DE102009036518B4 - Spacecraft and method of performing a landing thereof - Google Patents

Abstract

Spacecraft missile (1) with at least two thrusters (5, 6) producing a resulting thrust force in the direction of a thrust axis (7) of the spacecraft missile, the spacecraft missile (1) carrying a fuel and an oxidizer, with a) the Engines (5, 6) are pivotably articulated on a base body (2) with a degree of freedom of rotation for changing an angle (α) relative to the thrust axis (7), b) a control unit (16) is provided, c) at least one with the Control unit (16) in control connection is provided actuator (3, 3a) which is coupled to the engines (5, 6) such that actuation of the at least one actuator (3, 3a) by the control unit (16) ca) Swiveling the engines (5, 6) with a change in the assigned angles (α1, α2) and cb) results in a change in the resulting thrust force in the direction of the thrust axis (7), and d) a sensor (19) is provided which da) the distance of the space flight Flugk body (1) captured by a landing surface and db) coupled to the control unit (16) for controlling the resulting thrust by pivoting the engines.

Description

TECHNICAL FIELD OF THE INVENTION

The invention relates to an aerospace missile having two (or more) thrusters that produce a resultant thrust toward a thrust axis of the spacecraft, the spacecraft carrying both a fuel and an oxidizer. For example, the spacecraft is a rocket. Furthermore, the invention relates to a method for performing a landing operation of a spacecraft missile on a landing area, in particular a celestial body or the earth.

STATE OF THE ART

• – Bei Flüssigkeitsraketentriebwerken werden der Treibstoff und der Oxidator von der Rakete voneinander getrennt mitgeführt und mit (Turbo-)Pumpen oder Druckgas in das Triebwerk befördert. Als Treibstoffe einsetzbar sind insbesondere Kerosin, Hydrazin und seine Derivate oder verflüssigter Wasserstoff, während als Oxidator meistens flüssiger Sauerstoff, bei Hydrazin oder seinen Derivaten als Treibstoff jedoch Stickstofftetroxid verwendet wird. Die energiereichste Mischung für ein Flüssigkeitsraketentriebwerk ist mit flüssigem Wasserstoff als Treibstoff und flüssigem Sauerstoff als Oxidator gebildet. Derartige Flüssigtriebwerke sind ab- und anschaltbar, so dass auch kurze Schubimpulse erzeugt werden können. Eine Regelung der Schubkraft der Flüssigkeitsraketentriebwerke ist begrenzt und unter hohem Aufwand möglich, beispielsweise über eine geeignete Ansteuerung der erforderlichen Turbopumpen zur Förderung von Treibmittel und Oxidator.
• – Hingegen werden bei einem Feststoffraketentriebwerk sowohl der Treibstoff als auch der Oxidator als feste Stoffe gebunden mitgeführt. Feststoffraketen kommen unter Umständen völlig ohne bewegliche Teile, zusätzliche Vorrichtungen wie Pumpen oder Leitungen und ähnliches aus. Der Treibstoff und Oxidator sind einfacher und sicherer zu handhaben als flüssige oder gasförmige Treibstoffe, da diese nicht entweichen können. Die Schubentwicklung über die Brenndauer kann durch die sogenannte Abbrandcharakteristik a priori vorgegeben werden. Entsprechend kann die Beeinflussung der Brenndauer erfolgen. Eine Schubregelung während des Abbrands hingegen ist nicht möglich.

Missiles that move in the near-earth atmosphere usually carry only fuel, while a required for the combustion of the fuel and for generating the thrust oxidizer can be removed from the atmosphere. On the other hand, aerospace missiles, which may also move outside of the near-Earth atmosphere, require a particular design of the engines themselves, as well as carrying both the fuel and the oxidizer, since the oxidizer can not be removed from the atmosphere. Especially solid rocket engines and liquid rocket engines are known for this:

