INFLATABLE DEPLOYABLE SPACE STRUCTURES
TECHNOLOGY SUMMARY
R. E. Freeland ’
Jet Propulsion Laboratory
California Institute of Technology
G. D. Bilyeu ’
G. R. Veal *
L’Garde, Incorporated
M. M. Mikulas ’
University of Colorado
ABSTRACT
There has been limited interest in inflatable
deployable space structures since the 1950’s due to
their potential for low cost flight hardware,
exceptionally high mechanical packaging efficiency,
deployment reliability and low weight. A mnnber of
significant technology developments focused on the
demonstration of such potential include the Goodyear
antermas in the early 196O’s, the Echo Balloon series
from the late 1950’s to the early 196O’s, the Contraves
antennas and sun shades in the late 1970’s to the mid
1980’s and the L’Garde, Inc. inflatable decoys in the
1970’s and mid 1980’s and their space shuttle launched
Inflatable Antenna Experiment (IAE) in May 1996.
Results of this work, especially the IAE, have recently
attracted user interest. The determination of how well
the capability of this new class of space structures can
meet the requirements of specific applications is based
on a combination of issues that include structural
concept maturity, technology data base and the
capability for analytical performance simulation. The
maturity of the IAE antenna concept and the associated
technoloay data base are currently under evaluation for
a number of potential missions. Current analytical
simulations are addressing membrane reflector shape,
reflector error sources, stiffness and thermal stability of
rigidized inflatables and potential reflector shapes
resulting from changes of boundary condition.
*Technical Marlaser, IN-ST!ZP Inflatable Antenna Experiment
’ Pro-g-am Manager, Inflatable Antenna Experiment
’ Inflatable Antenna Experiment Principal Investigator
%ofessor, Center for Aerospace Structures
Copyright @ 1998 by the International Astronautical Federation.
The U.S. Government has a royalty-tie license to exercise all rights
under the copyright claimed herein for Governmental purposes. All
other ri$~ts are reserved by the copyright owner.
Collectively, these hardware demonstration results, the
current technology data base and the new analytical
tools for projecting orbital performance will enable
realistic estimates of the applicability of inflatable
space structures for specific applications.
INTRODUCTION
Concepts for inflatable deployable space
structures have been under development and evaluation
for almost 50 years. The potential for this class of
space structures for achieving low cost space hardware,
exceptional mechanical packaging efficiency,
deployment reliability, very small stowed volume and
low weight was recognized by only a limited segment
of the space structure community.
Over this period a number of different technology
developments have contributed to the validation of
mechanical performance and the establishment of
credibility for this new class of space structures. Such
validations were based on building inflatable
deployable structures up to 30 meters in size for both
ground based and on-orbit demonstrations. Such
results are the basis of the current technolo_y data base
for a munber of different structural concepts.
The results of these demonstrations have recently
generated serious interest within the antenna user
community. However, the determination of the
applicability of this technology to a specific mission
will require a) the capability for analytical simulation
of mechanical performance for each candidate system,
b) a well defmed structural concepts technolo*v data
base, and c) a system concept defmition for the specific
application, which will be based on the most mature
elements of the technology data base.
This publication will a) identify and describe
some of the most significant concept developments, b)
discuss the inflatable structures technology data base
associated with the IAE, c) overview previous
analytical performance projection capability and
identify new analytical capability along with sample
results, and d) summarize the types of interaction of
inflatable materials with the environment and show
recent results on the effects of long term radiation on
thin film materials.
SIGNIFICANT INFLATABLE STRUCTURES
CONCEPT DEVELOPMENTS
A number of different technology developments,
focused on inflatable deployable space structures, have
taken place over the past 50 years. Some were very
innovative for the time frame of their development and
many were not well documented. As examples of
significant structural concept developments, over a 40-
year period, antennas developed by Goodyear, the
Echo Balloons, the Contraves Antennas and Sunshades
and the NASA Inflatable Antenna Experiment will be
briefly summarized.
Goodvear Inflatable Structures
One of the original pioneers associated with the
development of inflatable deployable space structures
was Goodyear. In the time frame of the late 1950’s to
the mid 196O’s, they developed inflatable structural
concepts for their search radar antenna, radar
calibration sphere and lenticular inflatable parabolic
reflector.’
The inflatable search radar antenna is based on
using a truss type rigidizable support structure and
metallic mesh for the aperture surface, Figure 1. The
demonstration hardware, shown in Figure I, has an
aperture of approximately 10 meters in length, 3 meters
in width and has a parabolic profile. The mechanical
packaging technique appears to be based on folding the
long narrow structure into about 6 to 8 flat panels so
that the stowed volume was rectangular in geometry.
Figure 1. Inflatable Search Radar Figure 2. Radar Calibration Sphere
The radar calibration sphere structural concept is
based on using a large number of hexagonal shaped
flat membrane panels that are bonded at their
perimeters to adjacent panels to form a sphere when
inflated, Figure 2. The demonstration hardware shown
in Figure 2 is about 6 meters in diameter. Flight
hardware would be metalized to accommodate high RF
reflectivity.
The “lenticular inflatable parabolic reflector”
concept consists of a lenticular reflector structure,
which is supported around its periphery by a toroidal
structure, Figure 3.
The reflector is made from a number of metalized
“pie shaped” membrane gores thai are bonded together
to form a parabolic shaped surface. The canopy
structure is identical to the reflector, but does not have
to have the same high surface precision or metalized
surface for RF reflectivity or solar energy collection.
The torus is constructed from a number of curved
segments that are bonded together to produce a “ring”
shaped structure. The inflated structure, shown in
Figure 3 is about 12 meters in diameter overall and the
reflector itself is about 10 meters in diameter.
Rigidization techniques for this type of inflatable
structure have been evaluated2, which include flexible
and rigid foams of different density and thickness.
