DEVELOPMENT OF FLIGHT HARDWARE FOR A
LARGE, INFLATABLE-DEPLOYABLE ANTENNA EXPERIMENT
R. E. Freeland.
Jet Propulsion Laboratory
California Institute of Technology
Pasadena, California
G. D. Bilyeu**
G. R. Veal+
L’Garde, Incorporated
Tustin, California
Abstract
Large, space-based antennas are needed for a
variety of different applications. Since there is no
meaningful orbital assembly capability planned at this
time, any large space structures will have to be selfdeployable.
Current concepts for large, conventional,
mechanical, self-deployable space structures tend to be
,--very expensive and mechanically complicated. Current
mtenna-user requirements are so stringent (with respect
to the need for very low-cost, high-deployment
reliability, low weight, and packaged-volume and usable
aperture precision) that new and innovative approaches
to accommodate large space structures are needed.
Fortunately, a newly developed class of space structures,
called inflatable-deployable structures, has great potential
for satisfying these stringent user requirements. A
concept under development at L’Garde, Inc., for a large,
Z inflatable-deployable antenna represents an excellent
example of this new type of structure.
The NASA Office of Space Access and
Technology initiated l the In-Space Technology
Experiments Program (IN-STEP) specifically to
accommodate the verification and/or validation of
unique, innovative, and high-payoff technologies in the
space environment. The potential of the L’Garde, Inc.,
*Manager, IN-STEP Inflatable Antenna Experiment,
Applied Mechanics Technologies Section
**Program Manager, Inflatable Antenna Experiment
Txperiment Principal Investigator
Copyright d 1995 by the International Astmnu_ttical Federation. All tights
reserved.
concept has been recognized and resulted in its selection
for an IN-STEP experiment. The objectives of the
experiment are to verify low cost and light weight by
building a 14-meter-diameter flight-quality reflector
antenna structure, demonstrate deployable reliability in
a realistic environment, and measure the reflector
surface precision in a realistic gravity and thermal
environment. The approach will utilize the Space
Transportation System (STS)-launched, recoverable
Spartan spacecraft as the experiment carrier.
The flight-system functional performance
requirements originate from the experiment’s technical
objectives. The design requirements for the flight
hardware system are based on a combination of system
functional-performance requirements; basic inflatablestructures
capability; the L’Garde, Inc., technical data
base resulting from the development and launch of a
large number of inflatable “decoy type” structures; and
the space environment effects on a large, thin-film
structure in low earth orbit. The large-space
structure/environmental interactions include the effects
of atmospheric drag on the attitude stability of the
structure, the effects of the orbital thermal environment
and atomic oxygen on the thin-film materials, and loworbit
radiation effects on electronic components. Both
requirements and environmental effects are specified for
each subsystem; e.g., (a) the basic support structure that
houses the inflatable structure and the other subsystems
and interfaces with the experiment carrier, the Spartan,
(b) the inflatable-deployable antenna structure, (c) the
inflation system, (d) the surface-measurement system,
and (e) the electronic system. For each of the
subsystems, this paper will identify and describe the key
1
,-
and unique design drivers, the impact of the
environmental interactions, the type of analysis used for
simulating subsystem performance, the type of
developmental testing used for design refinement and
validation, the specialized processing, manufacturing and
assembly techniques, and description of the final design.
The experiment is being managed by the Jet
Propulsion Laboratory. The flight hardware
development at L’Garde, Inc., Tustin, California, is
currently in Phase C/b, aid
manifested to fly on STS 77
primary/sharing payload.
the experiment is
in April 1996 as
Introduction
Space-deployable antennas are needed for a
variety of applications that include space-based, verylong-baseline
interferometry (VLBI), mobile
communications, active microwave sensing, earthobservation
radiometry, synthetic aperture radar,
spacecraft communications, and DOD space-based
radar.’ Recent constraints on the availability of
,- resources for these types of applications within NASA,
Ile science community, the commercial sector, and even
the DOD, have resulted in stringent user application
requirements. Therefore, the real key to accommodating
these missions is the development of new concepts for
low-cost and mechanically reliable antenna structures.
Other important features include low weight, high
mechanical-packaging efficiency, usable aperture
precision, and long-term dimensional stability.
Realistically, however,
_. innovative concept
meaningful demonstrations of
capabilities will have to be
accomplished to attract any kind of serious user interest.z
A relatively new and unique concept for an
inflatable-deployable spice antenna structure that has
tremendous potential for accommodating such stringent
user requirements is under development by L’Garde,
Inc., Tustin, California.lp3 In fact, serious user interest
has resulted in the selection of this concept for a NASA
In-Space Technology Experiments Program (IN-STEP)
space-based experiment. This class of experiments is
based on demonstrating and evaluating the performance
of promising concepts with low-cost flight hardware.
