WO2001071126A1

WO2001071126A1 - Deployable space frame and method of deployment therefor

Info

Publication number: WO2001071126A1
Application number: PCT/US2000/007706
Authority: WO
Inventors: David P. Cadogan
Original assignee: Ilc Dover, Inc.
Priority date: 2000-03-23
Filing date: 2000-03-23
Publication date: 2001-09-27
Also published as: AU2000263338A1

Abstract

A space frame (100) capable of facilitating the deployment, and subsequent support, of a space structure includes a packageable, deployable, and rigidizable frame assembly (110), a packageable, deployable, and inflatable frame assembly shell (120) disposed around the frame assembly (110); means for attaching the frame assembly shell (120) to the frame assembly (10); and a shell inflator. The method of packaging and deploying the space frame (100) includes (a) collapsing the frame assembly (110) by packaging the shell (120) to provide a packaged frame assembly/shell; (b) controllably deploying the frame assembly (110) and the shell (120) from the packaged frame assembly/shell by employing the shell inflator to inflate the shell (120) while imparting a resistance to the shell (120) to resist deployment such that the internal gas pressure required to continue deployment is sufficient to fully inflate that portion of the shell (120) to which gas has been introduced; (c) continuing to resist deployment until the frame assembly (110) and the shell (120) are deployed; (d) terminating the introduction of gas into the shell (120); (e) rigidizing the frame assembly (110); (f) depressurizing the shell (120).

Description

DEPLOYABLE SPACE FRAME AND METHOD OF DEPLOYMENT THEREFOR

BACKGROUND OF THE INVENTION 1. Field of the Invention

This invention relates generally to a method and an apparatus capable of deploying and

subsequently supporting a lightweight space structure such as a solar array, reflector, sunshield,

radar array, antenna, or concentrator. The invention relates more specifically to a method and

lightweight apparatus that provide for the controlled deployment of an inflatable shell containing

a space frame assembly, and the subsequent rigidization of the assembly so as to support the space

structure.

2. Description of Related Art

Most conventional methods for deploying and supporting a space structure accomplish

the deployment by means of deployable truss structures consisting of relatively heavy elements

such as rigid members, hinges, latches, and cables. Increases in the number of satellites to be

launched over the next several decades, however, will emphasize the need for the reduction of

space hardware mass, stowage volume, and cost.

One approach to realizing these reductions is through the use of inflatable, deployable,

space structures. Inflatable structures offer many benefits over conventional deployable structures

because they are lower in mass and can be packaged into small volumes, which reduces launch

vehicle size and cost. The performance benefit margin of inflatable structures increases as the size

of the structure increases, thus making the technology more attractive for large-scale systems.

Examples of satellite components that benefit from the utilization of inflatable structures include

solar arrays, communications antennas, radar antennas, thermal/light shields, and solar sails. Although inflatable tubular structures weigh less than deployable truss structures

consisting of elements such as rigid members, hinges, latches, and cables, the weight of the

inflatable tubular structures is not insignificant.

Therefore, a need exists for a method and an apparatus capable of facilitating the

deployment, and subsequent support, of a space structure, but which can do so with an overall

apparatus weight which is less than that of conventional inflatable deployment structures.

SUMMARY OF THE INVENTION

It is an object of the present invention to provide a method and an apparatus capable of

deploying and subsequently supporting a space structure. It is a further object of the present

invention to provide a method and an apparatus capable of facilitating the deployment by

controlling both the rate and the directionality of deployment. It is a still further object of the

present invention to provide a method and an apparatus capable of facilitating the deployment,

and subsequent support, of the space structure, but which can do so with an overall apparatus

weight which is less than that of conventional inflatable deployment structures.

Accordingly, the present invention is directed to a deployable space frame comprising a

packageable, deployable, and ngidizable frame assembly; a packageable, deployable. and inflatable

frame assembly shell disposed around the frame assembly; means for attaching the frame assembly

shell to the frame assembly; and a shell inflator.