2
Figure 4. Echo I Balloon
Figure 3. Lenticular Inflatable Parabolic Reflector
Echo Balloons
The flight of the Echo Balloons in the late 1950’s
and early 1960’s represented a relatively tmrecognized
technical capability for the development, fabrication
and launch of large, high precision inflatable space
structures. This original and very innovative
technology development was done at NASA Langley
Research Center (LaRC) starting in May 1958 and the
fmal project responsibility was assigned to NASA
Goddard Space Flight Center (GSFC) at a later date.
The objective of the balloon series was to provide
passive, space based communications reflectors.
Echo I was made from a large number of gores of
mylar that were 12 pm thick, coated with 2000
angstroms of vapor-deposited aluminum, and bonded
together to form a sphere, Figure 4. The lOO&.
diameter sphere weighed 136 lbs. and was stowed in a
26 in. diameter, spherical container for launch. Thirty
pounds of sublimating powers were sifted into the
balloon structure to enable orbital deployment. Four
inflation tests were made in 41-ft. diameter vacuum
chambers at LaRC and five free-space environmental
tests were made in ballistic flights with rockets from
N.4SA’s Wallops Island Test Station. These tests were
used to develop and validate mechanical packaS&
ejection and inflation techniques. which resulted in the
successful deployment of Echo I on orbit.
Echo I was successfully launched on August 12,
1960 aboard a Delta to an initial orbit of 1000 miles,
which changed over a period of months from the
effects of solar pressure. The Echo I was operational
for a number of months which indicated that it
maintained a sufficiently large and reflective profile for
this time period.
This original, innovative and significant activity
was very well documented’s4*‘, with additional
references therein.
Contraves Inflatable Structures
The European Space Agency’s (ESA) interest in
the potential of inflatable deployable space structures
was signified by their sponsorship of the development
of reflector antenna and sun shade structural concepts
at Contraves Space Division in Switzerland. The
technology focus was for axisymmetric reflector
antennas for Very Large Baseline Interferometry
(VLBI), offset reflectors for mobile commtmications
and sun shade support structures for telescopes and
large sensors. This significant development was
initiated by ESA-ESTAC in the late 1970’s and
demonstrated in the late 1980’s. A 6-meter diameter
reflector antenna which was a 1/3 scale model of a
VLBI antenna, was built in the early 1980’s and
evaluated6. Subsequent to this, a IO x 12 meter offset
reflector antenna for land mobile communications at Lband
was built and evaluated for surface precision and
other mechanical characteristics’, F@re 5. The
measured reflector precision inflated, but not rigidized
was on the order of a few mm*s rms, which was quite
good for a structure that size. The construction of the
antennas was based on using two parabolic
membranes, supported at their periphery by a toroidal
Figure 5. 10 x 12 Meter Reflector Antenna Structure
structure. The membranes were made from multiple,
“pie shaped” gores, one being RF transparent and the
other metalized with aluminum to enable RF
reflectivity. The load carrying fibers in the gores were
Kevlar and matrix material was designed to become
rigid on orbit Tom solar heating, after deployment by
inflation.
A new structural concept for a sun shade support
structure for a submillimeter space telescope was
developed to the point of a functional scale model’,
Figure 6. This structure was based on a truss type
snucture that utilized the same materials and
rigidization techniques that were developed for
reflector antenna structures. This support structure was
intended to use flexible panels such as ML1 blankets or
equivalent in each bay. This would enable compact
mechanical packaging to be achieved below the
telescope structure in annular configuration, that is
smaller in diameter than the telescope reflector
structure.
These significant technology developments were
very well documented6*‘*’ with additional references
therein.
-
Inflatable Antenna Experiment .
NASA’s interest in demonstrating the potential of
this relatively new class of space suucture resulted in
their sponsoring the IN-STEP Inflatable Antenna
Experiment’*“, which flew on STS-77 on May 29,
1996, Figure 7. The antenna structural concept used
was developed by L’Garde, Inc. who have been
designing, manufacturing, ground and flight testing
inflatable space structures for the past 25 years. The
experiment objectives were to a) verify that large
inflatable space structures can be built at low cost, b)
show that large inflatable space structures have high
mechanical packaging efficiency, c) demonstrate that
this new class of space structure has high deployment
reliability, d) verify that large membrane reflectors can
be manufactured with surface precision of a few
millimeters rms, and e) measure the reflector surface
precision on orbit .
’
_
The inflatable structure comprised two basic
elements, the inflatable reflector assembly and the
Figure 6. Telescope Sun Shade Support Structure Figure 7. IAE on Orbit
torus/strut supporting structure, Figure 8. The reflector
assembly formed a 14 meter off-axis parabolic aperture
with a f/d of 0.5. The surface accuracy goal was I.0
mm rms as compared to a best fit parabola. The
reflector film, ‘A mil aluminized mylar, was stressed to
approximately 1200 psi by the inflation pressure of 3 x
IO4 psi. This stress level was sufficiently high to
assure a good reflective surface for the accuracy
measurement system. The canopy was also
constructed with 62 gores of % mil mylar but was left
transparent. The torus/strut structures were 24 and 18
inches in diameter, respectively, and were made with
12 mil thick neoprane coated Kevlar and locate the
reflector assembly at the effective center of curvature
of the reflector parabola as required for operation of
the Surface Accuracy Measurement Subsystem. The
torus also provides the rim support for the reflector
assembly without which the reflector assembly, when
inflated, will take a spherical shape.
The experiment was successfully flown on the
recoverable Spartan Spacecraft. A new, unique and
low cost space structures technology was demonstrated
on orbit by a) building a large inflatable space antenna
structure for on the order of $l,OOO,OOO, b)
demonseating extremely efficient mechanical
packaging by stowing a 14 by 28 meter inflatable
structure in a container the size of an office desk, c)
manufacturing an offset membrane reflector structure
with a surface precision on the order of a few mm’s
rms, and d) demonstrating the robusmess of
deployment for this new class of structure. The results
of this experiment were used specifically to establish
the technology data base and were the basis of a
technology road map for the continued development of
this type of space structure.