• - In liquid rocket engines, the fuel and the oxidizer are carried separately from the rocket and transported with (turbo) pumps or compressed gas into the engine. In particular, kerosene, hydrazine and its derivatives or liquefied hydrogen can be used as propellants, while as oxidizer mostly liquid oxygen, but with hydrazine or its derivatives as fuel, nitrogen tetroxide is used. The most energetic mixture for a liquid rocket engine is formed with liquid hydrogen as fuel and liquid oxygen as oxidizer. Such liquid engines are switched off and on, so that even short thrust pulses can be generated. A regulation of the thrust of the liquid rocket engines is limited and possible with great effort, for example via a suitable control of the required turbopumps to promote blowing agent and oxidizer.
• - On the other hand, in a solid rocket engine, both the fuel and the oxidizer are bound as solids. Solid rocket may come completely without moving parts, additional devices such as pumps or lines and the like. The fuel and oxidizer are easier and safer to handle than liquid or gaseous fuels because they can not escape. The thrust development over the burning time can be predetermined a priori by the so-called burn-up characteristic. Accordingly, the influencing of the burning time can take place. On the other hand, a thrust control during burning is not possible.

Finally, hybrid rocket engines are known that consist of a combination of liquid rocket engines and solid rocket engines.

DE 1 153 657 A discloses a multi-stage launcher which has a fixed main liquid engine and two more than 10 ° pivotable on all sides and independent of the main engine working independently fluid control thrusters with low thrust. The liquid main engine is to provide the drive power with short burning time and low gravitational losses in the vicinity of the earth, while the liquid control thrusters to ensure roll stabilization and to form a position control system. The fluid control thrusters are gimbaled with the design of the pivoting devices such that small pivoting and holding forces are applied. The adjustment allows a continuous adjustment of the liquid control engines.

DE 1 506 133 A discloses an engine for aircraft and spacecraft, in which two sub-engines can be pivoted together in the same pivoting direction to change the thrust direction. The required for such a pivoting number of actuators is reduced by a coupling of the pivoting movements via an actuating linkage.

US 6,227,494 B1 concerns a spacecraft with special landing gear on earth or a planet. This has a sensor assembly that senses signals caused by a "touch-down" of the spacecraft on Earth or the planet. In response, control of the operation of additional landing thrusters is provided.

WO 2009/094603 A2 discloses a method for performing a landing operation that requires control of axially and radially acting engines.

OBJECT OF THE INVENTION

The present invention has for its object to provide a spacecraft missile and a method for controlling the aerospace missile with altered, simplified and / or improved influencing possibilities for a change in a thrust force by the used engines is generated, in particular during a landing operation of the spacecraft.

SOLUTION

The object of the invention is achieved with a spacecraft missile according to the features of independent claim 1. Further embodiments of such a spacecraft missile arise in accordance with the dependent claims 2 to 5. A method for solving the problem underlying the invention results according to the features of independent claim 6. Further embodiments of such a method are defined in the dependent claims 7 and 8 ,

DESCRIPTION OF THE INVENTION

The present invention is directed to a particular missile, namely a missile, which is also or completely outside the near-Earth atmosphere can be used. Thus, the missile is not equipped with an "air-breathing" engine. Rather, engines are used in the missile according to the invention, which are not "air breathing" can be used. For this purpose, the aerospace missile carries with it both a fuel and an oxidizer. Furthermore, not only an engine is used in the missile. Rather, find at least two engines use that produce a resultant thrust in the direction of a thrust axis of the spacecraft missile. Here, the thrust axis coincide with the longitudinal axis of the spacecraft missile, be offset against this or (at least in sub-operating areas) with respect to this inclined for a change in attitude.

Furthermore, the present invention is based on the observation that, for example, for a soft landing of the spacecraft on a landing surface formed by a planet or satellite surface, the use of parachutes is required in principle. If a braking effect due to the parachute is insufficient or the use of a parachute is not possible for structural reasons or due to the lack of a sufficiently dense atmosphere, it is necessary that a thrust, in particular a braking thrust, must be adjusted during the landing process. A direct control of the engine is not possible for the special engines used here, technically very complex or possible only within predetermined limits.