The present invention is also directed to a method of packaging and deploying the space

frame. The method comprises (a) collapsing the frame assembly by packaging the shell to provide

apackaged frame assembly/shell; (b) controllably deploying the frame assembly and the shell from

the packaged frame assembly/shell by employing the shell inflator to inflate the shell while

imparting a resistance to the shell to resist deployment such that the internal gas pressure required to continue deployment is sufficient to fully inflate that portion of the shell to which gas has been

introduced; (c) continuing to resist deployment until the frame assembly and the shell are

deployed; (d) terminating the introduction of gas into the shell; (e) rigidizing the frame assembly;

and (f) depressurizing the shell. Thus, as the shell inflates, it deploys both the shell and the frame assembly contained

therein in a controlled manner. Once the deployed shell is fully inflated, the frame assembly is

rigidized to provide the support for the space structure. The shell is then depressurized and

assumes an essentially passive role, apart from providing environmental protection for the

rigidized frame assembly.

The advantages associated with the present invention are numerous. First, as indicated

above, most conventional methods for deploying and supporting a space structure accomplish

these tasks by means of either deployable truss structures consisting of relatively heavy elements,

or inflatable structures. The space frame, however, requires substantially less material to

accomplish the same structural performance as conventional inflatable structures, and is therefore

lower in mass. The present invention, therefore, provides a lightweight means for both deploying

and supporting a variety of space structures. Furthermore, by virtue of its inflatable means for

deployment and rigidizable means for support, the invention comprises a lighter-weight system

with fewer moving parts, thereby providing a higher degree of system reliability. Finally, the

ability to collapse the structure for launch results in a packed volume which is minimal when

compared with the aforementioned conventional structures.

Thus, while specifically facilitating as its primary application the deployment of lightweight

space structures, the invention also more generally provides for the deployment of any lightweight

structure that can then be maintained in place by the rigidity of the frame assembly. BRIEF DESCRIPTION OF THE DRAWINGS

Other objects, features, and advantages of the present invention will become more fully

apparent from the following detailed description of the preferred embodiments, the appended

claims, and the accompanying drawings. As depicted in the attached drawings:

FIG. 1 is a partial perspective view of a space frame according to a first preferred

embodiment of the present invention in which the shell and the frame assembly are in their

deployed position.

FIG. 2 is a cross-sectional view of the shell and the frame assembly depicted in FIG. 1.

FIG. 3 is a cross-sectional view of the shell and the frame assembly according to a second

preferred embodiment of the present invention.

FIG. 4 is a cross-sectional view of the shell and the frame assembly according to a third

preferred embodiment of the present invention.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

The present invention will be disclosed in terms of the currently perceived preferred

embodiments thereof.

The deployable space frame is an ultra-lightweight structural member that is simple to

manufacture and readily tailorable to meet the structural requirements of a specific application.

The space frame comprises a truss beam that comprises a series of connected box frames, or bays,

manufactured from low-mass rigidizable composite tubes, enveloped within an inflatable outer

shell. The deployable space frame requires substantially less material to accomplish the same

structural performance as conventional cylindrical inflatable tube structures, and is therefore lower

in mass. Furthermore, the deployed beam becomes a rigid structural member that does not

require the support of inflation pressure to maintain its shape. Referring to FIGS. 1 and 2. a deployable space frame 100 constructed in accordance with

a first preferred embodiment of the present invention is shown in its deployed position. In this

embodiment, the space frame 100 comprises a packageable. deployable. and rigidizable frame

assembly 110 having an assembly base end and an assembly tip end; a packageable. deployable.

and inflatable frame assembly shell 120 disposed around the frame assembly, the shell having a

shell base end and a shell tip end; means for attaching the frame assembly shell to the frame

assembly; and a shell inflator.

Frame assembly 110 comprises a plurality of connected thin- walled rigidizable composite

struts which define a series of connected box frames, or bays, and has a polygonal cross-sectional

shape (see FIG. 2). In the embodiment depicted in FIGS. 1 and 2. frame assembly 1 10 comprises

first, second, and third longerons 111 extending from the assembly base end to the assembly tip

end, and a plurality of connecting struts 112 interconnecting the first, second, and third longerons

so as to form a triangular cross-sectional strut configuration.

Struts 111 and 112 can comprise various materials of construction and various shapes

depending on the structural strength required. In a preferred embodiment, the struts are thin-

walled composite laminate tubes of a thermosetting or thermoplastic material, with a wall

thickness which is determined by the structural requirements of the specific application.

The means for attaching the frame assembly shell to the frame assembly serves to connect

the struts 111 to the inside surface of the shell 120. These connections facilitate the proper

packing and deployment of the shell as the frame assembly is packed and deployed. Suitable

means for attaching the frame assembly shell to the frame assembly include tabs that are affixed

to the inside surface of the shell and to the frame assembly. The tabs can be attached by means

such as bonding and/or stitching. The space frame properties are derived through the use of finite element beam modeling.

