbit, by top level igniting it is sent into transfer orbit again, be suitable for launching deep space probe.At present the carrier rocket of China does not also have top level, in emission process, separates after detector directly sent into track designation.By Yang Weilian, the Zhou Wenyan description in " Chang'e I moon probing satellite orbit Design " literary composition as can be known: in " Chang'e I " emission, carrier rocket is sent it into geostationary transfer orbit (GTO), utilize fuel enforcement change rail on the star afterwards, the goddess in the moon enters transfer orbit.

Deep space probe and higher moonlet or the load/higher large satellite of satellite total mass ratio of detection efficiency for farther distance in future, institute's carrying fuel quality on the star is restricted, need make full use of the emissive ability of China's carrier rocket, it is directly sent into transfer orbit.

Because transfer orbit is complex-shaped, and can not simply be described by track six roots of sensation number, therefore, need to seek the convenient and delivery side's coordinates transmission requirement of a kind of method as near-earth orbit.

Equivalent method is a method commonly used in science design, calculating, be widely used in a plurality of fields such as control, mechanics, Aero-Space, for example Zhai Guang, Yang Xiaoping once utilized the equivalent parameters method to analyze the dynamics of the moral honeycomb sandwich panel of aerospace field widespread use.Much relate to equivalent method can looking into also to have in the patent documentation, but in the orbit Design of deep space probe, also do not introduce " equivalence " notion.

In the actual engineering, whether the measurement detector succeeds in sending up and enters transfer orbit, only needs to weigh carrier rocket the energy that detector needs on this aspect on the transfer orbit is provided for detector at the injection point place, promptly has the correct speed of entering the orbit at correct injection position place.Consider that deep space probe institute is stressed still mainly from the earth near the earth, adopt near the disome kinetic model of the satellite of the earth still to have practical significance, therefore, adopting the equivalent parameters that obtains under this model to describe transfer orbit is a kind of feasible method, and the equivalent parameters method is easy to be widely accepted.

Summary of the invention

Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, a kind of method that adopts equivalent parameters to describe deep space detector equivalent transfer orbit is provided.

Technical solution of the present invention is: a kind of deep space detector equivalent transfer orbit is determined method, comprises the following steps:

(1), extracts the orbit parameter of designed transfer orbit according to dynamics of orbits modelling deep space probe transfer orbit;

(2) according to orbit parameter of extracting in the step (1) and space geometry relation, find the solution near the equivalent parameters of the described deep space probe transfer orbit earth.

Dynamics of orbits model in the described step (1) adopts disome model, trisome model, or many phantoms type.

Equivalent parameters in the described step (2) comprises orbital eccentricity e, semi-major axis of orbit a, orbit inclination i, right ascension of ascending node Ω, argument of perigee ω and epoch true anomaly θ constantly, and its computing formula is respectively:

Orbital eccentricity e:

$e=\left|\stackrel{\→}{e}\right|$

Semi-major axis of orbit a:

$a=\frac{{\mathrm{<figure-callout\; id="2"\; label="h"\; filenames="A20081022223100131.png"\; state="\{\{state\}\}">h</figure-callout>}}^2}{{\mathrm{\&mu;}}_e(1-e^2)}$ $h=\left|\stackrel{\&RightArrow;}{h}\right|$

Orbit inclination i is

With Between angle, and 0 °≤i＜180 °:

$i={\cos }^{-1}\left(\frac{h_K}{h}\right)$ $h=\left|\stackrel{\&RightArrow;}{h}\right|$

Right ascension of ascending node Ω is

With

Between angle:

$n=\left|\stackrel{\&RightArrow;}{n}\right|$

Argument of perigee ω is With

Between angle:

Epoch, true anomaly θ constantly was

With

Between angle:

Wherein, angular momentum vector

$\stackrel{\&RightArrow;}{h}=\stackrel{\&RightArrow;}{r}\&times;\stackrel{\&RightArrow;}{v}$

$=\left[\begin{array}{c}y\&CenterDot;v_z-z\&CenterDot;v_y\\ z\&CenterDot;v_x-x\&CenterDot;v_z\\ x\&CenterDot;v_y-y{\&CenterDot;v}_x\end{array}\right]$

Injection point place position of detector vector is $\stackrel{\&RightArrow;}{r}={\left[\begin{array}{ccc}x& y& z\end{array}\right]}^T,$ Velocity is $\stackrel{\&RightArrow;}{v}={\left[\begin{array}{ccc}{x}_x& v_y& v_z\end{array}\right]}^T;$ The ascending node vector

