EP1140629A1

EP1140629A1 - Device for supporting a payload in a launch vehicle of said payload

Info

Publication number: EP1140629A1
Application number: EP99964727A
Authority: EP
Inventors: Pascal Hubert
Original assignee: Centre National dEtudes Spatiales CNES
Current assignee: Centre National dEtudes Spatiales CNES
Priority date: 1998-12-29
Filing date: 1999-12-28
Publication date: 2001-10-10
Also published as: JP2002533269A; US6533221B1; FR2787765A1; WO2000038988A1; FR2787765B1

Abstract

The invention concerns a launch vehicle comprising a shroud (1a, 5a) supporting auxiliary propellers (2, 3) applying in operation excess flow of stresses to the shroud (1a, 5a), said device comprising means for coupling the payload (4) with the launch vehicle shroud. The invention is characterised in that the coupling means are designed to ensure a mechanical coupling between the payload (4) and the shroud along the direction of the launch vehicle longitudinal axis (X), and to cause them to be disengaged along radial and tangential directions. The invention is useful for neutralizing the effects on the payload of excess stresses applied on the launch vehicle shroud by the auxiliary propellers (2, 3).

Description

DEVICE FOR SUPPORTING A USEFUL LOAD IN A LAUNCHER OF SAID LOAD

The present invention relates to a device for supporting a payload in a launcher of said payload, said launcher comprising a casing supporting at least one auxiliary propellant applying in operation surges of force to said casing, said device comprising mechanical connection means of said payload to said casing of the launcher.

Such a device is known from the international patent application WO 98/32658 filed by the applicant, this device being designed to be integrated into a launcher of a payload in space, schematically represented in Figure 1 of the accompanying drawing where the launcher comprises a first stage 1 fitted with auxiliary thrusters' 2 and 3 fixed in diametrically opposite positions on the casing 1a of this stage, parallel to the longitudinal axis X thereof. FIG. 1 also shows a payload such as a satellite 4 mounted on the casing 5a of a second stage 5 of the launcher, by means of an adapter in the form of a frustoconical skirt 6, the satellite being protected conventionally by a cap 7 during the passage through the atmosphere. The envelopes 1a and 5a are cylindrical, coextensive and integral with one another.

As explained in the aforementioned patent application, the thrusts developed by the auxiliary thrusters 2,3 apply to the envelope of the first stage intense axial forces, in particular at the attachment points 8, 9 of the thrusters 2, 3 on the first floor. Is shown in Figure 2 of the accompanying drawing the intensity F of these forces observed at the periphery of the casing la. The forces are much more pronounced at the attachment points 8.9 (angular position 90 ° and 270 ° respectively) and their transmission to the skirt 6, through the envelope 5a of the second stage 5, would cause asymmetric deformation of it, damaging to the satellite, in the absence of countermeasures. There is shown in Figure 3 of the accompanying drawing the cross section of the "deformed" of the casing of the launcher, subject to the distribution of forces shown in Figure 2, by auxiliary thrusters 2,3 fixed in 8,9 respectively on this envelope.

According to the aforementioned patent application, the transmission of deformations to the satellite is prevented by installing a suspension device constituted by an annular chamber with flexible wall, filled with a fluid, between the satellite and the launcher, for example at the level of the large base 6b of the skirt 6, which receives the satellite at its other base 6a. The "surges" of forces received by the annular chamber in line with the attachment points 8, 9 are uniformly distributed annularly by the fluid in the chamber, the skirt 6 then undergoing perfectly symmetrical forces which do not deform this skirt as tightly . The integrity of the satellite is thus safeguarded during the operation of the auxiliary thrusters 2,3.

Means are provided, moreover, for varying the pressure of the fluid contained in the annular chamber. Thus this pressure can be lowered by a relatively high value ensuring good mechanical cohesion of the launcher and the skirt 6, necessary in particular during the operation of the auxiliary thrusters 2,3, to a lower value making it possible to filter out forces such as vibrations or shocks propagating in the launcher, in particular during the separations of the stages of the launcher and the cap 7.

The device described in the aforementioned patent application makes it possible to achieve the assigned goals. However, it involves either the use of an annular chamber with a flexible wall or, as a variant, a chamber with two complementary rigid annular walls, which are axially movable relative to one another, the sealing of this chamber. being provided by at least one annular seal whose length is that of the circumference of the chamber.

Due to its integration into a space launcher, such a device must have a high level of reliability. Obtaining it can be very expensive because of the sealing problems posed by the use of an annular chamber with flexible wall or an annular chamber comprising a long annular seal of the same extension.

When choosing the solution of an annular chamber with a flexible wall, it is further observed that when the pressure in this chamber is high, the flexible wall becomes very rigid, to the point that the forces transmitted by the device pass through this wall rather than by the fluid which it confines, which makes the device ineffective at high pressures.