Equivalent areas and moments of inertia are determined by applying unit loading to the finite

element models, determining deflections and using standard structural mechanics equations to

back calculate the properties. A geometrical configuration is determined such that the capacity

of the space frame is approximately equal to the required compressive force. For example, in one

application employing struts having a wall thickness of from 50 to 75 m, and a diameter of from

0.64 to 1.27 cm, the frame assembly cross-sectional dimensions are such that the frame assembly

cross section is capable of being inscribed within a 20.5 cm diameter circle.

Shell 120 is used for deployment, governs the deployed structure's straightness. and can

provide some protection of the rigidized frame assembly elements from environmental degradation

or thermal loads through the application of coatings, if required. The shell also provides the

system^'s structure while in the inflated-only state. Often, this is the governing parameter in the

system's design. The shell supports the loads from inflation and transfers those loads into the

rigidizable space frame assembly to tension the members prior to rigidization.

The strength of the space frame during deployment is derived from the inflatable shell.

Prior to rigidization. the inflatable shell governs the cross-sectional moment of inertia of the frame

assembly, which makes it advantageous compared to other approaches such as an inflatable truss.

Inflation of the shell yields a relatively high cross-sectional moment of inertia as compared to an

inflatable truss because the moment of inertia is dependent on a cubic value of the radius. The

larger radius of the inflatable portion of the space frame yields a relatively high stiffness during

deployment as compared to a truss with small diameter inflatable members. Therefore, the shell

provides the capability to withstand off-nominal loading during the deployment as well as during

flight modes. In a preferred embodiment, shell 120 comprises a polyimide film having a thickness of 12

μm, and in the embodiment depicted in FIGS . 1 and 2, has a substantially circular cross-sectional

shape. Once the frame assembly is deployed and rigidized. the shell is depressurized and takes

a passive role in system performance, apart from providing environmental resistance.

The shell inflator comprises a pressure-regulated gas source, such as nitrogen gas or a gas

generated on board the spacecraft once on orbit, a plurality of redundant valves, a plurality of

pressure sensors, and a plurality of pressure relief valves.

Referring to FIG. 3, a deployable space frame 200 constructed in accordance with a

second preferred embodiment of the present invention is shown in its deployed position. In this embodiment, frame assembly 210 comprises a plurality of connected thin- walled composite struts

and has a polygonal cross-sectional shape. In the embodiment depicted in FIG.3. frame assembly

210 comprises first, second, third, and fourth longerons 211 extending from the assembly base

end to the assembly tip end, and a plurality of connecting struts 212 interconnecting the first,

second, third, and fourth longerons so as to form a rectangular cross-sectional strut configuration.

The use of the term "shell" herein is meant to denote not only conduits such as the

aforementioned shell having a substantially circular cross-sectional shape, but also includes all

other cross-sectional shapes which are capable of being packaged and deployed according to the

present invention. For example, the use of a rectangular cross-sectional strut configuration

analogous to that depicted in FIG. 3 but in which the length of the rectangle is substantially

greater than the width of the rectangle enables the approximation of a non-circular cross-sectional

shell shape.

Furthermore, as depicted in FIG. 4, a deployable space frame 300 constructed in

accordance with a third preferred embodiment of the present invention comprises the same

triangular cross-sectional strut configuration as the embodiment depicted in FIG.2. but includes a shell 320 having a lobed cross-sectional shape. Additionally, even though the lobed cross-

sectional shape embodiment has been depicted in conjunction with the triangular cross-sectional

strut configuration, an analogous lobed cross-sectional shape could be employed with the

rectangular cross-sectional strut configuration depicted in FIG. 3 (i.e.. a shell having a lobe

associated with each side of the rectangle). Finally, since, as indicated above, the frame assembly

comprises a plurality of connected struts which define a polygonal cross-sectional shape, in the

general case, a lobed cross-sectional shape can be employed in which the number of lobes is equal to the number of sides of the polygon.

The method of packaging and deploying a space frame 100. 200, or 300 (referred to

herein in the description of the method as 100 for the purpose of brevity) comprises the following

series of steps. First, frame assembly 110 is collapsed from the assembly tip end to the assembly

base end by packaging shell 120 from the shell tip end to the shell base end to provide a packaged

frame assembly/shell. In a preferred embodiment, the packaging step is accomplished by rolling

the frame assembly-containing shell from the shell tip end to the shell base end. In another

possible embodiment, however, the frame assembly-containing shell could be folded from the shell

tip end to the shell base end in an accordian-like fashion.

