RU2289533C1 - Method of injection of spacecraft into inter-planetary trajectory - Google Patents

Abstract

FIELD: spacecraft inter-planetary flights with the aid of cruise jet engines, mainly electrical rocket engines.

SUBSTANCE: proposed method includes injection of spacecraft into heliocentric trajectory at distance of spacecraft from Sun followed by its approach to Sun. Active motion of spacecraft is realized in part of this trajectory behind Earth's orbit during operation of jet engines. Then, spacecraft returns to Earth at velocity increment and increases its heliocentric velocity in the course of gravitational maneuver near Earth. After spacecraft has crossed Earth's orbit in section of its approach to Sun and before entry into Earth's gravisphere, spacecraft is accelerated by repeated switching-on of cruise jet engines. At the moment of fly-by over Earth when gravitational maneuver is performed, angular motion of Earth and spacecraft relative to Sun are equalized. During fly-by over Earth, spacecraft is subjected to its gravitational field changing the vector of spacecraft heliocentric velocity, thus ensuring further acceleration of spacecraft and forming inter-planetary trajectory of flight to target.

EFFECT: reduction of time required for organization of gravitational maneuver in Earth's gravity field at injection of spacecraft into required inter-planetary trajectory.

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Description

The invention relates to rocket and space technology, and in particular to methods for launching a spacecraft (SC) on a flight path to celestial bodies.

Direct flights to distant celestial bodies during acceleration from near-Earth orbit require high speeds of departure, very significant expenditures of the mass of the device and are distinguished by the long duration of their achievement, comparable with the life of a person. These circumstances force us to look for other ways of their implementation. One of such ways is the use of gravitational fields of planets, in particular, directly the gravitational field of the Earth.

Existing analogues of the method of launching onto the interplanetary flight path for organizing the Earth’s flight suggest the use of additional gravitational maneuvers in the gravitational fields of the intermediary planets [1, p. 405, Fig. 151]. Thus, the United States, when implementing the Galileo and Cassini projects, used the gravitational impact on the flight path of the planet Venus to organize gravity maneuvers in the Earth's gravitational field. The disadvantages of this method are also associated with the long duration of its implementation (the flight from Earth to Earth lasts about two years), the implementation of this method is not possible every year, so it requires a certain relative position of the three celestial bodies of the solar system: Earth, Venus and the celestial body of the target of flight, as well as the difficulties associated with managing multiple spans by the apparatus of celestial bodies.

Closest to the proposed method of launch is the method of launching the spacecraft on the interplanetary flight path [1, p. 406 - prototype, Fig. 152], including its output to the heliocentric path, return to Earth with a speed increment and acceleration by the Earth's gravitational field. This method involves the launch of the spacecraft into a heliocentric orbit with an apogelium located beyond the orbit of the planet Mars, where it is braked to arrange a meeting with the Earth and to obtain an increase in the speed of the spacecraft during its passage.

The disadvantage of this method is the long duration of the spacecraft flight from Earth to Earth, about 2-3 years, which directly increases the flight time to the target.

The aim of the invention is to reduce the time of organization of the gravitational maneuver in the Earth's gravitational field during the launch of the spacecraft on the interplanetary flight path to the target.

The purpose of the invention is achieved by the fact that the method of launching a spacecraft onto an interplanetary flight path, including its output to the heliocentric path with the removal of the spacecraft from the Sun and its subsequent approach to the Sun, return to the Earth with a speed increment and increase in the heliocentric speed of the spacecraft as a result of gravitational maneuver Earth, that after the spacecraft crosses the Earth’s orbit in the area of its approach to the Sun and before the spacecraft enters the Earth’s gravisphere, they accelerate a apparatus using marching jet engines, ensuring equality of the angular movements of the Earth and the spacecraft relative to the Sun at the time of flight of the Earth.

The essence of the proposed method is illustrated in figure 1, which shows the heliocentric portion of the trajectory of the Earth-to-Earth flight.

Figure 2 for this trajectory presents the law of control of the thrust vector on the flight segment of the Earth-Earth.

Before carrying out the launch of the vehicle on the flight path to the celestial body for a specific composition and given values of the design parameters of the spacecraft, determine its flight program, ensuring equality of the angular movements of the Earth and the spacecraft at the time of its passage; determine the start date; Earth’s altitude and speed; moment of arrival to the target; the time of the impact on the spacecraft of a reactive force and its orientation in space. At the time of the passage of the Earth, the angular ranges of the spacecraft and the Earth can differ only by an integer number of revolutions relative to the Sun. In this case, trajectories with the same number of revolutions of the spacecraft and the earth relative to the sun are considered. To do this, they solve the optimization problem by finding the law of control of the jet thrust vector of marching engines that satisfies the required conditions. The mathematical apparatus used to determine the jet thrust vector control law, flight pattern parameters, and calculation algorithms are given in [2, 3].

Consider the sequence of operations of the proposed method of launching on an interplanetary trajectory. First, the spacecraft from the Earth’s surface is put into near-earth orbit. Next, the apparatus is affected by reactive force, a speed impulse is given to it, and it is brought to the heliocentric flight path. In the process of flying in the gravitational field of the Sun, in accordance with a given control law, the spacecraft is affected by reactive force and the trajectory of its return to the Earth’s gravisphere with increasing velocity is formed. During a flight in the Earth’s sphere of influence, the Earth’s gravitational field acts on the spacecraft, due to which the heliocentric velocity of the spacecraft increases and an interplanetary flight path to the target is formed.

Consider a possible device for a space rocket to implement this launch method. It may include a spacecraft equipped with the necessary systems for flight, including a propulsion system using chemical fuel or a propulsion system with an electric rocket engine, an upper stage, a multi-stage launch vehicle.