$\stackrel{\&RightArrow;}{n}=\stackrel{\&RightArrow;}{K}\&times;\stackrel{\&RightArrow;}{h}$

$=\left[\begin{array}{c}x\&CenterDot;v_z-z\&CenterDot;v_x\\ y\&CenterDot;v_z-z\&CenterDot;v_{\mathrm{<figure-callout\; id="0"\; label="y"\; filenames="A20081022223100131.png"\; state="\{\{state\}\}">y</figure-callout>}}\\ 0\end{array}\right]$

The excentricity vector

Its direction is pointed to the perigee of track by the earth's core:

$\stackrel{\&RightArrow;}{e}=\frac{1}{{\mathrm{\&mu;}}_e}[(v^2-\frac{{\mathrm{\&mu;}}_e}{r})\&CenterDot;\stackrel{\&RightArrow;}{r}-(\stackrel{\&RightArrow;}{r}\&CenterDot;\stackrel{\&RightArrow;}{v})\&CenterDot;\stackrel{\&RightArrow;}{v}];$

Direction vector for equator, the earth's core inertial coordinates system z axle;

Direction vector for equator, the earth's core inertial coordinates system x axle;

Direction vector for equator, the earth's core inertial coordinates system y axle.

The present invention compared with prior art beneficial effect is:

(1) the present invention utilizes equivalent parameters to represent near the earth transfer orbit of deep space probe, when carrier rocket is directly sent into transfer orbit with detector, delivery side needn't set up complicated deep space probe kinetic model, adopt existing model estimation launch accuracy just can more accurately estimate launch accuracy, solved the present carrier rocket of China and also do not had top level, the level igniting is sent detector into problem in the transfer orbit above can't adopting; And the present invention adopts the notion of equivalence to address this problem, and is simple and reliable.

(2) employing the inventive method can be by simple conversion Calculation, the transfer orbit of the complexity that deep space probe is formed in the deep space environment of complexity be converted to the earth near the similar simple elliptical orbit of satellite be described, it is simple to have method, easy to understand, physical significance is advantage clearly.

(3) transfer orbit described of equivalent parameters of the present invention is visual in image, and to the emission earth near the transfer orbit of aircraft similar, convenient and the coordination of delivery side.

(4) method is simple, is easy to promote.

Description of drawings

Fig. 1 equivalent parameters computing method process flow diagram;

Fig. 2 is near the orbital tracking synoptic diagram of the track earth;

Fig. 3 is the projection of transfer orbit in ecliptic plane in the embodiment of the invention.

Embodiment

A kind of deep space detector equivalent transfer orbit of the present invention is determined method, flow process as shown in Figure 1, concrete steps are as follows:

(1), extracts the orbit parameter of designed transfer orbit, the position and the velocity of particularly launching injection point according to dynamics of orbits modelling deep space probe transfer orbit;

The dynamics of orbits model can adopt disome model, trisome model, or many phantoms type, these models all are known to those skilled in the art, specifically can be referring to Shanghai space flight, 2003 the 4th phases, " moon exploration technology---an orbit Design and calculating " literary composition and PROGRESS IN ASTRONOMY, 2005 the 2nd phases, the relevant introduction in " being positioned at the relevant issues on the particular point in the solar system " literary composition about space probe.

In the survey of deep space activity, the task orbital distance earth of detector is remote, usually according to behind the detection mission design objective track, according to selected carrier rocket emissive ability, obtains the transfer orbit of detector by finding the solution the Lambert problem.The Lambert problem also is those skilled in the art's a common practise, can be referring to American Institude of Aeronautics and Astronautics publishing house content of the 276th～280 page in the 6th chapter of " An Introduction to The Mathematics and Methods ofASTRODYNAMICS " book of publishing in 1987 about the concrete introduction of Lambert problem.

(2) according to orbit parameter of extracting in the step (1) and space geometry relation (as shown in Figure 2), find the solution near the equivalent parameters of the described deep space probe transfer orbit earth.

Described equivalent parameters comprises orbital eccentricity e, semi-major axis of orbit a, orbit inclination i, right ascension of ascending node Ω, argument of perigee ω and epoch true anomaly θ constantly, and its computing formula is respectively:

Orbital eccentricity e:

$e=\left|\stackrel{\&RightArrow;}{e}\right|$

Semi-major axis of orbit a:

$a=\frac{{\mathrm{<figure-callout\; id="2"\; label="h"\; filenames="A20081022223100131.png"\; state="\{\{state\}\}">h</figure-callout>}}^2}{{\mathrm{\&mu;}}_e(1-e^2)}$ $h=\left|\stackrel{\&RightArrow;}{h}\right|$