Once on orbit, frame assembly 1 10 and shell 120 are controllably deployed from the

packaged frame assembly/shell by employing the shell inflator to introduce an inflation gas into

shell 120 so as to inflate the shell while imparting a resistance to the shell to resist deployment

such that the internal gas pressure required to continue deployment is sufficient to fully inflate that

portion of the shell to which said gas has been introduced. The introduction of the gas is

continued and the resistance to deployment is maintained until frame assembly 110 and shell 120

are deployed. Once frame assembly 1 10 and shell 120 are deployed, the introduction of gas into the shell

is terminated, the frame assembly is rigidized, and the shell is depressurized.

The method of packing the space frame minimizes volume and ensures deployment

reliability. The packing method is also low in mass, utilizes flight proven technology, and

provides the required protection of the system during launch vibration.

While the inflation system can take several forms, a preferred embodiment is bottled N₂

gas, or cold gas from an existing source on the spacecraft, in order to reduce system risks. The

shell is pressurized by a regulated gas source that is fed to the appropriate chambers at the

required rates and times via valves that are actuated by computer or by human intervention, such

as by radio signal, etc. Redundant valves and pressure sensing transducers are employed in the

system to reduce risk, and the shell is fitted with relief valves to prevent over-pressurization.

Regardless of which configuration of the frame assembly is being employed, successful

operation of the space frame depends on maintaining the stiffness of the inflatable shell while the

frame assembly and shell are being controllably deployed. To achieve a controlled deployment,

it is necessary to incorporate sufficient resistance to deployment such that the internal pressure

required to continue deployment is high enough to adequately stiffen the portion of the shell that

has already deployed. That is, the rate of deployment is controlled by a balance of forces between

the resistance of the controlled deployment device and the flow rate and pressure of the inflation

gas. During deployment, the internal pressure in the shell must be maintained at a prescribed level

in order to yield the required skin stress in the inflatable shell^" s wall to react to loads on the

system. This provides for a slow, controlled deployment which minimizes the possibility of film

stress and film surface rubbing.

The roll-up embodiment of the present invention, which causes the least damage to the

stowed shell and produces the smoothest deployment, comprises the rolled inflatable shell with a means for imparting resistance to deployment, i.e.. a means for controlling the rate of unrolling

when the inflation gas is introduced. The two general classes of deployment control which can

be used in the roll-up embodiment include: i) means embedded in either the interior or exterior

wall of the shell itself, and ii) means mounted at the deploying end of the shell.

Examples of the means for imparting resistance to unrolling which are embedded in the

shell wall itself and which provide for adhesion of the rolled-up exterior shell wall surfaces to one

another include a pressure sensitive adhesive, and a hook-and-loop fastener tape such as "NELCRO." In each of these means, the rolled shell remains in its packed state until the inflation

pressure overcomes the resistance provided by the means and initiates deployment.

In a first embodiment of the means for imparting resistance to unrolling, a pressure

sensitive adhesive is affixed to the exterior of the shell in longitudinal strips disposed at the 10 and

2 o'clock positions around the circumference of the shell. The adhesive can be used to control

deployment as well as assist in maintenance of the package shape during launch vibration. The

separation strength of the rolled shell from the outer wall of the shell, which dictates the internal

pressure, is determined by the adhesive peel strength and the width of the strips of adhesive. The

peel strength of this adhesive is constant over a wide temperature range about the predicted

deployment temperature. The adhesive comprises high molecular weight compounds having high

vacuum stability and therefore low outgassing characteristics. This embodiment requires only

a few grams per linear meter of adhesive, and represents the lowest mass approach possible.

In a second embodiment of the means for imparting resistance to unrolling, a plurality of

hook-and-loop fasteners embedded in the exterior wall of the shell can be used to control

deployment and assist in maintenance of the package shape. This is accomplished by adding four

independent strips of hook-and-loop fastener to the shelfs exterior: two strips of hook at thelO

and 2 o'clock positions, and two strips of loop at the 8 and 4 o'clock positions. The shell is then flattened and rolled about the 9 to 3 o'clock axis. When inflation gas is introduced, the shell

expands, causing the hook on the top side of the shell to detach from the loop on the bottom side

of the shell, thereby allowing the shell to unroll. By selecting various grades and widths of hook-

and-loop fastener, resistance to unrolling can be predicted and controlled, thus yielding the

internal pressure required to provide the specified shell stiffness.