Based on this observation, according to the invention engines are not rigidly articulated to a main body of the spacecraft missile. Rather, these are pivotably articulated with a degree of pivoting freedom, via which an angle α relative to the thrust axis, in particular with respect to the longitudinal axis, can be changed. A pivoting takes place in accordance with a control unit, which is in control connection with at least one actuator. A coupling of the at least one actuator and the pivotable engines is designed such that an actuation of the actuator or a plurality of actuators, which is effected by the control unit, a pivoting of the engines has the consequence, which leads to a change of the associated angle α. Here, the pivoting of at least two engines is correlated with each other such that a change in the resulting thrust force takes place in the direction of the thrust axis.

Such an embodiment is based on the finding that the resulting thrust force is dependent on the angle or the angles α. Thus, however, an influence on the resulting thrust, for example during a landing, not (only) via the control or specification of the thrust of the individual engines. Rather, an additional influence over the specification of the angle α is given. In extreme cases, no performance adjustment of the thrust of each engine can be done for a change in the resulting thrust and only a change in the angle α. While it is quite possible that the thrust axis is always oriented coaxially to the longitudinal axis of the spacecraft missile, it is also conceivable that the angle of the thrust axis to the longitudinal axis is variable, which can also be done via a change in the angle α of the individual engines. Thus, by specifying the angle α for an extended embodiment, both the resulting thrust force and the angle of the thrust axis relative to the longitudinal axis of the spacecraft missile can be influenced. In the context of the present invention, the control unit according to the invention or an operation described below as a control or control should also include a control unit or a regulation or a regulation.

It is possible that over the control unit a synchronization of multiple actuators takes place such that a synchronous pivoting of the engines takes place. In this case, a control can take place in which a control signal of the control unit correlates with a swivel angle of the engine. Also possible is the use of a scheme for which a swivel angle of the engine is detected by a sensor. According to a feedback signal of this sensor, the control unit can then bring about a desired angle α by acting on the actuator. For a further development according to the invention, however, is a Coupling device provided. This coupling device ensures a synchronous pivoting of several engines in mutually opposite pivot directions. When using such a coupling device is ensured even with an inaccurate control of the individual actuators that the relative position of multiple engines to each other within predetermined limits can be specified. It is also conceivable that by using the coupling device only one actuator is provided, over which a plurality of engines can be pivoted simultaneously.

The aforementioned measures according to the invention are used for a landing approach of a landing area. In order to be able to improve and / or automate such a landing approach, the invention proposes that a sensor be provided in the spacecraft missile. The sensor detects the distance of the aerospace missile from a landing area and can then make with increasing approach to the landing area an intervention in the resulting thrust, in particular generate a braking thrust force, so that the smoothest landing possible. For this purpose, the sensor is coupled to the control unit.

Quite conceivable is the pivoting of the engine in stages, which in an extreme case, only two angular positions are made possible for the engines. It will be appreciated that for an increase in the adaptability of the resultant thrust, it is advantageous to provide multiple stages for engine pivoting. For a stepless adjustment of the resulting thrust, however, the invention proposes that a continuous pivoting of the engines is possible. To enable a gradual or stepless pivoting of the engines, a suitable actuator must be used, which thus allows a stepless or stepped operation, and / or the pivoting degrees of freedom and any latches or detents for the pivoting of the engines must be continuously or gradually formed.

A further embodiment of the invention pays particular attention to the actuating and / or holding forces to be applied by the at least one actuator. While in principle such actuators must apply actuating and holding forces, which are dependent on the thrust of the associated engine, since the engines can generate a moment which is aligned to an undesirable pivoting of the engines from their desired pivot position, is for a further proposal of Invention arranged the articulation of the engines to the body upstream of the engines. Ideally, so that the thrust axis of the individual engine passes through the articulation or pivot axis, so that the required actuation forces of the actuator for a pivoting of the engine are not affected by the thrust of the engine. The reason for this is that the thrust of the engine passes through the Anlenkort.

In the event that at least one actuator is at least temporarily relieved of holding forces to secure a once achieved pivot angle of the engine, a locking or locking device is provided according to the invention. About the locking or locking device, which is preferably actuated automatically by the control unit, a pivot angle of an associated engine (or at the same time the pivot angle of several engines) can be fixed.

