CN114137993B - Ground fire short transfer orbit launching window searching method with deep space maneuver - Google Patents
Ground fire short transfer orbit launching window searching method with deep space maneuver Download PDFInfo
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Abstract
The invention provides a method for searching a firing window of a short transfer orbit of a ground fire with deep space maneuvering, which comprises the following steps: establishing a track optimization design model, converting the track optimization design problem into a multi-dimensional nonlinear programming problem, determining track optimization design parameters and optimization indexes, and giving out a constraint function during model optimization; acquiring initial values of track optimization design parameters by using a C3 energy contour diagram and a Lambert theory; and (3) substituting the initial values of the track optimization design parameters into the solving process of the multi-dimensional nonlinear programming problem, and after optimization convergence, extracting the key parameters of the track scheme design to complete the search of the launching track window. The method of the invention provides a deep space maneuver in the short transfer of the ground fire considering the carrying launching constraint, and an optimized method is adopted to search the window, so that the method can effectively adapt to the constraint of the carrier rocket on the detector, broaden the launching window of the Mars detector, and can be used for searching the launching window of the ground fire transfer orbit of the Mars detector.
Description
Technical Field
The invention belongs to the field of design of a deep space exploration interstellar transfer orbit, and particularly relates to a method for searching a firing window of a short earth-fire transfer orbit with deep space maneuver.
Background
There are various flight patterns from the earth to mars, including direct transfer, low energy transfer, low thrust transfer, etc. The most commonly used method for Mars detection is direct transfer, which is divided into short transfer and long transfer. The short transfer is that the detector reaches mars before reaching the far-day point of the elliptical orbit of the centroid, and the transfer time is short; the long transfer is that the detector reaches mars after reaching the far-day point of the elliptical orbit of the centroid, and the transfer time is long.
The first Mars detection task in China puts forward the requirement of adopting a short transfer scheme. However, short transfer solutions face engineering problems of high requirements on the conditions of the launch of the vehicle (e.g. launch direction of the vehicle, time to glide), narrow emission windows, low residual mass of the probe reaching the mars, etc. In order to solve the problems faced by the engineering, it is necessary to provide a method for searching the launching window of the short transfer orbit of the ground fire.
Disclosure of Invention
In order to overcome the defects in the prior art, the inventor of the invention carries out intensive research and provides a method for searching the launching window of the ground fire short transfer orbit with deep space maneuver.
The technical scheme provided by the invention is as follows:
a method for searching a firing window of a short transfer orbit of a ground fire with deep space maneuvering comprises the following steps:
s1, establishing an orbit optimization design model, converting the orbit optimization design problem into a multi-dimensional nonlinear programming problem, determining orbit optimization design parameters and optimization indexes, and providing a constraint function during model optimization;
wherein, the track optimization design parameters include: emission time t of detector L Ground fire orbit transfer time TOF and deep space maneuver time T DSM And hyperbolic overspeed V of the probe as it leaves the earth ∞ ；
The track optimization indexes are as follows: total propellant consumption m p Or the total speed increment size delta V is minimum, and the two are equivalent;
s2, acquiring initial values of the track optimization design parameters by using a C3 energy contour diagram and a Lambert theory;
and S3, substituting the initial values of the track optimization design parameters into a multi-dimensional nonlinear programming problem solving process, solving the multi-dimensional nonlinear programming problem in the step S1, and after optimization convergence, extracting key parameters of track scheme design to complete the search of the launching track window.
The method for searching the launching window of the short transfer orbit of the ground fire with the deep space maneuver, provided by the invention, has the following beneficial effects:
(1) According to the method for searching the ground fire short transfer orbit launching window with the deep space maneuver, one orbit maneuver is added in the ground fire transfer process, the plane and the speed of the detector flying around the sun orbit are changed, and compared with the scheme of the direct short transfer orbit, the scheme effectively makes up the defects of the existing carrying conditions in China under the condition that the ground fire transfer time is not remarkably increased, widens the launching window of the detector and increases the residual mass of the detector reaching mars;
(2) the method for searching the launching window of the ground fire short transfer track with the deep space maneuver adopts an optimization method to design the ground fire transfer track with the deep space maneuver, and ensures the optimality of the design result of the track transfer scheme of the detector.