An example of a means for imparting resistance mounted at the deploying end of the shell

is a means for imparting mechanical torque. Such devices function by torque reaction of the

rolled shell on the inflated portion of the shell in order to control deployment of the system. This

means can be, for example, a frictional device or a ratchet mechanism, such as a means for

imparting torque through plastic deformation of a wire or tape of material. Such devices are

typically low in mass and are highly reliable, but result in added tip mass which is unacceptable

with applications in which bending loads are expected to be encountered from satellite

attitude/reaction control system loads.

Successful deployment of the space frame is also dependent upon the ability of the frame

assembly struts to be collapsible for packaging, and then deployable to shape when the shell is

inflated. The struts, therefore, comprise a thermoplastic shape memory composite material. The

thermoplastic shape memory composite will return to its manufactured "set" shape when heated

above the second order transition temperature. Cooling the material below its second order

transition temperature will then cause the material to become rigid. This allows the struts to be

collapsed and packaged and later deployed to their final frame assembly shape.

In a preferred embodiment, the struts comprise thermoplastic composite laminate material

which rigidizes by cold rigidization. Other possible embodiments of the strut materials, however,

include materials such as ultraviolet radiation rigidizable materials and chemically hardened

structures. The functionality of the rigidization technique dictates the process by which the frame

assembly is deployed and rigidized. First, the packaged frame assembly is preheated either by the

heat given off by the spacecraft, by solar radiant energy, or by small heaters embedded in the

packed volume. The packaged frame assembly would have good conduction paths and would be

essentially the same temperature throughout the package. The preheating is necessary to warm

the composite material above its second order transition temperature and provide it with some

flexibility so as to allow it to be deployed via the inflation of the shell. During deployment, that

portion of the packaged frame assembly which has not yet been deployed is housed in a multilayer

insulation ("MLF') cover throughout its deployment. The cover contains the heat in the packaged

frame assembly during the slow inflation process, thus keeping it in its flexible state until fully

deployed.

Once the frame assembly emerges from the MLI cover it begins to give off heat and

harden. The composite material^* s cooling rate is dictated by the insulating capability of the shell

and any insulation layers added to the struts themselves. The insulation could be in the form of

vapor deposited aluminum ("NDA") coated polyimide, and can be tailored to release heat at any

rate to give the frame assembly structural capability, beyond that of the shell, even before

deployment is complete. The MLI cover also controls the fluctuations in the temperature of the

inflation gas and thus minimizes the quantity required if the system passes in and out of eclipse.

Both the thermoplastic and ultraviolet curable (thermoset) materials have been

demonstrated with success in the manufacture and laboratory test of 3 meter long. 15 cm diameter

booms. Advantages of thermoplastic materials, however, include the low coefficient of thermal

expansion of the composite, low outgassing, ease of manufacture, reversibility to facilitate

multiple functional tests in lab ambient conditions, and lower complexity. Numerous thermoplastic materials are possible as matrix resins for use in composite struts.

These materials are selected based on their performance properties, processing and manufacturing

capability, and service temperature in various environments. In a preferred embodiment, the

material is a modified thermoset that mimics a thermoplastic. Its properties include ease of

processing, mechanical properties, thermal performance (softening point >70 ° C for lab ambient

testing and thermal margin on orbit), low creep, and ability to function as a shape memory plastic.

The resin can be applied to various reinforcements such as graphite, "KEVLAR," glass

and/or ceramic fiber, poly(p-phenylene-2.6-benzobisoxazole)("PBO"), which is a rigid-rod

isotropic crystal polymer, and "VECTRAN." Various weave styles are available with the required

sizing to ensure wetting of the fibers and adhesion of the matrix resin. Hybrid weaves may also be a potential method of improving the reinforcement's capability while still retaining all of the

flexibility required for manufacture and packaging.