The experiment objectives are selected specifically to
.ralidate antenna-user criteria and to demonstrate the
.evelopment of large, flight-quality hardware for a lowcost,
high mechanical-packaging efficiency, low weight,
high deployment reliability, usable reflector-surface
precision, and thermal stability in a realistic
environment.
The experiment is currently in the final stage of
flight-hardware assembly and qualitication testing. It is
manifested to be flown on STS 77 in late April 1996.
This paper describes the design of the experiment flight
hardware and identifies the key issues for each of the
subsystems that comprise the experiment system. The
information contained in this paper and in References 1
and 2 is intended to provide a complete summary of the
experiment justification, technical approach and flight
hardware.
Exneriment-Svstem Performance Reouirements
The experiment-system functional requirements
are based on the experiment objectives and the inflatablestructures
concept capability, constrained by the NASA
experiment resources available and the capability of the
experiment carrier, the Spartan (Figure 1). The antenna
structural contiguration is based on the L’Garde, Inc.,
basic inflatable-antenna concept. The 14-meter-diameter
reflector size is based on an extrapolation of the g-meter
baseline structures data base and the current size limit
for manufacturing capability at L’Garde, Inc.
Moreover, this structure can be accommodated by
Spartan, and it is large enough to be used for real
applications, such as VLBI and commercial mobile
communications. The surface-precision goal of 1 mm
rms on orbit is based on the current analytical
performance projections, manufacturing, assembly, and
alignment capability at L’Garde, Inc. Validation and
characterization of the deployment sequence will be done
on orbit, which provides a realistic operational
environment. High mechanical-packaging efficiency will
be demonstrated by stowing the inflatable structure in a
small canister. The inflight single-orbit measurement of
surface precision and its thermal stability will provide a
measurement of the concept value for different potential
applications.
Qualification of the experiment’s hardware is
being accomplished by both analysis and test for
compliance with functional performance and STS safety
requirements.
Subsvstem Functional Requirements
The experiment-subsystem functional
requirements are driven by the system functional
requirements, with design parameters bounded by the
7
SPARTAN
l USER
REQUIRE~~ENTS
l ICD
l PIP
l OWITAL
SERVICES
l STS SAFETY
l SEPARATION
EXPERIMENT OBJECTIVES
l DEVELOP LARGE, l DEMONSTRATE
LOW COST DEPLOYMENT
ANTENNA RELIABILITY ON
STRUCTURE ORBIT
l VALIDATE l MEASURE
MECHANICAL SURFACE
PACKAGING PRECISION ON
EFFICIENCY ORBIT
SYSTEM FUNCTIONAL
PERFORMANCE REQUIREMENTS
l 14 METER DIAMETER
ANTENNA
l CHARACTERIZE ORBITAL
DEPLOYMENT SEQUENCE
l SURFACE PRECISION
I mm rms
l SURFACE MEASUREMENT
l SPARTAN
l BASED ON EXISTING
CONCEPT AND
COMPATIBLE MATERIALS
-
CONCEPT
CAPABILITY
l NINE METER
BASELINE
l DESIGN
l MATERIALS
PROCESSING
l INFLATION
TECHNIQUES
l MECHANICAL
PACKAGING
l SURFACE
MEASUREMEN
ELECTRONICS CANISTER INFLATABLE INFLATION SURFACE
STRUCTURE SYSTEM MEASUREMENT
SYSTEM
Fig. 1. Experiment Design Requirements
L’Garde, Inc., flight data base for inflatable structures
and the environmental-interaction effects on the
_- experiment hardware (Figure I). The subsystems needed
to accommodate the experiment include (a) the inflatable
structure, (b) the canister or bus structure, (c) the
inflation system, (d) the surface-measurement system,
and (e) the electronic system (Figure 2). The
combination of these subsystems represents the simplest
approach for satisfying the system functional
requirements. The design and performance of the actual
flight hardware will be based on how well the subsystem
functional requirements were satisfied. The key design
drivers, the types of analysis used, the functional
developmental testing, and pictures of the final designs
will be discussed in the appropriate subsystem sections
of this paper.
lflatable Structure
The subsystem functional requirements for the
inflatable structure are given by Table 1. The antenna
3
configuration is an off-axis parabolic reflector structure
consisting of (a) a 14-meter-diameter, multiple-gore
reflector structure and a transparent canopy (which is a
mirror shape of the reflector) to maintain gas pressure
on orbit, (b) a torus structure that supports the
reflector/canopy circumferentially, and (c) three 2%
meter-long struts that interface the torus structure with
the canister which is located at the center of curvature of
the reflector to accommodate operation of the surfacemeasurement
system.
