FR2664562A1 - Satellite of the geostationary type with an apogee manoeuvring system having liquid ergols and several nozzles - Google Patents

Abstract

Satellite (1) intended to be stabilised in autorotation on a geostationary orbit including, coaxial with an autorotation axis (-Z+Z), a satellite body (2) surrounded by a cylindrical solar generator (3), an apogee manoeuvring system (8) located along the axis and two pieces of equipment (7, 6) located axially, characterised in that the apogee manoeuvring system includes, opposite one of the pieces of equipment, a plurality of at least two nozzles (12) pointing parallel to the axis, but offset with respect to the latter by one and the same distance (L), distributed circumferentially in an even fashion around the other one of the pieces of equipment, these nozzles being connected to one and the same liquid-ergol powered orientation system (20-30).

Description

 The invention relates to the general structure of satellites, for example observation satellites, particularly terrestrial satellites. It is particularly aimed at the case of satellites stabilized in autorotation; it concerns in particular, but not exclusively, meteorological satellites.

Examples of autorotation-stabilized geostationary meteorological satellites have already been implemented in the United States, the Soviet Union, Japan, and the United States.
Europe under the WWW (WORLD WEATHER WATCH) weather observation program. In this respect, the GOES 1 and 4 satellites for the United States,
MOP in Europe and GOIS in Japan.

 Earth observation satellites have been, until now, both stabilized in autorotation and equipped with a solid propellant apogee maneuvering system.

 At present, both in the United States and in Europe in particular, various projects or pre-projects are under way for the establishment of so-called second-generation satellites as opposed to those already existing qualified thus of first generation; these second-generation satellites aim at performances, notably in number of channels and in precision, notably superior to those of the first generation satellites; this is accompanied by a dry mass much higher than that of the first generation satellites (from 800 to 1500 kg against approximately 300 to 350 kg previously).

 Since the days when first-generation satellites were designed and launched, there has been a major shift in the field of satellite propulsion, whether geostationary or not: the implementation of solid propellant apogee engines has become obsolete. in favor of liquid propellant propulsion systems.

 This generalization of liquid propellants at the expense of solid propellants is explained by a better specific impulse, the possibility of providing a unified supply system for the apogee maneuvering system and for the orbit correction and control system. attitude, a better operational flexibility (possibility of a climax maneuver in several phases) and finally a better flexibility of concept (one fills the tanks in liquid propellants before the launch according to the real mass, whereas before there During the development phase, it was necessary to choose the solid propellant engine with its tank depending on the possible mass of the satellite once finished). In fact, the use of liquid propellants results overall in a large mass gain at launch.

 One consequence of this development is that in Europe there are currently no solid propellant apogee engines that are approved and adapted to the masses of modern satellites (800 to 1500 kg of dry mass).

 In addition, the implementation of liquid propellant apogee tuyères poses a specific critical problem in the case of observation satellites (including meteorological) stabilized autorotation.

 Indeed, the mission of such observation satellites involves taking images of the earth and / or its atmosphere in infrared channels. The detectors used for this purpose must be cooled to low temperatures, of the order of 100 K, so as to have acceptable signal / noise ratio characteristics. These low temperatures are conventionally obtained by placing the focal plane of the observation instrument (where the infrared detectors are located) under a frustoconical radiator, itself placed on a transverse surface of the satellite, opposite the stellar void. , in order to minimize the flow of energy (especially solar energy) coming from the outside and likely to hinder the evacuation of calories by this radiator.The frustoconical lateral wall of this radiator is conventionally inclined at an angle slightly greater than 23, 50 with respect to a plane transverse to the axis of autorotation and is polished carefully (it is most often aluminum) so as to have an excellent coefficient of specularity and to reject any solar rays outside the radiator incidents, even in the worst cases (especially, winter solstice when the radiator is on the SOUTH side).

 For all first generation satellites, the field of view of the passive radiator to the cold space (to the south) was obtained by ejection of the solid propellant apogee engine after its combustion; the radiator was located just behind this apogee engine, at the level of the apogee satellite / engine interface, the liquid propellant apogee engine having from this point of view the advantage of constituting with its reservoir a compact assembly easily ejectable.

 it is recalled that the opposite face of the satellite (NORTH side) is conventionally occupied by the communication antennas with the ground, which comprise in particular conventionally a telecommunication mast disposed precisely along the autorotation axis of the satellite.