An advantage of the present space frame is that a thermoplastic material will have a

relatively high modulus unless heated to a temperatures which causes it to become very soft. This

is an important consideration during deployment of high aspect ratio frame assemblies because

the micro- wrinkles developed on the inboard side of the roll of a rolled shell can be impossible to

remove with inflation pressure if the material has any significant stiffness (modulus) during

deployment. The result, if this were to occur, would be curved structures. This concern is

mitigated in the frame assembly because the length delta of the wall of the individual rolled

members is relatively low and the truss members are under greater relative tension than

conventional cylindrical constant wall thickness booms. The fact that the rigidizable material is

also a shape memory plastic that is programmed to return to its original shape also mitigates the

risk of curved structures. The present invention, therefore, by making possible the use of a variety of lightweight

frame assemblies and devices for providing resistance to deployment, facilitates the deployment

of a variety of lightweight space structures such as solar arrays, reflectors, sunshields, radars,

antennas and concentrators. Although the invention' s primary application is in the deployment

of space structures, one skilled in the art can appreciate that the invention could be employed in

other environments that require the deployment of a lightweight structure.

The advantages of the space frame are numerous. First, the space frame minimizes the

potential for premature rigidization. both pre- and post-launch. Since the frame assembly is

manufactured from thermoplastic materials, it is impossible to have premature rigidization prior

to deployment. Heating of the package during ascent from free molecular heating during launch

will only preheat the assembly and is very desirable. The space frame is covered with MLI locally

as well as in the area over the rolled-up section to prevent premature heat loss and rigidization

during deployment. The area over the roll may also be covered with a high alpha/epsilon material,

such as VDA-coated film, to promote the retention of heat during deployment. The cover would

deploy to an area where it would allow the tip of the frame assembly to rigidize when the

structure was fully deployed. Therefore, the potential for premature rigidization is extremely low.

Secondly, the space frame provides tolerance to increasingly hostile environments, in

particular, increased thermal loads, radiation, and spacecraft-induced contamination, as well as

insensitivity to close proximity to either warm or cold structures. Environmental protection of

the frame assembly is provided by the resiliency of the rigidizable materials and the protective

capability of the MLI that locally wraps the struts, or the shell that also acts as part of the

insulation blanket.

Thirdly, the space frame minimizes stowage volume and accommodates stowage in

different stowage shapes. While only certain preferred embodiments of this invention have been shown and described

by way of illustration, many modifications will occur to those skilled in the art and it is, therefore,

desired that it be understood that it is intended herein to cover all such modifications that fall

within the true spirit and scope of this invention.

For example, the space frame can be tailored to meet the structural requirements of a

specific application. Each of these modifications has little effect on the manufacturing process

and cost of the system because the manufacturing techniques can easily accommodate many

changes. Several design variables of the space frame which can be altered to optimize the

structural characteristics include, for example, tapering the shell diameter from base to tip;

resizing bays by altering the length of the struts and interface locations to the longerons; altering

the material (fiber type, fiber orientation, resin type) and thickness of the structural members; and

changing the diameter of the structural members to effect the cross-sectional moment of inertia.

Optimization of the space frame design by changing materials, processing techniques or

geometry will also result in reduced linear mass densities of the space frame. The resulting

reductions in system mass would result in a reduction in spacecraft bus mass and/or spacecraft

expendables.

By way of further example of modifications within the scope of this invention, while the

frame assembly has been described as having either a triangular (FIGS. 1, 2. and 4) or a

rectangular (FIG.3) cross-sectional configuration, another embodiment could comprise any other

polygonal configuration (effected by altering the number of longerons) capable of being packaged

and deployed, and of providing the requisite structural support.

shell has been described as having a substantially circular cross-sectional shape in the first and

second preferred embodiments, other embodiments could comprise other shapes capable of being packaged and deployed, and of accommodating the configuration of the frame assembly, such as.

for example, the aforementioned lobed configuration of the third preferred embodiment.

space frame has been described in the context of a single straight frame, another embodiment

could comprise a configuration in which a plurality of space frames are interconnected to provide

a faceted torus or similar shape.

Another possible embodiment could comprise a configuration in which a plurality of space

frames are interconnected to provide a larger truss assembly. In the truss embodiment, the space

frame comprises a plurality of packageable, deployable. and rigidizable frame assemblies, each of

the assemblies having an assembly first end and an assembly second end; a plurality of

packageable, deployable, and inflatable frame assembly shells each corresponding to each of the

plurality of frame assemblies, each of the shells disposed around each of the corresponding frame

assemblies, and each of the shells having a shell first end and a shell second end; means for

attaching each of said shells to a corresponding one of each of said frame assemblies; means for

connecting said assembly second end to said assembly first end for each of said plurality of frame

assemblies, and means for connecting said shell second end to said shell first end for each of said

plurality of shells; and a shell inflator for inflating the plurality of shells.