Figure 8. Reflector Structure
The results of this experiment are very well
documented9.” with many additional references
therein.
CURRENT INFLATABLE STRUCTURES
TECHNOLOGY DATA BASE
The current data base for space inflatable
structures comes from both ground testing and flight
experience. The vast majority of the flight experience
is a result of the Air Force decoy programs at L’Garde,
Inc., plus the NASA/JPL/L’Garde Inflatable Antenna
Experiment9.ro. An overview of membrane materials,
rigidization techniques, deployment methods, inflation
techniques, and finally manufacturing and assembly of
large inflatable structures is provided.
Membrane Materials Characteristics/Orbital Radiation
Use of membrane materials in inflatable
structures is primarily .for the lenticular structure used
by parabolic reflectors”. These reflectors fmd
application in solar concentrators and radio frequency
(RF) antennas and others. The reflector portion of the
lenticular is normally metallized to reflect and focus
the solar or RF enera while the canopy which forms
the other half of the pressure vessel must be transparent
to the wavelength of interest. The material thickness
for these applications is normally on the order of % to
1 mil with the operating stress level ranging from 100
psi to 3000 psi, depending on the application. The
lower the stress level, the lighter the support structure
and the lower the make up gas weight. Usual
applications for these types of structures require a
lifetime of 5 to 10 years during which time the
structure must be able to retain its integrity, shape and
surface accuracy. Figure 9 lists the most important
properties of membrane materials and the current most
promising candidate materials along with their
corresponding properties”. Of the materials listed in
Figure 9, only the Kaptons are readily available in
production quantities and in the desired thicknesses.
The remainder are in various stages of development.
The damage threshold levels for ionizing
radiation are not available for several of the materials
in Figure 9. However, the results of tests for the
Vacuum Ultraviolet (VUV) and particle radiation
levels expected at Geosynchronous Earth Orbit (GEO)
and a typical Low Earth Orbit (LEO) orbiti indicated
that Kapton E, Aorimide, and CP2 performed well.
Atomic Oxygen (AO) reaction efficiencies also are not
available for many of the materials, however, tests with
exposure levels expected at LEO show Kapton E
5
Property Kapton H Kapton V Kapton E Aorimide
(DuPont) (DuPont) (DuPont) (Triton)
PBO
(Fast.
Milr.)
CPU2
(SRS)
Coefficient of Thermal
Expansion
PPM/C
(Yellow, TOR)
@-l?38C @SO-$OC @O!;OOC a-75 z2
MD -7.6 47tos1
TD +7.6
2ooc
Shrinkage
Coefficient of Hygroscopic
Expansion
%
PPM/%RH
0.17 0.03 0.03 NA NA NA
22 17 9 NA 0.8 NA
H20 Absorption % I .8 to 2.8 1.8 to3 2.4 2 to 8 0.8 NA
%SORH@23C
Modulus KPS.1 370 400 7.50 450 MD 6000 315 to 420
TD 3000
Yield Strength TD PSI 10000 10000 15000 8800 27500 NA
MD 9600
Creep (Total strain % (@applied NA NA O.O065(3OOpsi) NA 0.0055 NA
after 76 days) stress) (I SOOpsi)
Solvent Resistance excellent excellent excellent excellent excellent sol. in MEK
MIBK,
CHC13
Uniformity
(thickness), Mils
rmsx100 NA NA 2.4-2.5 2..7-11.7 15.9 10
Space Env. A0 Re(cc/AO)xl O”-24 3 3 0.14 0.6 NA
VUVIAO Re(cc/AO)xl O”-24 3.07 3.07 0.17 NA NA
W V % Prop. Retained loo(Ts)@mooHIs loocrs)@loooHn EXCEL@4SWES NA NA
Ionizing Rad. Rad Thresh, Rad, TS 5xlOA9 5x109 NA H NA NA
Rad Thresh, Rad, %E 1xlW 1x109 NA NA NA NA
NA
Outgassing CVCM
TML
Bondability
Metallizability
%
%
0.02 0.02 NA 4 NA NA
0.77 0.77 NA NA NA
Yes Yes Yes Yes Yes Yes
Yes Yes Yes Yes Yes Yes
Figure 9. Membrane Materials Properties
performs as well as Kapton H and that PBO, as well as
CPI, performed satisfactorily. The Aorimide mass loss
was roughly 20% of the other materials. Evaluation
testing of these materials is continuing in order to
obtain values for the remainder of the properties shown
in Figure 9.
Rigidization Techmoues
The only practical applications of purely
inflatable space structures are for reflector and
concentrator structures. These structures are made of
light-weight thin films and, therefore, are very lightly
loaded. They are usually operated at very low
pressures - on the order of lW6 psi. It is therefore
possible to provide make-up gas to account for the
losses due to leakage through punctures caused by
space debris and micro meteoroids. This is true
especially for lifetimes of 5 or 10 years. All other
inflatable space structures would need to operate at
considerably higher pressures in order to carry the
much greater applied loads. So, unless their lifetime is
extremely short, it is necessary to rigidize them after
deployment.
The ideal rigidization system would exhibit the
following properties: a) high modulus after rigidization
for suuctural stifmess, b) process reversibility for
testability, c) high flexibility for dense packaging, d)
zero coefficient of thermal expansion for thermal
6
stability, and e) resistance to the space environment,
and f) minimal change of shape during the rigidization
process.
There are many rigidization techniques that have
been developed to date. The most common of these
are: a) fabric impregnated with resin that is cured by
exposure to ultraviolet light, b) fabric impregnated
with water soluble resin that rigidizes as the water
evaporates, c) fabric impregnated with a resin that
rigidizes when it is cooled below its glass transition
temperature, d) thermal set plastic resin that cures upon
the application of heat, e) a laminate of aluminum foil
and thin Kapton film that rigidizes when the aluminum
is strained beyond its yield point. Examples of these
are shown in Figures 10 through 14. While all of these
rigidization systems meet some of the requirements,
none meet all of them.