 Such a possibility of ejection of the apogee maneuvering system seems excluded in the case of a liquid propellant system (especially in the case of a unified feed), given that the propellants intended for the propulsion of apogee are brought to the apogee nozzle (oriented precisely along the axis of autorotation) from tanks located inside the body of the satellite by pipes and that it is not realistic, for reasons of sealing, to provide a cutoff between the tanks and the apogee nozzle for the purpose of ejection thereof.

 The design of the geostationary observation satellites stabilized in autorotation thus supposes to solve the following technical problem: how to implant the antennas, arranged axially, the system of maneuver of apogee with liquid propellants, classically arranged axially, and the radiator of the system of observation, conventionally arranged axially, while respecting the constraints arising from the autorotation of the satellite (in particular, the observation system conventionally comprises moving optical parts that it appears essential to maintain as close as possible to the axis of autorotation to subtract them from possible crippling centrifugal efforts). More generally, the problem is to reconcile on a stabilized satellite in rotation, an axial propulsion system and two equipment (here an antenna tower and a radiator) to be arranged in the axis of rotation.

 The invention proposes for this purpose, a satellite intended to be stabilized in autorotation on a geostationary orbit comprising, coaxial with an autorotation axis, a satellite body surrounded by a cylindrical solar generator, an apogee maneuvering system arranged along the axis and two equipment arranged axially, characterized in that the apogee maneuvering system, fixed permanently opposite one of the equipment, comprises a plurality of at least two thrusters oriented parallel to the axis , but offset with respect to it from the same distance, distributed circumferentially in a regular manner around the other equipment, these nozzles being connected to the same liquid propellant supply system.

 It will be appreciated that, according to a characteristic of the invention, original in itself, it is allowed to use several nozzles which allows the implementation of previously approved nozzles despite the increase in the mass of satellites.

 According to preferred arrangements of the invention, possibly combined - said other equipment disposed between the nozzles is a set of communication antennas, - the equipment to which the climax maneuvering system is attached is a radiator belonging to an observation system with infrared detectors, - according to another alternative, the equipment opposite which is fixed the apogee maneuver system is the set of communication antennas, the equipment located between the nozzles can be the aforementioned radiator, - the nozzles are two in number, arranged symmetrically on either side of the autorotation axis (beyond two nozzles can be spoken of a substantially regular ring of nozzles), - the offset distance of the nozzles relative to the axis is at least half the radius of the cylindrical solar generator, - the offset distance of the nozzles relative to the axis is at least two x third of the radius of the cylindrical generator, - this ejectable cover is a cold source, heaters being arranged around the radiator and intended, during their implementation, to cause decontamination of the lower part of the satellite, - the radiator is covered by an ejectable cover, - the ejectable cover is connected to the satellite body, radially outside the radiator, by frangible links pyrotechnically controlled, - elastic energy storage elements are interposed axially between the ejection cover and the satellite body, adapted to cause the ejection of the cover, - these elastic elements are axially compressed spiral springs, - a heat shield is interposed between the set of antennas and the nozzles, - this heat shield comprises adapted hinged flaps to clear the field of view of the antennas towards the Earth after the apogee maneuver, - the system of supply Liquid propellant ion also feeds orbit correction and attitude control nozzles.

 It will be appreciated that such a concept with multiple apogee nozzles was, a priori, hampered by unacceptable reticence on the part of a person skilled in the art. In particular, the question arose as to what would be the behavior of the satellite during the combined thrusts of two or more nozzles in case of eventual desynchronization of these thrusts and / or in the event of differences in the offset at the nozzle axis, or even in case of amplitude differences of said thrusts. On the other hand, what was the extent of the changes to be made to the satellite's propellant feed system to enable it to reliably and synchronously supply two or more nozzles. Finally, the location of equipment (such as that radiator or antennas radiator near the outlet of the nozzles did not lead to excessive pollution (contamination and thermal effect) due to the operation of said nozzles.