Another solution of the problem underlying the invention is given by a method for performing a landing of a spacecraft missile. In such a method, a determination is first made of a resulting thrust force required for the landing. The resulting thrust force is less than the sum of the amounts of engine thrust at least in partial phases of the landing process. Alternatively or cumulatively, the determination of a resulting transverse thrust force required for the landing can take place. This resulting shear force deviates from the sum of the amounts of engine thrust. If, for example, the distance from a landing area is known via a sensor, a distance-time setpoint signal can be predetermined for a desired landing process in the control unit. The required resulting thrust then correlates with the second derivative of this distance-time command signal. By specifying a corresponding resultant thrust force according to the invention, the said distance-time setpoint signal course can thus at least be approximated. About the design of the transverse thrust transverse misalignment of the spacecraft missile can be compensated so that the spacecraft missile sits as central as possible or at a desired location of the landing area. For this purpose, it is advantageous if a sensor is provided which can also detect such an offset.

In the next method step of the method according to the invention, a determination is made of changed pivoting angles α of the engines. In this case, the changed pivoting angles are determined in such a way that at least an approximation to the determined resultant thrust force and / or the determined transverse thrust force takes place.

If the changed swivel angle of the engines is known, the control unit performs a control of the at least one actuator such that the engines are pivoted to the determined changed pivoting angles α.

It is also quite possible, however, for the actuators to be actuated in accordance with a control signal course predefined a priori, so that the swivel angle α is changed in accordance with a predefined course, for example triggered by an approach of the spacecraft with a desired distance to the landing area. It is also possible that such a course of the swivel angle from a map, different characteristics or functional dependencies is derived in the control unit, wherein a dependency can be taken into account by a current speed of the spacecraft, the weight of the spacecraft, the distance of Spacecraft missile from landing area and the like.

In a further development of the method according to the invention, a distance from a landing surface is determined, taking into account a determination of the resulting thrust force and / or the resulting transverse thrust force and the changed pivot angle α.

While it is quite possible that in addition to a change in the swivel angle α influencing the power output, a particularly simple method for influencing the thrust and for controlling a landing process is given if a change in the resultant thrust and / or the resulting shear thrust by sole change the swivel angle, but without affecting the power output of the engines.

Advantageous developments of the invention will become apparent from the claims, the description and the drawings. The advantages of features and of combinations of several features mentioned in the introduction to the description are merely exemplary and can come into effect alternatively or cumulatively, without the advantages having to be achieved by embodiments according to the invention. Further features are the drawings - in particular the illustrated geometries and the relative dimensions of several components to each other and their relative arrangement and operative connection - refer. The combination of features of different embodiments of the invention or of features of different claims is also possible deviating from the chosen relationships of the claims and is hereby stimulated. This also applies to those features which are shown in separate drawings or are mentioned in their description. These features can also be combined with features of different claims. Likewise, in the claims listed features for further embodiments of the invention can be omitted.

BRIEF DESCRIPTION OF THE FIGURES

In the following the invention will be further explained and described with reference to preferred embodiments shown in the figures.

1 shows a schematic representation of components of a spacecraft with pivoting angles of engines, which have a maximum resulting thrust result.

2 shows a schematic representation accordingly 1 with increased pivot angles α of the engines, which lead to a reduced resulting thrust.

three shows the balance of power on two engines, which are pivoted by an angle α relative to a thrust axis or longitudinal axis of the spacecraft missile.

4 shows a schematic block diagram for the interaction of different components of a spacecraft missile according to the invention.