” _ _- -~-- _ .
. .
Figure 12. Carbon Fabric Impregnated with
Low Temperature Rigidized Resin
Figure 10. W Cured, Gloss Fabric Impregnated
with W Cured Resin
Fi_gure 13. Carbon Fabric Impregnated with
Thermally Activated Epoxy
P
F@re 11. Kevlar Fabric Impregnated
with Water Soluble Resin
Figure 14. Kapton, Aluminum Foil, Kapton Laminate
7
With proper selection of a structural fabric, all of
the resin/fabric systems can be designed for a modulus
greater than IO x IO6 psi, which is also about the value
of the aluminum laminate. The UV cured and
thermally cured thermoset plastic systems have the
disadvantage of being non-reversible and, therefore, it
is not possible to test the rigidized configuration that is
to be flown. The aluminum laminate system is not
reversible; however, it can be rigidized for ground
testing, then repackaged for flight and rigidization in
space. All of the systems can be packaged to volumes
much smaller than their deployed dimensions. The
fabric based systems can be packaged somewhat more
densely than the aluminum laminate. None of the
current systems have a zero coefficient of thermal
expansion, therefore, it is necessary to provide
Multilayer Insulation (MLI) to reduce the temperature
gradients sufficiently to prevent warping due to nonuniform
heating. This is a major disadvantage for the
UV cured systems, since they require exposure to the
sun to cure. A disadvantage of the water soluble resin
system is that as the water evaporates for rigidization,
it could possibly condense on nearby surfaces, but only
if they are extiemely cold, such as the case for infiared
telescope mirrors. All of th.e systems developed to date
have exhibited resistance to the space environment.
Deplovment Technioues
The current deployment techniques are directed at
members such as tubes and struts. Normally these
members are used to move the remainder of the
inflatable system into position for inflation. For
example, the struts on an antenna would be used to
move the torus and lenticular into position where they
would simply be inflated. Deployment methods must
keep the deploying inflatable structure within a
predictable envelope, and provide a well-defined
deployment rate that is slow enough to prevent
significant loads on the spacecraft. They must also
provide restraint for the structure during launch, as
well as large well-defined passages to vent entrapped
gas during ascent to orbit. There are several techniques
available to satisfy these requirements, three of which
are shown in Figures 15 through 17.
Roll-Out Method
This method is similar to the well known party
favor. The primary difference is that the coiled spring
that provides the deployment resistance in the party
favor is replaced with Velcro@ on the top and bottom
of the tube. By varying the area and location of the
Velcro@ it is possible to vary the pressure necessary to
Figure 15. Roll-Out Method
Outer
Depioymg
, Tube
, Tube
Figure 16. Fan Fold Method
Figure 17. Mandrel Method
deploy the tube, hence, its rigidity during deployment.
Depending on the flexibility of the strut material, it
may be necessary to provide a method to assure the
vent path.
Mandrel Method
The Mandrel method shown in Figure 16 features
well defined venting paths to the packaged strut. The
strut is packaged beneath a mandrel with a conical
shaped lead in. Application of pressure causes the
packaged tube to be pulled over the mandrel. The
mandrel provides the directional control of the
deploying strut. In addition, friction between the
mandrel and interior surface of the tube provides the
resistance to deployment, which in turn controls the
rate of deployment and strut stiffhess during
deployment.
Fan Folded Method
mis method was used for the ground test space
chamber tests of the ITSAT solar arrayi4. Directional
control for this example is provided by folding the
rigidizable tube 90’ to the solar array folds. The
deployment resistance is provided by the bending
strength of the tube itself which, in this case, is made
of Kapton-Aluminum-Kapton laminate. Because of the
stiffness of the tube material, well defmed vent paths
were inherent in the folded tubes.
Inflation Subsvstem
Numerous types of inflation systems have been
used for inflatable space structures. The most used are
systems using nitrogen gas and subliming powders.
High vapor pressure fluids have been used where the
pressures from sublimating powders are not adequate
and using nitrogen is impractical. Hydrazine systems
are now being evaluated because of their capability to
give medium level inflation pressures with low weight
and volume when compared to nitrogen systems.
Hydrazine application issues are handling, safety and
cost.
Nitrogen systems are used where moderate
hardware cost and low development cost are of prime
importance. In the case of IAE9v”, component hardware
common to the Spartan was the majority portion of the
system since it was ah-eady Shuttle qualified. Weight
and space optimization was not a primary requirement
for IAE, whereas cost, reIiability, qualification, and
availability were very important considerations.
The key design drivers for the IAE inflation
subsystem included a) high-pressure nitrogen gas
storage for the inflatable structure, b) sensors, valves,
and reguIators for implementing the control of
inflation, c) using a functional concept based on
previous successful L’Garde, Inc., designs, and d)
maximizing the use of Spartan cold-gas attitude
control-system components.
The functional design of the nitrogen inflation
subsystem is nearly identical in concept to the ones
successfully flown by L’Garde, Inc., for much smaller
inflatable structures. Analysis of mass flow was used to
establish component requirements. Component
selection was based on previously qualified hardware
used for the Spartan attitude-control, cold-gas system
and on previous L,‘Garde, Inc., flight systems. The
supporting structure used for mounting the tanks,
plumbing, and components was an ahrminum
honeycomb panel similar to that used for the canister.
The two large structural composite gas tanks utilize the
same mounting configuration as that for Spartan to
minimize re-qualification costs. No attempt was made
to develop a light-weight, highly compact inflation
system for this experiment because of cost limitations.
Sublimating inflation systems have been used
since the fust orbital test of the Echo balloons. The
operating principle for sublimating powders is to
release the powder within the interior of the inflatable
structure after orbit insertion. In space conditions the
powders will sublime into a gas that provides vapor
pressures in the range of l@’ to 104 atmospheres,
depending on the gas temperature within the inflatable.