The object of the present invention is precisely to provide a device for supporting a payload in a launcher of said payload in space, which does not have these drawbacks and which, in particular, makes it possible to prevent transmission of surge. efforts from the launcher to the satellite, while having high operating reliability and a moderate production cost.

This object of the invention is achieved, as well as others which will appear on reading the description which will follow, with a device for supporting a payload in a launcher of said payload, said launcher comprising a casing supporting at least one auxiliary propellant applying in operation surges of force to said casing, said device comprising means of connection mechanical of said payload to said casing of said launcher, this device being remarkable in that said connecting means are designed to ensure mechanical coupling of the load and the launcher in the direction of the longitudinal axis of the launcher, and a decoupling of these following the radial and tangential directions.

As will be seen below, these connecting means prevent a significant transmission of the surges of effort to the payload, while ensuring the mechanical cohesion of the launcher / payload assembly.

According to a preferred embodiment of the device according to the invention, the connection means comprise an annular radial rim secured to the payload, an annular groove secured to said casing, said groove slidingly accommodating said rim, in any radial or tangential direction , and stopping any relative movement of said edge in the axial direction. According to another characteristic of the device according to the invention, the connection means further comprise at least one stop means mechanically linking the envelope and the load in at least one of the radial and tangential directions, said stop means being arranged at a node of the cross section, by a plane passing through said stop means, of the deformation of said envelope under surge. Even more preferably, the stop is arranged at a tangential displacement node of the deformation and acts so in a tangential direction passing through this node.

Other characteristics and advantages of the present invention will appear on reading the description which follows and on examining the appended drawing in which:

- Figure 1 shows schematically a part of a satellite launcher of the prior art, described in the preamble to this description, - Figure 2 is a graph showing the peripheral distribution of the axial forces applied to the casing of the launcher when the auxiliary thrusters 8 and 9 operate,

FIG. 3 represents a cross section of the casing of the launcher, deformed by the forces applied to this casing by these propellants, FIGS. 2 and 3 being commented on in the preamble to this description,

FIG. 4 is a schematic plan view of the device according to the present invention, and

- Figures 5 and 6 are sectional views of the device of Figure 4, along the section lines III and IV, respectively.

The device according to the invention is arranged at a plane P perpendicular to the longitudinal axis X of the launcher (see FIG. 1), this mean plane possibly being at any level between the bases 6a and 6b of the skirt 6, advantageously at the base 6b. More generally still, this device could be installed at any level, between the payload and the attachment points 8,9 of the auxiliary thrusters 2,3.

In Figure 4, it appears that the device shown generally takes a circular shape. It is essentially composed of annular members superimposed axially, between which are installed a plurality of elastic stops 12i, a plurality of hydraulic stops 13 _j and, according to the present invention, a plurality of mechanical connection stops 14.

The structure and operation of the elastic stops 12i and of the hydraulic stops 13 _j are the subject of the French patent application filed today by the applicant and entitled "Device for suspending a payload in a space launcher". We can refer to this request for more details concerning these stops. In short, their function is to filter the shocks mentioned in the preamble to the present description, by an appropriate temporary lowering of the axial stiffness of the launcher, at the level of the plane P, shocks which are observed in particular during the release of the cap 7 and during separations of the stages of the launcher.

It will be noted that the distributions of the abovementioned stops are regular and nested one inside the other. By way of illustrative and nonlimiting example only, the device shown thus comprises four elastic stops 12i (i from 1 to 4), four mechanical connection stops 14i and eight hydraulic stops 13 _j (j from 1 to 8) regularly spaced angularly with a pitch of 22.5 °. At the base 6b of the skirt 6, the casing 1a, 5a of the launcher has substantially the circular shape of the device shown in FIG. 2, which follows the outline of said casing. When the auxiliary thrusters 2 and 3 fixed at 8 and 9 respectively on this casing operate, these apply to the casing la, 5a of the launcher of surges of force around the attachment points 8 and 9, as is discussed above in connection with Figure 2. These surflux have the effect of deforming the envelope, 5a (see Figure 3) including in the plane passing through the stops 12i, 14i and 13 _j. It is observed on the model, and it is confirmed by calculation, that this deformation passes through four points regularly spaced angularly from each other by 90 °, on the right section of the casing of the launcher. This distribution of points is spaced at an angle α from an axis passing through the attachment points 8, 9, as shown in FIG. 4.

This rotation is explained by the composite nature of the material used to constitute the casing of the launcher. By way of illustrative and nonlimiting example only, this angle α can be of the order of approximately 10 °.

It is understood that the propagation to the skirt 6, then to the satellite 4, of the deformations which are printed on the casing of the launcher is liable to damage the satellite.