Orbit inclination i is

With

Between angle, and 0 °≤i＜180 °:

$i={\cos }^{-1}\left(\frac{h_K}{h}\right)$ $h=\left|\stackrel{\&RightArrow;}{h}\right|$

Right ascension of ascending node Ω is

With

Between angle:

$n=\left|\stackrel{\&RightArrow;}{n}\right|$

Argument of perigee ω is

With

Between angle:

Epoch, true anomaly θ constantly was

With

Between angle:

Wherein, angular momentum vector

$\stackrel{\&RightArrow;}{h}=\stackrel{\&RightArrow;}{r}\&times;\stackrel{\&RightArrow;}{v}$

$=\left[\begin{array}{c}y{\&CenterDot;v}_z-z\&CenterDot;v_y\\ z\&CenterDot;v_x-x\&CenterDot;v_z\\ x\&CenterDot;v_y-y\&CenterDot;v_x\end{array}\right]$

Injection point place position of detector vector is $\stackrel{\&RightArrow;}{r}={\left[\begin{array}{ccc}x& y& z\end{array}\right]}^T,$ Velocity is $\stackrel{\&RightArrow;}{v}={\left[\begin{array}{ccc}{x}_x& v_y& v_z\end{array}\right]}^T;$ The ascending node vector

$\stackrel{\&RightArrow;}{n}=\stackrel{\&RightArrow;}{K}\&times;\stackrel{\&RightArrow;}{h}$

$=\left[\begin{array}{c}x\&CenterDot;v_z-z\&CenterDot;v_x\\ y\&CenterDot;v_z-z\&CenterDot;{\mathrm{<figure-callout\; id="0"\; label="v\; y"\; filenames="A20081022223100131.png"\; state="\{\{state\}\}">v</figure-callout>}}_y\\ 0\end{array}\right]$

The excentricity vector

Its direction is pointed to the perigee of track by the earth's core, integration two body motion equation, and arrangement obtains:

$\stackrel{\&RightArrow;}{e}=\frac{1}{{\mathrm{\&mu;}}_e}[(v^2-\frac{{\mathrm{\&mu;}}_e}{r})\&CenterDot;\stackrel{\&RightArrow;}{r}-(\stackrel{\&RightArrow;}{r}\&CenterDot;\stackrel{\&RightArrow;}{v})\&CenterDot;\stackrel{\&RightArrow;}{v}];$

Direction vector for equator, the earth's core inertial coordinates system z axle;

Direction vector for equator, the earth's core inertial coordinates system x axle;

Direction vector for equator, the earth's core inertial coordinates system y axle.

According to above-mentioned steps, can solve near the radical of the track of the earth shown in Figure 1.

Embodiment

" Kua Fu " plan is to survey and the international cooperative project of forecasting at solar-terrestrial physics weather, and wherein the task track of " Kua Fu A " star is the halo orbit of ordering around day ground L1, and this track and earth distance are not less than 1,270,000 kilometers.It is 200 kilometers that " Kua Fu A " star is directly sent into perigee altitude by rocket, and the apogee is 1,270,000 kilometers a transfer orbit." Kua Fu A " star not only is subjected to the terrestrial gravitation effect on transfer orbit, also be subjected to gravitation influences such as the bigger moon and the sun, the space mechanics circumstance complication, and its trade shape no longer is simple ellipse, as shown in Figure 3; Be the projection (reference coordinate be the earth's core ecliptic rotating coordinate system) of accurate transfer orbit in ecliptic plane shown in Fig. 3, this shows that this is the space curve of a strip complexity.

The injection point place of transfer orbit shown in Figure 3 detector's status be [6137.4801411748.967405 1594.532700-3.912119-8.524762-5.707643] (km, km/s); At the transfer orbit end, inject halo orbit point place detector's status for [332081.5119341103576.439016 601764.922985-0.524630 0.187722 0.054832] (km, km/s).

By above-mentioned data as can be known: at transfer orbit injection point place, detector is 200km apart from the height of floor level, relatively the speed 10.97968km/s in the earth's core; At the place, solstics of the liftoff ball distance of transfer orbit, about 1,290,000 kilometers of detectors, the speed 0.55990km/s in the earth's core relatively apart from the earth.Obviously, these 2 speed of locating do not meet the characteristics of motion on the elliptical orbit.Yet detector only needs to obtain corresponding speed at the injection point place, so under the effect of the non-coplanar force of complexity, and the detector transfer orbit shown in Figure 3 that just can fly out, wherein, the A point is represented injection point, and the B point is represented and is injected a little.