Temperature control of the balloon interior is
maintained through proper thermal design of the
balloon system. These powders provide self pressure
regulation if excess powder is carried and allowed to
sublimate as makeup gas. Sublimating systems have
the advantage of handling ease by being non-corrosive
and solid at room temperature. Their toxicity is
reasonably low and they are low cost. L’Garde has
used these systems in flight tests of target balloons
launched from sounding rockets.
Manufacturing and Assembly Methods
Inflatable structures require unique
manufacturing methods and techniques because of the
thin flexible materials that are used. For the IAE onefourth
mil (.00025 inches) polyester films were used
for the lenticular structure. For the torus and struts
eleven mil rubberized Kevla& fabric was used. The
handling and cmting of gores from these materials
present unique challenges, especially because of the
dimensional precision required to attain overall system
performance. Precision cutting of gores from thin
films is a very important factor in the making of an
accurate reflector or structure such as those used for
the IAE. Precision templates are often used for cutting
of the gores for smaller reflectors/concentrators. This
method gives a hia degree of gore dimensional
control, but the templates become very expensive as
the size of the reflector or structure increases. There is
a practical limit on how large the templates can be and
still give reasonable handling and cutting results. This
9
method is also labor intensive. A more desirable
method is to make use of an automated cutting system
that can produce accurately cut gores with greater
precision and lower cost. L’Garde’s computer
controlled gore cutting machine is show in Figure IS.
This equipment has the capability to produce gores for
reflectors up to 25 meters, and larger with the addition
of table segments.
The gores are seamed together using tape and
space qualified adhesive. The finished membrane is
then mounted on a fixture that allows the membrane to
be pressurized for accuracy measurement. This is
accomplished using photogrammetry.
Photogrammeny, as its name implies, is a
3-dimensional coordinate measuring technique that
uses photographs as the fundamental medium for
metrolo* (measurement). The basic principle used is
triangulation. By taking photographs (or videographs)
from at least two different locations, so-called “lines of
sight” can be developed from each camera to points on
the object under test. These lines of sight (sometimes
called rays owing to their optical nature) are
mathematically intersected to produce 3-dimensional
coordinates of the points of interest. The technique of
photogrammetry is used in the V-STARS system by
Geodetic Services, Inc. of Melbourne, Florida. In the
VSTARS, the film camera is replaced by a high
resolution digital camera and the pictures - video
images in this case are stored in the computer. The
system includes user-friendly software that carries out
all the laborious calculations to determine the
coordinate points of interest from the video images of
the same article. The accuracy of the VSTARS is about
0.001 inch for every 100 inches.
Supporting structures, such as struts or torii, are
generally constructed using similar techniques, except
the materials are generally thicker and the accuracy
requirements a little less stringent.
Figure 18. Automated Gore Cutting Machine
ANALYTICAL CHARACTERIZATION OF
MEMBRANE REFLECTORS
The analytical characterization of mechanical
performance of inflatable deployable space structures
involves many different possible approaches depending
on type of the structure involved. The most
challenging is the characterization of the pressurized
membrane reflector structures, in particular the
reflector surface precision as a function of the
parameters that contribute to the error. A review of
previous analytical work on the shape of inflated
membrane reflectors has been done and new capability
has been developed and experimentally verified.
Additionally, these new tools have been used to
determine the sensitivity of reflector error sources and
to characterize the change of reflector shape as a
function of edge displacement.
Reflector Shape Analvsis
Interest in the analysis and design of pressurized
fihn structures can be traced back to the 19lO’.s, when
the power series solution of a homogeneous isotropic
linear elastic circular membrane under lateral pressure
--- Hen&y’s problem” --- was published. This paper
and subsequent contributions (which include advances
of Hencky’s approach, variational solutions based on
assumed pressurized shapes, and other theoretical
works) are generally restricted to an initially flat
circular geometry and typically render the problem
solvable via approximations on slopes and rotations.
These approximate solutions remain of interest today,
even in the age of advanced numerical solutions with
fmite difference and fmite element methods. The
reasons for this include that the numerical solution of
membrane problems can be very fragile and that the
verification of the results is difficult. Available
pressurized membrane test results are limited and
original tests are costly and delicate. In this scenario of
several solution options with mixed advantages and
disadvantages, it is important to be aware of the errors
associated with different approaches. Murphyi tended
to this need in 1987 when he compared the accuracy of
a number of solutions in the context of solar
concentrators. His study, however, is limited in a
number of respects. First, it is restricted to initially flat
membranes. Second, the accuracy requirements
addressed are those for heliostats: error tolerances for
the high precision RF reflectors currently considered
are much more restrictive. Third, Murphy measures
accuracy via slope, rather than wavefront errors, while
for RF application the latter is relevant. Finally,
10
analytical solutions based only on approximate
variational formulations are included.
-
Precision space reflector applications are
associated with accuracy beyond the customary
tolerances of structural engineering. This makes the
reliability of related analytical predictions critical
whenever test verification is difficult, as is the case for
large inflatable membrane reflectors. For pressurized
membranes, however, numerical analysis can be
troublesome and classic solutions exist only for the
simplest configurations. Furthermore, the accuracy of
classical solution options is not well understood due to
approximations made to facilitate the exact solutions.
Research” was conducted which contributes to filling
this gap by exploring the impact of representative
solution approximations on the accuracy of analytical
shape predictions. The selected approximations were
individually addressed via a parametric study of
axisymmetric linear elastic isotropic membranes.