DESCRIPTION OF THE FIGURES

1 shows components of a spacecraft missile 1 , here is a basic body 2 with an actuator three and a coupling device 4 and engines 5 . 6 , In 1 a thrust axis 7-7 is indicated coaxial with or offset from the longitudinal axis of the spacecraft missile 1 can be oriented or can be oriented with respect to this with a constant or variable angle. The engines 5 . 6 own thrust axles 8th . 9 , The engines 5 . 6 are each at the same distance to the thrust axis 7 swiveling over storage units 10 . 11 stored, whose pivot axes in 1 run vertically to the drawing plane. The push axes 8th . 9 in this case run through the pivot axes of the bearing units 10 . 11 , The engines 5 . 6 are each with swivel angles α ₁ , α ₂ to the bearing units 10 . 11 pivotable. This pivoting is caused by an actuation of the actuator three , For the illustrated schematic embodiment, the coupling of the actuator takes place three through as striving 12 . 13 trained coupling device 4 , wherein an end portion of the struts 12 . 13 on the actuator three is hinged while another end of the struts 12 . 13 each on an engine 5 . 6 is articulated (linkages in 1 not shown). An axial displacement of the articulation points on the actuator three along the thrust axis 7 thus leads to a tensile or compressive force in the struts 12 . 13 that with introduction into the engines 5 . 6 leads to a swivel moment, which on the engines 5 . 6 in the direction of pivoting about by the bearing units 10 . 11 predetermined pivot axes leads. According to 1 are the push axes 8th . 9 parallel to the thrust axis 7 oriented. Accordingly, the resulting thrust of the engines results 5 . 6 from the sum of the amounts of thrust of each engine 5 . 6 ,

According to 2 is the actuator three actuated in such a way that increased angles α ₁ , α ₂ result. For the illustrated embodiment, the coupling device 4 designed such that the pivoting angles α ₁ , α _{2 are the} same size, so that they are also referred to below with the angle α. It should be noted, however, that according to 2 a pivoting of the engines 5 . 6 by the same amount α, but with opposite sense of direction. For example, the angle α is 10 °, 15 °, 20 ° or 25 °.

The ratios of the thrust forces are in the free body picture according to three shown. The engines 5 . 6 generate shear forces F _S , which are also denoted by the reference numerals 14 . 15 Marked are. As a result of the pivoting of the engines 5 . 6 with the swivel angles α with respect to the thrust axis 7 can the shear forces F _S decompose into a proportion F _S7 , which is parallel to the thrust axis 7 is oriented, as well as a proportion F _SQ , which is transverse to the thrust axis 7 oriented, wherein the transverse shear forces F _{SQ of} the two shear forces 14 . 15 are oriented opposite to each other. For the amounts of the force components, the following relationship applies: F _S7 = F _S · cos α; F _SQ = F _S · sin α.

While the transverse shear forces F _SQ cancel each other out for equal angles α ₁ = α ₂ = α, a resultant thrust force F _RES results, which results from the sum of the magnitudes of the force components F _S7 , so that the following applies: F _RES = 2 × F _S × cos α

From the above formula for the determination of the resulting thrust force F _RES follows directly that an influence on the resulting thrust force F _{RES can be} brought about by influencing the angle α ₁ , α ₂ . Is a pivoting of the engines 5 . 6 not at the same angle α ₁ , α ₂ , but at different angles, so can also change the orientation of the thrust axis 7 opposite a longitudinal axis of the spacecraft missile 1 by the pivoting of the engines 6 . 5 be initiated.

The orientation and installation direction of the engines with respect to the spacecraft missile 1 can be arbitrary. For example, the engines can 5 . 6 in the 1 to three to be located down in a landing approach while the thrust forces 14 . 15 are formed as braking forces, thus generating a resultant thrust, namely a braking force in three oriented upwards. It is understood that the inventive idea is not based on the use of two pivotable engines 5 . 6 is limited. For example, it is possible for three, four, five ... engines to be uniform in the circumferential direction about a longitudinal axis of the spacecraft missile 1 are distributed and can be pivoted apart with swivel angles α, so that the associated thrust axes of the engines lie on the lateral surface of a cone whose opening angle corresponds to the angle α. For example, for three engines used in the context of the invention, a "pivoting of the engines in opposite directions" to be understood such that the engines of the thrust or longitudinal axis 7 be pivoted away with an increase in the opening angle of the cone on which lie the thrust axes of the engines.