 In fact, the invention is based on the gyroscopic rigidity of the satellite, because of its stabilization in autorotation even before apogee maneuvers, which rigidity appeared, after tests, quite sufficient, in the case of satellites of second generation, given the mass and rotational speed imposed on them, to maintain at quite acceptable levels the adverse effects due to the aforementioned construction or operating defects concerning the nozzles.

 Furthermore, the invention has recognized realistic methods for limiting, the pollution of the equipment because of the nozzles, in the case of a radiator by the temporary filling of the latter through an ejectable cover of simple structure being satisfied with a super-isolation of conventional type, or more generally by ejectable or articulated screens (flaps that fold over the nozzles after the apogee maneuver).

 It has been found that the use of multiple nozzles makes it possible to achieve almost all the mass gain obtained for other types of satellite due to the transition from a solid propellant propulsion to a liquid propellant propulsion and a single nozzle. .

Objects, features and advantages of the invention appear from the description which follows, given by way of non-limiting example, with reference to the accompanying drawings in which:
FIG. 1 is a diagrammatic axial view, partially cut away, of a satellite according to the invention, according to the arrow I of FIG. 2, in the volume available in a launcher of the Ariane type,
FIG. 2 is another axial view along arrow II of FIG. 1;
FIG. 3 is a schematic view in perspective,
FIG. 4 is an enlarged detail view of the lower part of FIG. 1,
FIG. 5 is a simplified hydraulic diagram of the liquid propellant circuit of the apogee maneuvering system, of orbit correction and of attitude control,
FIG. 6 is a schematic axial view with partial tearing of the optical observation system according to arrow VI of FIG. 7,
FIG. 7 is a similar view along arrow VII of FIG. 6,
FIG. 8 is an axial sectional view of the radiator of said observation optical system,
FIG. 9 is a schematic figure showing this radiator equipped with its ejectable cover,
FIG. 10 is a view similar to FIG. 1 but corresponding to a variant in which the nozzles are arranged around the tower of antennas, and
- Figures 11 and 12 are views similar to those of Figures 2 and 3 corresponding to this variant.

 Figures 1 to 3 show an observation satellite for autorotation stabilization around a -Z + Z axis in a geostationary orbit, generally designated 1.

 In FIG. 1, this satellite is represented in an envelope in phantom marked A, representing the volume assigned to it in a carrier and double launch structure of the ARIANE type.

 This satellite comprises mainly a body denoted 2 as a whole surrounded by a cylindrical envelope 3 covered with solar cells and constituting a solar generator, an infrared observation optical system noted 4, provided with an inlet baffle 5 disposed radially and a radiator 6 disposed on a transverse face SUD of the body, a set of telecommunication antennas and image transmission noted 7 as a whole and arranged projecting on a transverse face NORD, opposite to the transverse face having the radiator , and a climax propulsion system noted as a whole, and arranged axially as a whole.

 As is apparent from Figures 1 and 2, the satellite body 2, sometimes called service module, typically comprises a main platform 2A disposed transversely and a tube 2B disposed axially integral with the platform.

 Inside this tube 2B is arranged the observation system 4 shown in more detail in FIGS. 6 and 7, the baffle 5, not shown in FIG. 2, passing through this tube and being arranged opposite an opening 3A. arranged in the solar generator (see Figure 3).

 At its upper part, this satellite body 2 is closed by an upper platform 9 to which the antennas 7 are attached. These antennas, not forming part of the present invention, will not be further detailed here; they are of any known type suitable for carrying out the telecommunication and image transmission missions imposed on the satellite.

 Around the tube 2B is disposed the apogee propulsion system 8. In fact, this system comprises a unified liquid propellant supply circuit also connected to the attitude control and orbit correction nozzles of any known type. appropriate (shown only in Figure 5 as 35).

 This apogee propulsion system 8, shown in detail in FIG. 5, thus notably comprises propellant reservoirs 10 and 11 held in place in the satellite body by various stiffening and fixing crosspieces of any appropriate known type, some of which are shown in Figures 1, 2 and 4 under the reference 100.