The object of the present invention is precisely to prevent such propagation of these deformations, so as to protect the physical integrity of the satellite. According to the present invention, this object is achieved by decoupling in the radial and tangential directions the deformations of the skirt 6 and the deformations of the envelope 5a of the second stage and by disposing, in each of the four points mentioned above, a stop mechanical connection 1 i coupling the skirt 6 and the casing 5a of the second stage 5.

Reference is now made to FIGS. 3 and 4 of the appended drawing which represent partial views in axial section of the support device according to the invention, taken respectively along the section lines III and IV of FIG. 4, passing through a current section of the device and one (14 ₃ ) of the stops 14i, respectively, it being understood that the other stops of the same type are identical to those which will be described. The support device according to the present invention is described, by way of illustrative and nonlimiting example only, as associated and integrated with the suspension device described in the French patent application filed today by the applicant and entitled "Suspension device from 'a payload in a space launcher'. Of course, these two devices could be used independently of one another.

In FIGS. 3 and 4, 10 and 11 of the first and second annular members are referenced, respectively, the annular member 10 being fixed to the skirt 6 by a circular distribution of bolts such as that referenced 15, and the annular member 11 being mounted on the second stage 5 of the launcher by means of connecting means (20a, 20b, 22, 23) forming part of the present invention.

The annular member 11 itself comprises first and second annular flanges 17,18 respectively, secured by bolts such as that referenced 19. The first flange 17 takes the form of a ring with an axis parallel to the X axis of the launcher, this ring extending from two axially spaced radial annular flanges, one 17a projecting outward to be crossed by bolts such as that referenced 19, the other 17b being turned inward to be caught in an annular groove limited by rings 20a, 20b assembled by bolts such as that referenced 21. The ring 20b is itself fixed to the second stage 5 by bolts such as that referenced 16. The members 17bf20 _a , 20 _b are part of the connecting means of the device according to the present invention.

Skids 22, 23 allow sliding of the rim 17b of the flange 17 in the groove limited by the rings 20a, 20b, transverse to the axis X of the launcher, for a purpose which will be described later in detail. A clamping of the flange 17 can be established by the pads 22, 23 adapted to prevent detachment of this flange relative to these pads, under the action of external forces applied to this flange. The second flange 18 carries the elastic and hydraulic stops mentioned above (not visible in Figures 3 and 4) and described more fully in the aforementioned patent application to which reference will be made for more details concerning them. These stops have the effect of securing the annular member 10 and the flanges 17 and 18, the member 10 then being pressed against the flange 18.

For the purposes of the description of the present invention, it will be considered that the member 10 and the flanges 17,18 form a unitary annular assembly (10,17,18), that is to say a non-deformable block, which is not the case when this assembly is implemented so as to filter shocks, as described in the aforementioned French patent application.

Referring now to Figure 4 of the accompanying drawing, which shows a section of the device according to the invention showing one (14 ₃ ) of the mechanical connection stops forming part of the connecting means of the support device according to the invention .

According to a preferred embodiment of the present invention, this stop 14 ₃ takes the form of a stud with a radial Y axis (see FIG. 4), passing through two radial holes 24.25 drilled respectively in the flange 17 and in a lug 26 projecting from the ring 20a and facing the flange 17. The stud 14 ₃ comprises, from one of its ends, a head 27, a screw part 28 suitable for being fixed in the hole 24, threaded to this effect, and a part 29 with a smooth surface suitable for sliding freely in the hole 25 drilled in the tab 27. It is understood that thus the ring 17 is free to move away or to approach the tab 26, in the radial direction Y.

It is clear that screwing the stud in the hole 25 and free mounting of this stud in the hole 24 would allow the same relative displacements of the tab 26 and the ring 17.

The annular assembly (10,17,18) is therefore free to deform radially at the stops 14i under the action of surges developed by the auxiliary thrusters 2,3.

Because they maintain the coaxiality, on the Y axis, of the holes 24 and 25, the studs 14χ, however, prevent any tangential displacement (in a direction T perpendicular to the axes X and Y, see Figure 4) of the flange 17 relatively to lug 26, and therefore to skirt 6 and satellite 4 secured to the assembly (10,17,18), relative to the envelope 5 of the second stage and to the envelope of the first stage, integral with that of the second floor. To resist crushing under tangential forces, the hole 25 can be lined with a ring 30 made of a suitable metal.

According to the invention, by placing the studs 14i at the tangential displacement nodes of the casing of the launcher under surge, a mechanical connection is established between the satellite 4 and the launcher which ensures the overall cohesion of the assembly which it forms, without hampering radial and / or tangential deformations of the flange 17 relative to the ring 20a, necessary for the decoupling of these members vis-à-vis the surge of force developed by the auxiliary thrusters on the casing of the launcher. The tangential coupling established by the studs or stops 14ι has no effect on this decoupling since these stops are arranged at the tangential displacement nodes, that is to say at points where they are not exerted no tangential forces between the envelope and the annular assembly (10,17,18).