According to 4 is in the spacecraft missile 1 a control unit 16 provided, via suitable control signals 17 . 18 actuators three . 3a actuated, their influence on the tilt angle of the associated engines 5 . 6 via the coupling device 4 coupled together. The control unit 16 is preferably the signal of a sensor 19 supplied, the the distance of the spacecraft missile 1 captured from a landing area. Furthermore, the control unit 16 Signals of any other device 20 to process. For example, in the facility 20 a map, a functional dependency or the like may be stored, to which the control unit 16 recourse. Also possible is that in the facility 20 a mass of the spacecraft is detected or approximated, for example by determining the duration of flight, which correlates with the reduction of the propellant and the oxidizer. Likewise, the facility 20 determine a current speed of the spacecraft missile and the control unit 16 provide. A consideration of any other operating or environmental parameters is possible accordingly.

LIST OF REFERENCE NUMBERS

1

Space Missile

2

body

3

actuator

4

coupling device

5

engine

6

engine

7

thrust axis

8th

thrust axis

9

thrust axis

10

storage unit

11

storage unit

12

strut

13

strut

14

thrust

15

thrust

16

control unit

17

control signal

18

control signal

19

sensor

20

Facility

Claims (8)

Spacecraft ( 1 ) with at least two a resultant thrust force in the direction of a thrust axis ( 7 ) spacecraft-producing engines ( 5 . 6 ), the spacecraft missile ( 1 ) carries a fuel and an oxidizer, wherein a) the engines ( 5 . 6 ) pivotable on a base body ( 2 ) are articulated with a degree of pivoting freedom for changing an angle (α) relative to the thrust axis ( 7 ), b) a control unit ( 16 ) is provided, c) at least one with the control unit ( 16 ) in control connection actuator ( three . 3a ) provided in such a way with the engines ( 5 . 6 ), that an actuation of the at least one actuator ( three . 3a ) by the control unit ( 16 ) ca) a pivoting of the engines ( 5 . 6 ) with a change in the associated angles (α ₁ , α ₂ ) and cb) a change in the resultant thrust force in the direction of the thrust axis ( 7 ) and d) a sensor ( 19 ), the da) the distance of the spacecraft missile ( 1 ) from a landing area and db) with the control unit ( 16 ) is coupled to control the resulting thrust by pivoting the engines. Spacecraft ( 1 ) according to claim 1, characterized in that a coupling device ( 4 ) is provided, the synchronous pivoting of multiple engines ( 5 . 6 ) ensured in mutually opposite directions of pivoting. Spacecraft ( 1 ) according to one of the preceding claims, characterized in that the pivoting of the engines ( 5 . 6 ) is infinitely possible. Spacecraft ( 1 ) according to one of the preceding claims, characterized in that the articulation of the engines ( 5 . 6 ) on the base body ( 2 ) upstream of the engines ( 5 . 6 ) is arranged. Spacecraft ( 1 ) according to one of the preceding claims, characterized in that a latching or locking device is provided, via which a pivoting angle of an engine ( 5 . 6 ) is determinable. Method for performing a landing operation of a spacecraft ( 1 ) according to one of the preceding claims on a landing surface, in particular a celestial body or the earth, characterized by the following method steps: a) determining aa) a resulting thrust force required for the landing, the resulting thrust deviating from the sum of the amounts of the partial thrust forces ( 14 . 15 ) of the engines ( 5 . 6 ), and / or ab) a resulting transverse thrust force required for the landing, the resulting lateral thrust differing from the sum of the magnitudes of the partial thrust forces of the engines; b) Determination of changed pivoting angles (α) of the engines ( 5 . 6 ), for which at least an approximation to the determined resultant thrust force and / or the determined transverse thrust force takes place; c) driving the at least one actuator ( three . 3a ) such that the engines ( 5 . 6 ) are pivoted to the determined changed pivot angles (α). A method according to claim 6, characterized in that a distance from a landing surface is determined, taking into account a determination of the resulting thrust force and / or the resulting transverse thrust force and the changed pivot angle (α). A method according to claim 6 or 7, characterized in that a change in the resulting thrust force and / or the resulting transverse thrust force by solely changing the pivoting angle, but without affecting the power output of the engines ( 5 . 6 ) he follows.

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