Drawings
Fig. 1 is a schematic view of the ground fire short transfer track of the present invention.
Detailed Description
The features and advantages of the present invention will become more apparent and appreciated from the following detailed description of the invention.
The invention provides a method for searching a launching window of a ground fire short transfer orbit with a deep space maneuver, which is characterized in that the deep space maneuver is carried out once in the ground fire short transfer, and an optimized method is adopted for searching the window, so that the constraint of a detector on a carrier rocket can be effectively reduced, the launching window of a Mars detector is widened, the method can be used for searching the launching window of the ground fire short transfer orbit of the Mars detector, and the schematic diagram of the ground fire short transfer orbit is shown in figure 1. The method specifically comprises the following steps:
S1, establishing an orbit optimization design model, converting the orbit optimization design problem into a multi-dimensional nonlinear programming problem, determining orbit optimization design parameters and optimization indexes, and providing a constraint function during model optimization;
in the present invention, the track optimization design parameters include: emission time t of detector L Ground fire orbit transfer time TOF and deep space maneuver time T DSM And hyperbolic overspeed V of the probe as it leaves the earth ∞ (3 × 1 matrix).
In the invention, the purpose of track design is to maximize the residual mass or minimize the total speed increment after the detector is braked by fire, and the optimization indexes are as follows: total propellant consumption m p Or the total speed increment size av min, which are equivalent.
The total velocity increment size Δ V is calculated as follows:
(1) according to the emission time t of the detector L And the ground fire orbit transition time TOF is obtained to obtain the time t of the detector reaching the Mars A ；
(2) From t L And t A Obtaining the positions R of the earth and the mars at the emission time and the arrival time of the detector according to the ephemeris e ,R m And velocity V e ,V m ；
(3) According to the emission time t of the detector L Hyperbolic overspeed V when leaving the earth ∞ And the position R of the earth e And velocity V e The state of the emission moment of the detector (including the position R) can be determined 0 And velocity V 0 )；
Emitting the detector at the moment t through track forecast L State (including position R) 0 And velocity V 0 ) Forecasting deep space maneuvering time T DSM Obtaining the state of the detector before deep space maneuver, including the position R DSM0 Velocity V DSM0 ；
(4) By solving the time T of the detector maneuvering from deep space DSM Position R of DSM0 To the time of arrival of Mars A Position R of m The constructed Lambert problem can obtain the state (including the position R) of the detector after the deep space maneuver DSMt Velocity V DSMt ) And state of arrival at Mars (including position R) a Velocity V a )；
(5) Determining deep space maneuvering speed increment V DSM ＝V DSMt -V DSM0 (ii) a By reaching state R of Mars a 、V a And speed V of Mars m And calculating the near-fire braking speed increment V according to the orbit parameters after the acquisition of the detector Mars b ；
(6) Total speed increment Δ V ═ V DSM |+|V b |。
In the invention, a constraint function establishes a constraint relation between a variable and an optimization target, and the specific constraint function is as follows:
(1) emission time t of detector L Usually between given maximum and minimum ranges, i.e. t Lmin ≤t L ≤t Lmax ；
(2) The ground-fire orbit transit time TOF is generally between a given maximum and minimum range, i.e. TOF min ≤TOF≤TOF max ；
(3) Deep space maneuvering time T DSM Usually between a given maximum and minimum range, i.e. T DSMmin ≤T DSM ≤T DSMmax ；
(4) The magnitude of the deep space maneuver speed increment is typically limited to less than a given maximum value, i.e. | V DSM |≤V DSMmax ；
(5) The magnitude of the increase in the rate of the near fire braking is generally limited to less than a given maximum value, i.e. | V b |≤V bmax ；
(6) Hyperbolic overspeed V of probe when leaving earth ∞ Constraint V ∞ | 2 ≤C 3Lmax In which C is 3Lmax Related to carrying capacity, determined by carrying capacity;
(7) overspeed by hyperbola V ∞ Can calculate declination lambda of the time of leaving the earth 0 Satisfy lambda 0 ≤i Lmax Argument of near place omega 0 Satisfy omega min ≤ω 0 ≤ω max Wherein i Lmax And minimum and maximum argument ω of perigee min 、ω max Related to carrying capacity, is determined by carrying capacity.