But, only is 200km at transfer orbit injection point place detector apart from the height on ground, the gravitation of the earth that detector mainly is subjected on this height can adopt the disome kinetic model to describe near the mechanical environment of this point, therefore can obtain near the equivalent parameters of the transfer orbit of the earth.

Utilize said method can find the solution the equivalent parameters that obtains transfer orbit:

Semi-major axis of orbit: a=624,057.5km

Orbital eccentricity: e=0.98946

Orbit inclination: i=35 °

Right ascension of ascending node: Ω=5 °

The argument of perigee: ω=155 °

Perigee altitude: h _p=200km

Altitude of the apogee: h _a=1,235,159km

What this group equivalent parameters was described is a highly elliptic orbit, and position, the velocity at its place, perigee are consistent with position, the velocity at actual transfer track injection point place.But under the space mechanics environment of reality, this track can not exist, and it is near the equivalent track of the actual transfer track earth.

The unspecified part of the present invention belongs to general knowledge as well known to those skilled in the art.

Claims (3)

1, a kind of deep space detector equivalent transfer orbit is determined method, it is characterized in that comprising the following steps:

(1), extracts the orbit parameter of designed transfer orbit according to dynamics of orbits modelling deep space probe transfer orbit;

(2) according to orbit parameter of extracting in the step (1) and space geometry relation, find the solution near the equivalent parameters of the described deep space probe transfer orbit earth.

2, a kind of deep space detector equivalent transfer orbit according to claim 1 is determined method, it is characterized in that: the dynamics of orbits model in the described step (1) adopts disome model, trisome model, or many phantoms type.

3, a kind of deep space detector equivalent transfer orbit according to claim 1 is determined method, it is characterized in that: the equivalent parameters in the described step (2) comprises orbital eccentricity e, semi-major axis of orbit a, orbit inclination i, right ascension of ascending node Ω, argument of perigee ω and epoch true anomaly θ constantly, and its computing formula is respectively:

Orbital eccentricity e:

$e=\left|\stackrel{\&RightArrow;}{e}\right|$

Semi-major axis of orbit a:

$a=\frac{h^2}{{\mathrm{\&mu;}}_e(1-e^2)}$ $h=\left|\stackrel{\&RightArrow;}{h}\right|$

Orbit inclination i is

With

Between angle, and 0 °≤i＜180 °:

$i={\cos }^{-1}\left(\frac{h_K}{h}\right)$ $h=\left|\stackrel{\&RightArrow;}{h}\right|$

Right ascension of ascending node Ω is the angle between I and the n:

$n=\left|\stackrel{\&RightArrow;}{n}\right|$

Argument of perigee ω is With

Between angle:

Epoch, true anomaly θ constantly was

With

Between angle:

Wherein, angular momentum vector

$\stackrel{\&RightArrow;}{h}=\stackrel{\&RightArrow;}{r}\&times;\stackrel{\&RightArrow;}{v}$

$=\left[\begin{array}{c}y\&CenterDot;v_z-z\&CenterDot;v_y\\ z\&CenterDot;v_x-x\&CenterDot;v_z\\ x\&CenterDot;v_y-y\&CenterDot;v_x\end{array}\right]$

Injection point place position of detector vector is $\stackrel{\&RightArrow;}{r}={\left[\begin{array}{ccc}x& y& z\end{array}\right]}^T,$ Velocity is $\stackrel{\&RightArrow;}{v}\text{=}{\left[\begin{array}{ccc}{x}_x& v_y& v_z\end{array}\right]}^T;$ The ascending node vector

$\stackrel{\&RightArrow;}{n}=\stackrel{\&RightArrow;}{K}\&times;\stackrel{\&RightArrow;}{h}$

$=\left[\begin{array}{c}x\&CenterDot;v_z-z\&CenterDot;v_x\\ y\&CenterDot;v_z-z\&CenterDot;v_y\\ 0\end{array}\right]$

The excentricity vector Its direction is pointed to the perigee of track by the earth's core:

$\stackrel{\&RightArrow;}{e}=\frac{1}{{\mathrm{\&mu;}}_e}[(v^2-\frac{{\mathrm{\&mu;}}_e}{r})\&CenterDot;\stackrel{\&RightArrow;}{r}-(\stackrel{\&RightArrow;}{r}\&CenterDot;\stackrel{\&RightArrow;}{v})\&CenterDot;\stackrel{\&RightArrow;}{v}];$

Direction vector for equator, the earth's core inertial coordinates system z axle;

I is the direction vector of equator, the earth's core inertial coordinates system x axle;

Direction vector for equator, the earth's core inertial coordinates system y axle.

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