Limits of applicability were considered for
pressurization levels and accuracy requirements of
current professional interest for radio frequency (RF)
applications. Initially flat and curved membranes were
studied; the latter designed to assume exact parabolic
shapes when pressurized. Although specific diameters
were studied, the results are applicable via newly
developed scaling laws to any dish diameter. A
sofhvare package, AM (Axisymrnetric Membrane)“, is
a high precision numerical tool for the study of
pressurized axisymmetric membranes capable of
modeling wrinkling and of determining initial shapes
which inflate to desired pressurized contours. In the
development of AM, both large rotations and
displacements are included in the analytical
formulation. While the guideline tolerance for the
parametric study” reflects current interest in RF
applications, AM surpasses that accuracy with a
precision applicable to optical frequencies. Results
from the AM study were compared with the FEM
programs FAIM”, and NASTIUN, for initially flat
membranes and correlation was obtained well within
the range of accuracy requirements for RF
applications.
To provide benchmark solutions and to serve as a
testbed for parametric studies, the software package
AM for the high precision analysis of pressurized
axisymmetric membranes was exercised on several
problems for demonstration purposes. AM is written
in C and it consists of a number of modules which are
easily combined into special purpose or general
programs. These modules include: a) numeric shape
solver, to calculate via direct integration the loaded
state (shape, stresses, and strains) for axisymmetric
membranes of arbitrary initial shapes subject to
pressure, thermal, and kinematic (edge displacement)
loads. The modeling of wrinkling, as well as
simulating various analytical approximations are
optional; b) symbolic solver, for classical
axisymmetric membrane solutions. AM currently
includes Hencky’s solution, as well as other
approximate solutions; c) inverse solver, to provide
the initial shape for a desired pressurized contour and
pressure; d) shape comparator, to compare
axisymmetric shapes and evaluate the error between
via any of a number of error measures; e) shape
optimizer, to adjust a given shape (via axial translation
and/or scaling) to achieve a best fit to a particular
surface; and fl parabolic assessor, to produce the bestfit
paraboloid to a given membrane shape and evaluate
the (RF rms or other) error between. The numeric
shape solver at the heart of the numeric solver module
is an integrator to solve the meridian as an initial value
problem from any known point onward. In each
integration step along a discretized contour, the state
and position of a point is iteratively determined from
the known adjacent point. The point iteration continues
until, the equilibrium, constitutive, and kinematic
equations are satisfied. Complete details of the
program and solution methods are presented in
detail*‘,i8.
Reflector Error Sources
As discussed earlier, the dimensional
requirements for precision space reflectors involve
accuracy beyond the customary tolerances of structural
engineering. Thus, to achieve a high precision
reflector, consideration must be given to all possible
error sources. A listing of possible error sources is as
follows:
.
.
.
.
.
.
.
.
.
.
.
to
Material stiffness properties and area1 variation
Material thickness and area1 variation
Creep
Moisture effects
Material “wrinkling” or creasing due to handling
and packaging
Fabrication
Analytical shape prediction
Edge support conditions
Pressure level
Thermal distortions
Gravitational effects in earth testing.
A primary purpose of the analytical work”,” was
understand the relative accuracy of available
11
analytical tools in predicting the shape of thin film
structures. The results of that study indicate that the
commonly used tools for shape analysis, FAIM and
NASTR4N can readily provide acceptable accuracy
for RF class precision reflectors. The application of
these tools to higher precision applications such as
optics will require further careful study and evaluation.
Thus, it would seem that observed errors in ground
tests are the result of the other sources listed above
The major advantage of inflatable thin film
reflectors over mechanically deployed structures lies in
the fact that these structures can be packaged into
extremely small volumes for launch. This advantage
increases for very large diameter reflectors; diameters
in the 25 to 50 meter range. Additionally, for these
reflectors to be practical, the internal pressure must be
kept very low to minimize the amount of makeup gas
required for leakage through micrometeoroid
punctures. This in turn requires that the operational
stresses in the thin films be very low. On the other
hand, it would be desirable for the elastic deformations
due to pressure to be large compared with fabrication
errors so that the reflector would achieve its desired
analytically predicted shape under load. This then
leads to the observation that it would be highly
desirable to have thin films with a very low modulus.
Current space qualifiable thin polymeric films have a
modulus on the order of 500,000 to 800,000 psi. A
challenge for the materials community would be to
develop thin polymeric films with an order-ofmagnitude
lower modulus than currently available.
This would significantly enhance the feasibility of
reliably achieving high precision inflatable reflectors.
Another major factor in inflatable reflector accuracy is
the thermal distortions. For precision RF reflectors, it
would be desirable to have a coefficient of thermal
expansion less than the current value of 12 ppm per
degree F that exists for KAPTON E.
To provide confidence that these large diameter
reflectors will perform as expected in space will
require the development of a ground validation
program. This ground validation program will involve
a systematic protocol consisting of carefully conducted
tests coupled with high accuracy simulation analyses to
properly bridge the gap between ground tests and space
operation. It is likely that gronnd tests on 25 to 50
meter diameter reflectors will be impractical and
possibly impossible. For example. if desired materials
are developed to enable reflector operation at
extremely low pressures, the effects of gravity could be
greater than that of the pressure. In fact, for large
diameter reflectors, the weight per unit area of the film
may be greater than the internal pressure. For such
cases it would be impossible to provide test verification
through earth based tests. However, it is possible to
tind a smaller diameter reflector for which this pressure
anomaly is not an issue. A set of constant thickness
scaling laws have been developed by the authors which
permits the results of the small scale tests to be scaled
up to larger diameters in a rational fashion. Such
small-scale tests along with the scaling laws will play a
critical role in the development of a ground based
verification protocol for large inflatable reflectors.
One-Meter Tests - A series of tests on
pressurized flat membranes one-meter in diameter as
shown in Figure 19 were recently conducted as part of
a program to validate analytical tools and to understand
the roll of small scale testing in verifying antenna
performance. The test procedure, metrology system,
and results are presented”. As part of the program,
careful tests were conducted on the constituent
materials to establish the elastic properties to be used in
the analysis for correlation. Because of availability,
the materials selected for testing were 1/2-mil MYLAR
and KAPTON HN. Two materials were used to better
understand the roll of material properties. During the
modulus testing of the MYLAR, highly anisotropic
properties were observed as shown in Figure 20. The
material was not only orthotropic but the principal
material axes were found to occur at an angle to the
machine and transverse directions of the material.