Reflector
The major challenge is to design and fabricate a
14-meter-diameter, multiple-gore reflector structure with
an orbital surface precision on the order of 1 mm rms
and with enough reflectivity to accommodate orbital
operation of the Surface Accuracy Measurement System
(SAMS). The system reflector-gore geometry is
determined by the L’Garde, Inc., FLATE code, which
uses the desired orbital-membrane configuration, its
Fig. 2. Experiment Orbital Configuration
operating stress levels, and materials properties of the
membrane and bonded seams to define the number of
gores, their zero stress shape, and the operating gas
pressure.4 The membrane-materials properties (which
include <- thickness and non-linear modulus) are
experimentally characterized for use in the detail design.
The techniques for handling, laying out, marking,
cutting, and the butt-joint bonding of the one-quarter-mil
mylar membrane were developed on previous programs
at L’Garde, Inc., and’ demonstrated on a g-meterdiameter
reflector structure. hJylar was selected for this
experiment because of its availability, low cost, and its
extensive use in previous flight applications of inflatable
structures.
Fabrication of the 14-meter-diameter reflector
was based on using 62 individual one-quarter-mil
aluminized mylar gores (Figure 3). The gores were
assembled on full-scale, specialized tooling that was
designed to account for the difference in curvature
etween the zero strain-assembly condition of the
membrane and the orbitally loaded configuration. The
seams are butt-joined, utilizing a doubler of the same
material on one side only. The adhesive was a standard,
space-qualified, flexible material used on a number of
previous programs. The ground handling of the
membrane required the development of special folding
techniques so that the material could be stowed in a
small compact package for ease of handling and
deployment.
The surface precision of the as-manufactured
reflector structure is determined by mounting the
membrane on a fixture that simulates its interface with
the canopy structure and torus. A pressure differential
across the structure (equivalent to that on orbit) produces
a surface that represents the “upper bound” of reflector
precision that would be achieved if the assembly were
perfect. The flight-reflector structure has a measured
surface precision on the order of 1 to 2 mm rms.
Canopy
The primary design requirement for the canopy
is that it should be a “mirror” image of the reflector
structurk. That is, the design should be based on using
4
,
Table 1. Inflatable Structure Subsystem Requirements
.
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-I
Based on L’Garde, Inc., inflatable-deployable
antenna concept
14-meter off-axis parabolic reflector
Surface accuracy goal of 1 mm rms
Optically reflective surface on reflector
Clear canopy
Torus and struts provide basic support structure
Packagable in a container compatible with
carrier vehicle
f/D = 1/2 (Parent)
Deployment time compatible with single-orbit
experiment
Structural Stability
Maximum reflector deflection with respect
to the spacecraft - 0.50 cm
Dimensional Stability
Torus kO.70 cm on diameter
Strut h2.5 cm on length
the same materials and number of gores, but would not
require the same surface precision or vapor-deposited
_. aluminum as that used on the mylar. This approach will
result in a force balance between the two doubly curved
surfaces at their interface where they are assembled to
form a lenticular structure. Since this canopy structure
is made and assembled with the same tooling as that
used for the reflector, no special tests were done to
verify the canopy’s configuration. Assembly of the
canopy with the reflector is accomplished by bonding
both structures at their outer perimeter to a flexible ring
structure. This interface structure provides sufftcient
stiffness to transfer the loads from the inflated lenticular
structure to the torus, yet it is flexible enough to be
stowed with the membrane structure. Mechanical
packaging tests of this structure are accomplished after
assembly of the “lenticular” structure.
Torus
The design of the torus structure is driven by the
size of the reflector and the circumferential tension loads
from the lenticular structure. These loads produce
compression and bending in the torus structure. The
material selected was neoprene-coated kevlar because (a)
L’Garde, Inc., has extensive experience with the
handling and bonding of this material, (b) it is
commercially available and inexpensive, (c) it stows
efficiently, and (d) it has adequate strength and stiffness
for accommodating the experiment. The detail design of
the torus is based on a L’Garde, Inc., specialized code,
which uses as input the external loading on the torus, its
geometry, operating stress level, and the material
properties to determine the required diameter and
operating pressure.5 The material properties for the
analysis are experimentally characterized. A full-scale
engineering model torus is shown in Figure 4.