Claims

What is claimed is:

1. A deployable space frame, said space frame comprising:

a packageable, deployable, and rigidizable frame assembly, said frame assembly having an

assembly base end and an assembly tip end;

a packageable, deployable, and inflatable frame assembly shell disposed around said frame

assembly, said shell having a shell base end and a shell tip end;

means for attaching said frame assembly shell to said frame assembly; and

a shell inflator.

2. A deployable space frame according to claim 1 , wherein said frame assembly

comprises a plurality of connected struts.

3. A deployable space frame according to claim 2, wherein said struts are thin-walled

composite tubes having a wall thickness of from 50 to 75 μm.

4. A deployable space frame according to claim 3. wherein said struts comprise a

thermoplastic material and have a diameter of from 0.64 to 1.27 cm.

5. A deployable space frame according to claim 3, wherein said struts comprise a

thermoset material and have a diameter of from 0.64 to 1.27 cm..

6. A deployable space frame according to claim 1. wherein said frame assembly has

a polygonal cross-sectional shape.

7. A deployable space frame according to claim 6. wherein said frame assembly comprises first, second, and third longerons extending from said assembly base end to said

assembly tip end, and a plurality of connecting struts interconnecting said first, second, and third

longerons.

8. A deployable space frame according to claim 1, wherein said shell has a

substantially circular cross-sectional shape.

9. A deployable space frame according to claim 1, wherein said shell has a lobed

cross-sectional shape.

10. A deployable space frame according to claim 1, wherein said shell comprises a

polyimide film material.

11. A deployable space frame according to claim 8, wherein said shell has a thickness of l2 μm.

12. A deployable space frame according to claim 1, wherein said shell comprises

means for imparting resistance to inflation.

13. A deployable space frame according to claim 12, wherein said means for imparting

resistance to inflation is strips of adhesive tape or hook-and-loop fastener tape.

14. A deployable space frame according to claim 1. wherein said shell inflator

comprises a pressure-regulated gas source, a plurality of redundant valves, a plurality of pressure

sensors, and a plurality of pressure relief valves.

15. A deployable space frame, said space frame comprising: a plurality of packageable, deployable. and rigidizable frame assemblies, each of said

frame assemblies having an assembly first end and an assembly second end;

a plurality of packageable. deployable, and inflatable frame assembly shells each

corresponding to each of said plurality of frame assemblies, each of said shells disposed around

each of said corresponding frame assemblies, and each of said shells having a shell first end and

a shell second end;

means for attaching each of said shells to a corresponding one of each of said frame

assemblies;

means for connecting said assembly second end to said assembly first end for each of said

plurality of frame assemblies, and means for connecting said shell second end to said shell first end

for each of said plurality of shells; and

a shell inflator for inflating said plurality of shells.

16. A method of packaging and deploying a space frame, said space frame comprising

(i) a packageable, deployable, and rigidizable frame assembly, said frame assembly having an

assembly base end and an assembly tip end; (ii) a packageable. deployable, and inflatable frame

assembly shell disposed around said frame assembly, said shell having a shell base end and a shell

tip end; (iii) means for attaching said frame assembly shell to said frame assembly: and (iv) a shell

inflator. said method comprising: (a) collapsing said frame assembly from said assembly tip end to said assembly base end

by packaging said shell from said shell tip end to said shell base end to provide a packaged frame

assembly/shell; (b) controllably deploying said frame assembly and said shell from said packaged frame

assembly/shell by employing said shell inflator to introduce an inflation gas into said shell so as

to inflate said shell while imparting a resistance to said shell to resist deployment such that an

internal gas pressure required to continue deployment is sufficient to fully inflate that portion of

the shell to which said gas has been introduced;

(c) continuing to resist said deployment until said frame assembly and said shell are

deployed;

(d) terminating the introduction of said gas into said shell;

(e) rigidizing said frame assembly; and

(f) depressurizing said shell.

17. A method of packaging and deploying a space frame according to claim 16,

wherein said step (a) packaging is accomplished by rolling.

18. A method of packaging and deploying a space frame according to claim 16,

wherein said step (a) packaging is accomplished by folding.

19. A method of packaging and deploying a space frame according to claim 16,

wherein at least said steps (b), (c), and (d) are conducted outside the atmosphere of the earth.