 As is apparent in detail in FIG. 4, this apogee propulsion system 8 does not comprise a single nozzle disposed along the autorotation axis, as in the known autorotation stabilized observation satellites, but a plurality of identical nozzles 12 noted arranged at equal distances L of the axis and angularly distributed regularly around it. In the example considered, these nozzles 12 are two in number, arranged symmetrically with respect to the axis along a diameter, the distance L being at least half the radius R of the constituent cylindrical envelope of the solar generator 3 .

 These nozzles are substantially disposed at the same level axially as the radiator 6 of the observation system 4.

 In the example shown, these nozzles 12 are in fact slightly set back towards the inside of the satellite body. According to a variant not shown, the nozzles are further radially outwardly shifted (at least 2/3 of R), but less recessed with respect to the radiator than in FIG. 4, taking into account the available volume in the launcher.

 As is apparent from FIG. 5, the unified liquid propellant supply circuit is very close to the conventional unified supply circuit known in the case of a single apogee nozzle. This circuit comprises a pressurization stage 20 connected to a source of helium 21 and a supply circuit 22 or 23 specific to each of the two liquid propellants used, here monomethylhydrazine (abbreviated to MMH) in tanks noted 24 and 25 and nitrogen peroxide (N204) in reservoirs 26 and 27.

 Tests have shown that the passage of one to two tuyere (s) implied, for reliable operation, in addition to a simple duplication of the propellant feed lines 28 and 29 and the pressurization line 30, only a slight increasing a few percent of the diameter of the feed lines conventionally used for the known apogee tuyères, so as to compensate for the pressure drop resulting from the division of the propellant flow into two branches.

 It should be noted that the nozzles 12 are identical to conventional nozzles used alone in the known climax propulsion systems (in practice thrust nozzles equal to 400 N).

 Compared to a conventional single-nozzle system, the mass penalty due to the use of two nozzles instead of one is only about 3.1 kg (2.5 kg for the additional nozzle and about 0.6 kg for the increase of diameter of the lines), which is negligible compared to the gain in mass (which can reach 200 kg approximately) resulting from the passage of a propulsion by solid propellants to a propulsion by liquid propellants.

 This circuit of Figure 5 being very close to a conventional circuit, the aforementioned reservations, it will not be detailed further here. It is recalled that the valves marked with a P are pyrovalves, normally closed (+) or normally open (-); those marked with an I are Latch Valve lock valves; those marked M are manual valves normally closed before launch.

 The observation system 4 is conventional in itself and it is only as a reminder that it is presented in Figures 6 and 7.

 It mainly comprises a mirror 30 which reflects incident radiation S, parallel to the axis, on a P-shaped focusing plane in FIGS. 6 to 8 located in a housing 32 containing infrared detectors of any known type along a radiative surface. cross-section 33, for example covered with a suitable known white paint surrounded by the frustoconical radiator 6.

 During the entire launch phase and until after the end of the apogee maneuver, the radiator 6 is closed by an ejectable cover 40 shown schematically in FIG. 9. This cover comprises a thick central portion 40A centered in the radiator 6 and a thinner peripheral ring 40B temporarily connected to a securing ring 41 integral with the satellite body and disposed around the radiator. The connection of the lid to the securing crown is ensured by elements 41A and 40C respectively connected to this ring and to this lid secured to each other by pyrotechnic control pins schematized at 42 in FIG. 9.

 The ejection of the lid is provided by elastic members 43 interposed axially between this ejection cover and the peripheral ring and released during the pyrotechnic control of the aforementioned pins.

These elastic members here are axially compressed spiral springs arranged in close proximity to the elements secured by the pins.

 This lid is covered with a super-insulation (it can be conventionally a composite coating based on titanium sheets). This cover is used as a cold source for trapping the contamination resulting from the combustion of propellants for this purpose, the super-insulation layer is advantageously covered with a layer of white paint.

 Similarly the peripheral ring is covered with a super-insulation; it is for example a composite material comprising a titanium layer parallel to doubly aluminized Kapton layers, and kept spaced relative to each other by fiberglass spacers. Marked heaters 44 are arranged near the peripheral ring (these are for example simple heating resistors). These heaters have the function, after the end of the apogee maneuver, to cause a vaporization of the majority of the products deposited from the combustion of the propellants, the gases thus obtained being trapped by condensation by the cold source that constitutes the lid .