In summary, the track design problem can be converted into a multidimensional nonlinear programming problem, which is described as follows:
optimizing parameters:
X＝[t L ,TOF,T DSM ,V ∞ ]
optimizing indexes:
J＝ΔV→min
constraint function:
t Lmin ≤t L ≤t Lmax
TOF min ≤TOF≤TOF max
T DSMmin ≤T DSM ≤T DSMmax
|V DSM |≤V DSMmax
|V b |≤V bmax
|V ∞ | 2 ≤C 3Lmax
λ 0 ≤i Lmax
ω min ≤ω 0 ≤ω max
the above-described multidimensional nonlinear programming problem can be solved by a sequential quadratic programming method (SQP).
And S2, acquiring initial values of the track optimization design parameters by using a C3 energy contour diagram and a Lambert theory.
Before the model is optimized by the SQP method, an initial value of an optimization variable needs to be given, and the initial value can obtain the optimal emission time t of a detector of a two-pulse orbit by utilizing a C3 energy contour diagram and a Lambert theory L Time of earth-fire orbit transition TOF, hyperbolic overspeed V at detector leaving earth ∞ (ii) a In addition, the deep space maneuver is assumed to be located at the middle moment of the flight time, so as to generate the initial value of the orbit optimization design parameter.
S3, substituting the initial values of the track optimization design parameters into the solving process of the multidimensional nonlinear programming problem, solving the multidimensional nonlinear programming problem in the step S1, and after optimization convergence, extracting the key parameters (t) of the track scheme design L 、TOF、T DSM 、V ∞ ) And finishing the search of the launching track window.
In the invention, in order to improve the calculation efficiency of the model, the input parameters involved in the optimization process of the model are normalized, and the quality, the length and the time unit (M, L, T) involved in the normalization process are as follows:
in the formula, m 0 Is the initial mass of the probe;
the revolution orbit radius of the earth is one astronomical unit; μ is the gravitational constant of day.
The invention has been described in detail with reference to specific embodiments and illustrative examples, but the description is not intended to be construed in a limiting sense. Those skilled in the art will appreciate that various equivalent substitutions, modifications or improvements may be made to the technical solution of the present invention and its embodiments without departing from the spirit and scope of the present invention, which fall within the scope of the present invention. The scope of the invention is defined by the appended claims.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Claims (4)
1. A method for searching a launching window of a short transfer orbit of a ground fire with deep space maneuvering is characterized by comprising the following steps:
s1, establishing an orbit optimization design model, converting the orbit optimization design problem into a multi-dimensional nonlinear programming problem, determining orbit optimization design parameters and optimization indexes, and providing a constraint function during model optimization;
wherein, the track optimization design parameters include: emission time t of detector L Ground fire orbit transfer time TOF and deep space maneuver time T DSM And hyperbolic overspeed V of the probe as it leaves the earth ∞ ；
The track optimization indexes are as follows: the total velocity increment size Δ ν is minimal;
the total velocity increment magnitude av is obtained by:
(1) according to the emission time t of the detector L And the ground fire orbit transition time TOF is obtained to obtain the time t of the detector reaching the Mars A ；
(2) From t L And t A Obtaining the positions R of the earth and the mars at the emission time and the arrival time of the detector according to the ephemeris e ,R m And velocity V e ,V m ；
(3) According to the emission time t of the detector L Hyperbolic overspeed V when leaving the earth ∞ And the position R of the earth e And velocity V e The state of the emission moment of the detector, including the position R, can be determined 0 And velocity V 0 ；