Thanks to the invention, it is possible to very substantially reduce the axial load differences due to the operation of the auxiliary thrusters 2,3, at the interface between the casing 5a and the skirt 6. In a particular implementation of the device according to the invention, invention, it was possible to reduce to less than 5 N / mm, the amplitude of the load variations at this interface, ie to a level which does not cause any appreciable asymmetrical deformation of the skirt 6, which correctly protects the satellite carried by this skirt of a transmission of such deformations.

This result is obtained thanks to the connection means described above which ensure mechanical coupling of the launcher and the satellite along the longitudinal axis X of the launcher, thanks to the cooperation of the annular rim 17b and the groove delimited by the rings 20a, 20b, and a decoupling of the launcher and the satellite in the radial and tangential directions (except at the nodes of tangential displacement), this decoupling being able to prevent any sensitive transmission of surge of forces to the satellite.

Of course, the invention is not limited to the embodiment described and shown which has been given only by way of example. Thus the mechanical coupling by stops of the ring 20a and the annular assembly (10,17,18) could be established at nodes of radial displacement of the deformation of the casing of the launcher, rather than to the tangential displacement nodes. One can calculate, and control on model, the position of these radial displacement nodes, as for the tangential displacement nodes. The stops used to establish this coupling are very consistent with the stops 14ι except that the axes of the studs are tangential instead of being radial and therefore pass in coaxial holes on a tangential axis such as T (see Figure 4) drilled in corresponding tabs depending on the annular assembly (10,17,18) and the ring 20a, respectively. Experiments and calculations have shown that the mechanical cohesion of the launcher / satellite assembly obtained by such "radial" stops, arranged where the deformations of the envelope in the radial direction are substantially zero is good, although slightly less to that obtained by the "tangential" stops.

Alternatively, one could ensure said cohesion using a set of stops comprising both radial stops and tangential stops. Similarly, the device according to the invention has been described as integrated into the suspension device with hydraulic and elastic stops described in the aforementioned French patent application. It is clear that the support device according to the invention could be physically removed from such a suspension device, or even be used alone in the event that there is no need to treat the shocks that the suspension device has for purpose of filtering.

Claims

CLAIMS 1. Device for supporting a payload (4) in a launcher (1) of said payload (4), said launcher comprising an envelope (la, 5a) supporting at least one auxiliary propellant (2,3) applying during operation of the surges of effort to said envelope (la, 5a), said device comprising means for mechanical connection of said payload (4) to said envelope (5) of said launcher, characterized in that said connection means (10 , 17, 18, 14i) are designed to ensure mechanical coupling of the load (4) and the casing (la, 5a) in the direction of the longitudinal axis (X) of the launcher, and a decoupling of these ci following the radial and tangential directions.

2. Device according to claim 1, characterized in that said connecting means comprise an annular flange (17b) integral with said payload (4), an annular groove (20a, 20b) integral with said envelope (la, 5a) , said groove slidingly accommodating said flange (17b) in any radial (Y) or tangential (T) direction, and stopping any movement of said flange (17b) relative to said groove in the direction of the longitudinal axis (X) of the launcher.

3. Device according to claim 2, characterized in that sliding means (22,23) are interposed between the transverse faces opposite said flange (17b) and said groove (20a, 20b).

4. Device according to any one of claims 1 to 3, characterized in that said connecting means further comprises at least one stop means (14i) mechanically connecting the envelope (5) and the payload (4) along at least one of said radial and tangential directions.

5. Device according to claim 4, characterized in that said stop (14j _. ) Is arranged at a displacement node of said envelope in a predetermined direction, so as to oppose an overall displacement of the envelope (the , 5a) with respect to the payload (4).

6. Device according to claim 5, characterized in that said predetermined direction (T) is tangential to the envelope (la, 5a).

7. Device according to claim 5, characterized in that said direction (Y) is radial relative to the envelope (la, 5a).

8. Device according to claim 5, characterized in that it comprises at least one stop (14i) arranged to oppose relative displacement in a tangential direction (T) and at least one other stop arranged to oppose relative displacement in a radial direction (Y).

9. Device according to any one of claims 4 to 8, characterized in that said stop

(14i) is constituted by a cylindrical stud passing through two coaxial holes (24,25) drilled respectively in a part (20a) integral with the envelope (la, 5a) and in a part (17) integral with the payload (4 ).

10. Device according to claim 9, characterized in that said stud (14i) is slidably mounted in at least one of said holes (24,25).

11. Device according to claim 10, characterized in that said stud (14i) is fixed to one of said parts.

12. Device according to any one of claims 4 to 11, characterized in that it comprises a plurality of stops (14i) each disposed at a displacement node, in a predetermined direction, of the envelope (la, 5a ) under surge.

13. Space launcher equipped with the device according to any one of claims 1 to 12.