By way of example, in the case of a 1-tonne satellite dry weight and a unified MMH / N204 dual propellant propulsion system, these combustion products typically have the following analysis.
- H20: 90 mg
- N2: 125mg
- CO: 50 mg
- H2: 5mg
- C 2: 25mg
Most of these products of combustion, being trapped by the lid, is removed during the ejection thereof, typically 24 hours after the apogee maneuver, the heaters being implemented for example as soon as the end of this maneuver.

A typical mass specification for the system of Figure 9 is - super insulation: 750 g - heaters: 500 g - lid and springs: 2.500 g
Table 1 lists the results obtained for various cases of asymmetry of the structure or operation of the nozzles 12.

 For each of the cases considered, Table 1 indicates the amplitude of the considered defect, the autorotation speed considered, the maximum observed nutation, the speed increment loss and the resulting speed increment angle offset for the satellite.

 It should be noted that a difference of 5 mm between the distances to the axis of the nozzles results in minute disturbances.

 An angular difference of 0.50 on the thrust vector likewise leads to minute perturbations.

 In the case of both a difference of 5 mm between the distances to the axis of the two nozzles, an angular displacement of 10 of the thrust vector and a dispersion of 16 N on the thrust, nutation is obtained. maximum of 3.30, a speed increment loss of 2.2 m / sec and a speed increment angle offset obtained of 0.360, which corresponds to disturbances that are quite acceptable and can be corrected thereafter. It is recalled that the nominal speed increment is typically of the order of 1500 m / sec.

 The above results correspond to autorotation speeds of 15 rpm, which is well below the autorotation speeds provided for second-generation satellites, which can reach 100 rpm.

 It has even been considered the case where one of the nozzles would fail which would result in a sensitive lever arm: at 15 rpm, the disturbances resulting from such a failure become important but when the speed of rotation approaching the 50 rpm disturbances come back within allowable limits.

 It is clear from the tests in this Table 1 that the gyroscopic rigidity associated with the desired autorotation speeds is sufficient to maintain at an acceptable level the disturbances likely to occur as a result of the implementation of a plurality of no particular precautions are necessary to ensure the fitting and operating accuracy of these nozzles.

 FIGS. 10 to 12 correspond to a variant embodiment of the nozzles, in which the latter marked 12 'are located no longer around the equipment that constitutes the radiator, but around the equipment disposed on the opposite face, namely the tower of antennas.

 In these FIGS. 10 to 12, elements similar to those of FIGS. 1 to 3 have the same reference signs, with the addition of a "prime" index when these elements have differences, either by their overall structure or by their implantation.

 The protection of the equipment, here constituted by the tower 7, with respect to the jets apogee nozzles (thermal impact, pollution ...) is here obtained by a thermal screen of any known type appropriate, noted 40 'in its assembly, whose constituent elements are provided with a composite coating based on titanium, for example.

 These elements can be fixed or removable.

 As a preferred example, these elements are articulated flaps interposed between the nozzles and the antenna tower, adapted in deployed configuration (shown) to protect the tower (at a time when it is not useful) and, after the apogee maneuver, to fall back (transversely to the axis) against the satellite body, eventually closing the apogee nozzles whose role is completed and clearing the field of view of the antennas to the Earth.

 It goes without saying that the above description has been proposed by way of non-limiting example and that many variants can be proposed by those skilled in the art without departing from the scope of the invention.