Emitting the detector at the moment t through track forecast L Predicting the state of the deep space maneuver time T DSM Obtaining the state of the detector before deep space maneuver, including the position R DSM0 Velocity V DSM0 ；
(4) By solving the time T of the detector maneuvering from deep space DSM Position R of DSM0 To the time of arrival of Mars A Position R of m The formed Lambert problem can obtain the situation that the detector is behind the maneuver time in deep spaceIncluding the position R DSMt Velocity V DSMt And a state of arrival at Mars, including position R a Velocity V a ；
(5) Determining deep space maneuvering speed increment V DSM ＝V DSMt -V DSM0 (ii) a By reaching state R of Mars a 、V a And speed V of Mars m And calculating the near-fire braking speed increment V according to the orbit parameters after the acquisition of the detector Mars b ；
(6) Total speed increment Δ V ═ V DSM |+|V b |；
S2, acquiring initial values of the track optimization design parameters by using a C3 energy contour diagram and a Lambert theory;
and S3, substituting the initial values of the track optimization design parameters into a multi-dimensional nonlinear programming problem solving process, solving the multi-dimensional nonlinear programming problem in the step S1, and after optimization convergence, extracting key parameters of track scheme design to complete the search of the launching track window.
2. The method for searching the launching window of the ground fire short transfer orbit with the deep space maneuver as recited in claim 1, wherein the constraint function comprises:
(1) Emission time t of detector L Usually between given maximum and minimum ranges, i.e. t Lmin ≤t L ≤t Lmax ；
(2) The ground-fire orbit transit time TOF is generally between a given maximum and minimum range, i.e. TOF min ≤TOF≤TOF max ；
(3) Deep space maneuvering time T DSM Usually between a given maximum and minimum range, i.e. T DSMmin ≤T DSM ≤T DSMmax ；
(4) The magnitude of the deep space maneuver speed increment is typically limited to less than a given maximum value, i.e. | V DSM |≤V DSMmax ；
(5) The magnitude of the increase in the rate of the near fire braking is generally limited to less than a given maximum value, i.e. | V b |≤V bmax ；
(6) The detector is away from the groundHyperbolic overspeed V in time of sphere ∞ Constraint | V ∞ | 2 ＜C 3Lmax In which C is 3Lmax Related to carrying capacity, determined by carrying capacity;
(7) overspeed by hyperbola V ∞ Can calculate declination lambda of the time when the earth leaves 0 Satisfy lambda 0 ≤i Lmax Argument of near place omega 0 Satisfy omega min ≤ω 0 ≤ω max Wherein i Lmax And minimum and maximum argument ω of perigee min 、ω max Related to carrying capacity, is determined by carrying capacity.
3. The method for searching the launching window of the ground fire short-transfer orbit with the deep space maneuver as claimed in claim 1, wherein in step S2, before the model is optimized by the SQP method, the initial value of the optimization variable is required to be given, and the initial value is used for obtaining the optimal launching time t of the detector of the two-pulse orbit by utilizing the C3 energy contour diagram and the Lambert theory L Time of earth-fire orbit transition TOF, hyperbolic overspeed V at detector leaving earth ∞ (ii) a And (4) assuming that the deep space maneuver is positioned at the middle moment of the flight time, and generating initial values of the track optimization design parameters.
4. The method for searching the launching window of the ground fire short transfer orbit with the deep space maneuver as claimed in claim 1, wherein the input parameters involved in the model optimization process are normalized, and the normalized parameters relate to the following mass, length and time unit:
in the formula, m 0 Is the initial mass of the probe;
the revolution orbit radius of the earth is one astronomical unit; μ is the gravitational constant of day.
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