Subsequent discussions with DuPont revealed that this
anisotopy is a function of the manufacturing process
and could vary among batches. Thus, the use of this
material for structural testing must be accompanied by
stiffness testing of the batch used. The KAPTON
material tested was very isotropic and a completely
different manufacturing process explains this.
Figure 19. One-Meter Diameter, % Mil Mylar, Flat
Circular Test Specimen
12
I
Figure 20. Young Modulus for % Mil Mylar as a
Function of Angular Orientation
The correlation between test results and the fmite
element analysis FAIM are shown in Figure 21 for the
MYLAR specimens. The results are shown for two
different pressure levels corresponding to maximum
film stresses of 636 psi and 3076 psi. These two
widely different fihn stresses were chosen to provide
correlation with analysis over a large ranSe of loading.
It should be noted that the analysis conducted in this
study was nonaxisymmetric due to the anisotropic
nature of the material. As can be seen in Figure 21 the
global correlation between test and analysis appears
excellent. However, due to the extreme nature of the
accuracy requirements for precision reflectors, this
1 -meter Diameter hfybzr Membrane
f.60 , I
- --
-.
-=.
p = .018 pi
FAIM-calculated
b
\
\
b
-ti \
p = .0017 psi .
.
.
l\
0.00 II.,,,, ,,:I/,,,,,,,,,,,,~,,,,~~,,,,,,,
0 2 4 6 8 f0 72 f4 t6 18 2 0
Radial Distance (inch)
F@re 21. Measure Mylar Profile from Figure 19 Figure 22. Deviation in Membrane Later
Test Article. Membrane Center Stress Equals Displacement Between Finite Element Code
636 psi and 3076 psi FAIM and Experiment of Fi_gne 19
Slobal look at correlation is not adequate. To enable a
finer look at the correlation, the differences between
test and analysis are plotted in Figure 22. The
maximum difference between test results and analysis
is for the high stress case a with a value of 0.006 in (6-
mils). This difference amounts to about OS”h of the
total elastic deformation. This range of error is within
the level of accuracy of the materials properties
(modulus and thickness variations). It is concluded
that modem numerical analysis tools are able to predict
membrane deformations under ranges of loading of
interest in realistic designs. Thus, the majority of
errors that occur in thin film reflectors will be a result
of the other errors discussed previously. Since it may
be practically impossible to eliminate all of these error
sources, it may be prudent or even necessary to
consider on-orbit shape adjustment as a means of
obtaining and ensuring required shape accuracy.
Potential shane adiustment approaches - The
major approaches for providing shape adjustment of a
pressurized membrane reflector are as follows:
. Adjust feed position
. Adjust edge radial displacements
l Change internal pressure
. Provide thermal gradient over reflector surface
. Integrate electro-piezoelectric into the thin film
surface.
To evaluate the feasibility of these different shape
adjustment approaches it is necessary to be able to
conduct reliable shape change simulation analyses.
The research” and validation” have demonstrated that
such simulations can be carried out with existing
structural analysis tools. Each of these shape
13
adjustment approaches are being studied by different
researchers as to their effectiveness and practicality.
Of the above mentioned shape adjustment
potential approaches, the concept of adjusting the edge
radial displacement is particularly attractive due to the
relative implementation simplicity of the concept.
Results” demonstrate the potential shape changes that
can result from the application of radial edge
displacement. The results of that study are shown in
Figure 23. In this figure the shape deviations from a
perfect paraboloid are plotted as a function of radial
distance from the center for three different t7Ds. In
order to provide shape control of a reflector surface by
changing the radial edge displacement, it would be
necessary for the resulting effect to be transmitted
throughout the surface. As can be seen in Figure 23,
the effect of an applied edge displacement is a strong
function of the reflector curvature (VD). For shallow
reflectors (f’D = 2), the effects of an applied radial
edge displacement are dramatic and provide a change
over the entire surface. However, for deep reflectors,
(fD = 0.5) the effect of the edge displacement results
primarily in a 2 translation of the surface with
nonuniform effects occurring only in a botmdary layer
near the edge. The implications of this study are that
radial edge control will not be effective for deep
reflectors, while some control can be achieved for
shallow reflectors. Since most reflector applications
involve fD values greater than 1, the approach of
using edge displacements for shape control will not be
possible.
STRUCTURES/ENVIRONMENTAL
INTERACTIONS
Of all the known types of space structure at this
time, the inflatable structures have the most significant
interaction with the space environment. This is a
consequence of the materials used, the large size
structures that are needed for many applications, the
rigidization techniques used and, in particular, the need
to maintain inflation pressure in some of the structural
elements such as the reflector/canopy structure.
A major challenge for space inflatable structures
at this time is the development of flexible materials for
both inflated and rigidized membranes. Thin
membrane materials are usually used for high precision
reflector structures that are tmder constant pressure
loading and planar array structures that are under
constant in plane tension loading. In addition to being
resistant to the orbital radiation environment, these
membrane materials usually have stringent
, I
I I
arz t-l
1 -_ J7D = 2.0 1
60 h z#-p= -93.7! kk 40
\TiY
- . .
0 O S 1.0 1.5
Figure 23. Effect of In-Edge Displacement on
Deflected Shape
requirements for mechanical properties such as ultralow
modulus, low long-term creep, while at the same
time lending themselves to handling, processing,
metalizing, bonding and high density mechanical
packaging. Rigidizable flexible inflatable materials
also have stringent requirements for mechanical
properties such as high-deployed stiffness, low thermal
expansion, and low long-term creep while operating in
high radiation environments such as orbiting the moons
of Jupiter.
14
These interactions and others such as the effects
of solar pressure and micrometeoroid penetration must
be accounted for in the design, materials selection,
structural configuration, and orbital scenarios for all
applications under consideration for this new class of
space structure.