TABLE I

SPEED <SEP> NUTATION <SEP> LOSS <SEP> SHIFT
<tb> SOURCE <SEP> OF <SEP> DISTURB <SEP> AMPLITUDE <SEP> OF AUTHORIZATION <SEP> MAXIMUM <SEP> OF INCREMENT <SEP> OF <SEP> ANGULAR
<tb> t / m <SEP> () <SEP> SPEED <SEP> m / s <SEP> INCREMENT <SEP> FROM
<tb> SPEED <SEP> ()
<tb> DIFFERENCE <SEP> FROM <SEP> OFFSET <SEP>
<tb> TUYERES <SEP> BY <SEP> REPORT <SEP> + <SEP> 5 <SEP> mm <SEP> 15 <SEP> 0.6 <SEP> 0.3 <SEP> 0.06
<tb> A <SEP> AXIS
<tb> DIFFERENCE <SEP> ANGULAR <SEP> FROM
<tb> PUSH <SEP> 0.5 <SEP> 15 <SEP> 0.2 <SEP> 0.2 <SEP> 0.02
<tb> FAILURE <SEP> OF A <SEP> TUYERE <SEP> 15 <SEP> 44 <SEP> 144 <SEP> 9.6
<tb> FAILURE <SEP> OF A <SEP> TUYERE <SEP> 50 <SEP> 5.2 <SEP> 1.7 <SEP> 0.78
<tb> EN <SEP> COMBINATION
<tb> - <SEP> difference <SEP> of <SEP> offset <SEP> + <SEP> 5 <SEP> mm
<tb> - <SEP> Angular <SEP> gap <SEP> 1 <SEP> 15 <SEP> 3.3 <SEP> 2.2 <SEP> 0.36
<tb> - <SEP> difference <SEP> of <SEP> thrust <SEP> 16 <SEP> N
<Tb>

Claims (16)

 A satellite (1, 1 ') for autorotative stabilization in a geostationary orbit having, coaxial with an autorotation axis (-Z + Z), a satellite body (2) surrounded by a cylindrical solar generator ( 3), an apogee maneuvering system (8, 8 ') arranged along the axis and two axially arranged equipment (7, 6), characterized in that the apogee maneuvering system has the opposite of the one of the equipment a plurality of at least two nozzles (12, 12 ') oriented parallel to the axis, but offset with respect thereto by the same distance (L), distributed circumferentially in a regular manner around the other equipment, these nozzles being connected to the same orientation system liquid propellant (20-30).  2. Satellite according to claim 1, characterized in that said other equipment disposed between the nozzles (12) is a radiator (6) belonging to an observation system with infrared detectors.  3. Satellite according to claim 1 or claim 2, characterized in that the equipment (7) opposite which is mounted apogee maneuvering system (8) is a set of communication antennas.  4. Satellite according to claim 1, characterized in that said other equipment disposed between the nozzles (12 ') is a set of communication antennas.  5. Satellite according to claim 4, characterized in that the equipment opposite which is mounted apogee maneuvering system (8 ') is a radiator (6) belonging to an observation system with infrared detectors.  6. Satellite according to any one of claims 1 to 5, characterized in that the nozzles are two in number, arranged symmetrically on either side of the autorotation axis.  7. Satellite according to any one of claims 1 to 6, characterized in that the offset distance of the nozzles relative to the axis is at least half the radius of the cylindrical generator.  8. Satellite according to claim 7, characterized in that the offset distance of the nozzles relative to the axis is at least two-thirds of the radius of the cylindrical generator.  9. Satellite according to any one of claims 1 to 8, characterized in that the radiator is covered by an ejectable cover (40).  10. Satellite according to claim 9, characterized in that this ejection cover is a cold source, heaters (44) being arranged around the radiator and intended, when implemented, to cause decontamination of the lower part of the satellite .  11. Satellite according to claim 9 or claim 10, characterized in that the ejectable cover (40) is connected to the satellite body, radially outside the radiator, by frangible links pyrotechnically controlled (40C, 40A, 42). ).  12. Satellite according to any one of claims 9 to 11, characterized in that elastic energy storage elements (43) are interposed axially between the ejection cover and the satellite body, adapted to cause the ejection of the cover ,  13. Satellite according to claim 12, characterized in that these elastic elements (43) are axially compressed spiral springs.  14. Satellite according to claim 4 or claim 5, characterized in that a heat shield (40 ') is interposed between the set of antennas and the nozzles.  15. Satellite according to claim 14, characterized in that this heat shield comprises articulated flaps adapted to clear the field of view of the antennas to the Earth after the apogee maneuver.  16. Satellite according to any one of claims 1 to 15, characterized in that the liquid propellant feed system also feeds orbit cores and attitude control nozzles.