We give a specific example of the launch of the spacecraft on the flight path to the planet Jupiter and some of its characterizing figures. With the help of a rocket carrier, an initial mass of 8120 kg, consisting of a chemical booster block (SC) and a spacecraft with a solar electric propulsion system (specific thrust of 1800 s, nominal reactive thrust of 34.4 g), is launched into low Earth orbit. The RB starts on July 30, 2011 from near-Earth orbit and carries out the launch of the spacecraft from the Earth’s gravitational field to the heliocentric flight path with a speed of departure from the Earth of ≈1.74 km / s and is separated. After separation of the RB, the mass of the apparatus is ≈2230 kg. In this case, the heliocentric flight path lasting 400 days (see figure 1) consists of three active and three passive sections. The x axis indicates the direction of the vernal equinox. The orbit of the Earth is depicted by a thin line, active sites are represented by bold lines, passive - by dots. The symbol α indicates the apogelium of the earth's orbit. The radii indicate the positions of the spacecraft at the time of launch from the near-Earth orbit T _c and at the time of its entry into the Earth's gravisphere T _s , which coincide with the positions of the Earth in its orbit at the corresponding moments.

In the heliocentric portion of the flight, the motion control is carried out using marching engines in accordance with the specified law of reactive thrust vector control (see figure 2) and the spacecraft returns to the Earth’s gravitosphere with a given velocity increment, for this trajectory the reactive thrust vector control law is presented, and the notation is introduced: P is the value of jet thrust; α is the angle between the thrust vector and the normal to the instantaneous plane of the flight path; β is the angle between the projection of the thrust vector onto the instantaneous plane of the flight path of the vehicle and the radius vector of the Sun-KA; Δ - engine on-off function (if Δ> 0, then the engine is on; if Δ <0, then the engine is off).

At the beginning of the heliocentric section, the spacecraft continues to accelerate, then the passive flight section follows, and then the flight brake section. In this area, the apparatus continues to move away from the Sun, but, gradually losing speed, it begins to approach the Sun and crosses the Earth’s orbit after turning off the engine.

Then the engines turn on again to accelerate, the device gradually begins to approach the Earth and enters its gravisphere, flying over the Earth’s surface at an altitude of 800 km at a speed of 9.1 km / s. At this point in time, the mass of the apparatus is approximately ≈1800 kg. As a result of gravity maneuver in the Earth's gravitational field, the value of the heliocentric velocity of the vehicle increases by 5.57 km / s. 1020 days after the gravitational maneuver in the Earth's gravitational field, the spacecraft enters the gravitosphere of Jupiter. The flight time, starting from the start from Earth orbit to the entrance to the Jovian gravisphere, is approximately equal to 1420 days, and the mass of the device at the entrance to the Jovian gravisphere is ≈1580 kg. The use of solar electric rocket engines as part of the apparatus allows them to be used on the Earth-Jupiter flight section.

The proposed method of launching the spacecraft on an interplanetary trajectory during a flight to Jupiter allows reducing the time to reach Jupiter by two years compared to the method implemented by the USA in the Galileo project (start on October 18, 1989, meeting with Jupiter on December 7, 1995). In this project, one maneuver in the gravitational field of Venus and two maneuvers in the gravitational field of the Earth were carried out to bring the spacecraft to the flight path to Jupiter. It is not entirely correct to compare the mass characteristics of both methods of launching the spacecraft onto the interplanetary trajectory, as different booster blocks and spacecraft of different composition were used. However, we note that the masses of the devices at the entrance to the Jovian gravisphere are only 860 kg different, although in the Galileo project, a vehicle with an initial mass of more than 17 tons was launched from near-Earth orbit.

The use of marching engines for organizing gravity maneuvers near the Earth significantly expands the possibilities of providing the necessary increment of the Earth’s flight speed at a given time of its flight. The use of electric rocket engines as marching engines with a large specific impulse can significantly reduce the mass consumption of the device, and the small spacecraft removal from the Earth’s orbit is very favorable for using solar energy as a source of their power. An important advantage is the annual repeatability of this flight scheme, and the ability to control the time of the spacecraft return to Earth and its flight speed can significantly expand the spacecraft launch windows from near-Earth orbit.

The technical result is an increase in the efficiency of flights to distant celestial bodies. It is characterized by the following points: shortening the flight time to the target, increasing the spacecraft’s transport capabilities, the possibility of implementing the proposed flight scheme annually, cheapening and simplifying the organizational and technical support for flights to distant celestial bodies with ground-based controls, and the universality of the sequence of operations when flying to a particular celestial body.

LITERATURE

1. V.I. Levantovsky. Elementary Space Flight Mechanics, 3rd edition, revised and revised. - M .: Nauka, 1980, 512 p.

2. Fedotov G.G. Optimization of flights between the orbits of artificial satellites of two planets using a combination of high and low thrust. Space Research, 2002, Volume 40, No. 6, pp. 616-625.

3. Fedotov G.G. Optimization of flight paths of spacecraft with electric propulsion when using gravitational maneuver. Space Research, 2004, Volume 42, No. 4, pp. 404-413.

Claims (1)

The method of launching a spacecraft onto an interplanetary flight path, including its introduction to the heliocentric path with the spacecraft moving away from the Sun and its subsequent approach to the Sun, returning to the Earth with a speed increment and increasing the spacecraft heliocentric speed as a result of the Earth’s gravitational maneuver, characterized in that after the spacecraft crosses the Earth’s orbit in the area of its approach to the Sun and before the spacecraft enters the Earth’s gravisphere, the vehicle is accelerated using rshevye jet engines, ensuring equality of angular motions of the Earth and the spacecraft relative to the Sun at the time of passage of the Earth.

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