CONCLUSION
The potential of inflatable, deployable space
structures for enabling some specific classes of
application seems to have been recognized for about 50
years. At that time innovative and unique technology
developments were initiated. Such developments
continued at a relatively low level of investment until
the advent of the IAE in 1996. The results of that
experiment provided a) a technology data base for a
reflector antenna structural concept, b) validation of
the potential for this new class of space high precision
structures and c) illumination of the specific
technologies needed to enable such a concept and
project the potential performance for specific
applications. Subsequent to 1996 a number of
different technology developments were implemented.
In particular, a) the analytical characterization of
inflatable structures mechanical performance has been
addressed and the results are significant, b) the
experimental characterization of the effects of orbital
radiation on a number of different thin film materials
has been initiated and c) the concept development for
rigidization techniques, fabrication of seamless
membrane reflectors and others at a number of
different manufacturing organizations. Collectively,
the current technology data base for specific structural
concepts, structural designs that account for
environmental interactions and the new analytical tools
for, projecting orbital performance will enable realistic
estimates of the applicability of inflatable structures for
specific applications.
ACKNOWLEDGMENTS
The work described in this paper was carried out
at the Jet Propulsion Laboratory, California Institute of
Technology, under a contract with the National
Aeronautics and Space Administration.
The authors would like to acknowledge and thank
Richard M. Dickinson of Jet Propulsion Laboratory for
providing technical data, information, photographs and
references for the inflatable structures developments
-
accomplished by Goodyear.
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
REFERENCES
Dickinson, Richard M. and Cramer, P.W. Jr., ‘&A
Survey of Unfurlable Spacecraft Antennas”
presented at the IEEE Antenna and Propagation
Society Meeting, Boston, MA 1966.
Harris, J. T., and Stimler, F. J., “Expandable
Structures for Space”, Astronautics, Volume 6,
No. 4, pp. 30,3 1, 92-94, April 196 1.
Jaffe, Leonard, ‘Project Echo Results,”
Astronautics, May 196 1.
Wilson, A., ‘<A History of Balloon Satellites”,
British Interplanetary Society, Volume 34, pp. IO22,
Jan. 1981.
Hansen, James, “The Big Balloon”, Air & Space,
April/May 1994.
Reibaldi, G.G. and Bemasconi, M.C. (1985)
“QUASAT Progrannne: The ESA Reflector”
Paper IAF-85-400 presented at the 36*
International Astronautical Congress, Stockholm,
Oct. 7-12; also: Acta Astronautica 15 (1987), 181-
187.
Bemasconi, M.C., E. Pagana and Reibaldi, G.G.
(1987) “Large Inflatable, Space-Rigidized
Antenna Reflectors: Land Mobile Services
Development” Paper IAF-87-3 15 presented at 38”’
International Astronautical Congress, Brighton
(UK), October 1 O- 17.
Bemasconi, M.C. and Koese, S. (1988) “The
Space-Rigidized Thermal Shield for the ESA FarInfrared
Space Telescope (FIRST)” Paper
presented at the 3rd European Symposium on
Space Thermal Control and Life Support Systems,
Noordwijk (NL), October 3-6, ESA SP-288, 165-
173.
Freeland, R. E., Bilyeu, G.D., Veal, G.R., Steiner,
M.D. and Carson, D.E., “Large Inflatable
Deployable Anterma Flight Experiment Results,”
IAF Paper 97-1.3.01, presented at the 48*
Congress of the International Astronautical
Federation, Turin, Italy, October 6-10, 1997.
Freeland, R.E. and Veal, G.R., “Significance of
the Inflatable Antenna Experiment Technology”
presented at the 39*
AIAAIASMEIASCEIAHS~ASC Structures,
Structural Dynamics and Materials Conference,
AIAAfASMElAHS Adaptive Structures Forum,
.4pril20-23, 1998, Long Beach, California.
Jenkins, C.H., Freehand, R.E., Bishop, J.A., and
Sadeh, W.Z. An Up-to-Date Review of Inflatable
Structures Technolo=g for Space-Based
Applications,” presented at the Space 98
Conference, Albuqueque, NM, April 27, 1998.
15
12. Stuckey, WK., Meshishnek, M.J., Hanna, W.D.,
and Ross, F.D., Space Environment Test of
Materials for Inflatable Structures, Aerospace
Corporation Report No. TR-98( 1055)-I, dated 16
April, 1998.
13. Space Environmental Effects on Spacecraft: LEO
Materials Selection Guide, NASA Technical
Report #466 1, August 1995, Two Volumes
14. Malone, P.K., Williams, G.T., “Lighmeight
Inflatable Solar Array”, Journal of Propulsion and
Power, Vol. 12, No. 5, September-October 1996.
15. Henky, H. Uber den Spannungszustand in
kreisrunden platten mit verschwindender
Biegmgsteifigkeit. Zeitschrifl fur Mathematik
und Physik, 63:311-317, 1915.
16. Murphy, L.M. Moderate axisymmetric
deformations of Optical Membrane Surfaces.
Journal of Solar Energy Engineering, Transactions
ofthe ASME, 109(2):111-120, May 1987.
17. Greschik, G., Mikulas, M.M., Palisoc, A.
Approximations and Errors in Pressurized
Asisymmetric Membrane Shape Predictions. 39*
AIAA SDM Conference Proceedings, Long
Beach, CA, April 20-23, 1998. AIAA-98-21 OlCP.
18. Greschik, G., Palisoc, A., Cassapakis, C., Veal, G.,
Mikulas, M.M. Approximating Paraboloids with
Axisymmetric Pressurized Membranes. 39ti
AIAA SDM Conference Proceedings, Long
Beach, CA, April 20-23, 1998. AIAA-98-2102-
CP.
19. Palisco, A., Veal, G., Cassipakis, C., Greschik, G.,
Mikulas, Martin. Geometry Attained by
Pressurized Membranes. SPIE Paper Number
3356 46, February, 1998.
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