The specialized tests for the torus include
mechanical packaging efficiency and neutral buoyancy
flotation to determine the attachment plane for the
torus/lenticular reflector assembly under zero g loading
conditions. The mechanical packaging consists of
repeated stowing of the torus in the canister structure,
Fig. 3. Reflector Membrane - 14.0 meters Diameter
5
Fig. 4. Torus Structure - 15.2 meters Diameter
using different “folding” techniques. Success for this
test is determined by the best packaging efficiency,
coupled to a final folding configuration that lends itself
to “deployment by inflation”, as established by previous
flight-hardware experience. The neutral buoyancy state
for the torus is achieved by (a) completely filling the
structure with water, (b) locating it in a trough filled
with water, (c) pressurizing the torus to a differential
pressure of 0.02 MPa, which simulates orbital loading,
and (d) applying a small amount of distributed flotation
to offset the negative buoyant forces resulting from the
tluid displaced by the volume of the fabric used for the
structure. This technique worked so well that the torus
could be manually displaced in the trough with
essentially no measurable restoring forces observed. In
this unloaded state, the mounting plane for the lenticular
structure was located with a rotating laser beam and
marked for subsequent attachment of the interface
hardware.
Assembly of the torus/lenticular structure is
based on using 62 discrete mechanisms located along the
mounting plane that can be adjusted individually to
impart a uniformly distributed load to the torus. The
adjustments are interactively made to simulate the
circumferential geometry of the lenticular structure when
properly aligned on its mounting fixturing. After the
integration of the two structures, additional mechanical
packaging tests are done, starting with the folding
configurations successfully developed with the torus
structure.
Struts
The design drivers for the struts include (a) a
Lequired length of 28 meters, (b) a minimum structural
frequency of 0.25 Hz, (c) manufacturing tolerances of
0.5 cm for bending and torsion distortions, and (d) an
operating pressure the same as that for the torus. The
detail design is based on using the same neoprene-coated
kevlar material as used on the torus. The diameter and
resulting bending stiffness are based on a requirement
for a minimum natural frequency to accommodate the
orbital stability needed for the experiment. The
minimum diameter required was 35.6 cm. The
materials-processing and fabrication techniques used are
near identical to the ones used on the torus. A full-scale
engineering model is shown in Figure 5.
Specialized testing included manufacturing
evaluation for bending, twist, and mechanical packaging
efficiency. The quality of the manufactured struts was
established by “floating”
trough.
them in a full-length water
Calibration marks on the ends of the tubes
indicated the degree of relative rotation. Measurement
of lateral translation along the length of the strut, as it is
rotated in the water trough, is a direct measure of the
bending as a result of manufacturing. These tests
showed that the torsional permanent set in the structure
was about IS”, but did not affect the functional
performance, and the bending was on the order of
50 mm maximum deflection.
The end fittings for the struts are essentially flat
plates machined from aluminum and bonded to the
interior area of the end of the tubes. Such fittings are
simple, inexpensive, and easy to interface with the
canister panel-support fittings and the torus-to-strut
interface fittings. The geometry of these fittings,
because of their size and rigidity, has a significant
impact on the mechanical packaging techniques. Folding
patterns were developed that kept the thin membrane
material from the vicinity of the end fittings.
The results of the mechanical packaging tests of
the canopy, torus, and struts individually were used for
Fig. 5. Structure - 30.5 meters Long
6
.
arriving at the packaging techniques for the complete
- till-size inflatable structure. Validation of this approach
for stowing the inflatable structure was done by adding
ML1 blankets and electrical wiring to the inflatable
model and successfully packaging it in the canister.
Canister
The design drivers for the canister include (a)
providing the load-carrying structure for all elements of
the experiment, except the equipment panel that remains
with the Spartan, (b) interface structure with the Spartan,
(c) deployable panels to accommodate ejection of the
stowed inflatable antenna structure, (d) smooth surface
compartment to house the stowed inflatable structure, (e)
interface with the struts, and (0 high structural-design
margins to minimize the need for expensive structuralqualification
verification testing (Table 2).
The design approach for the canister (or bus
structure) employs aluminum honeycomb panels for the
basic load-carrying elements of this structure. The
justification for this selection is that (a) it is very stiff for
its weight, (b) it is available and relatively inexpensive,
cc) the smooth face sheets provide the appropriate
urfaces for the stowed inflatable structure, and (d)
L’Garde, Inc., has extensive experience using this type
of structure.
The configuration development of the canister
starts with (a) the volume required for the stowed
inflatable antenna structure, (b) the volume and mounting
surfaces required for the inflation system, electronics,
Table 2. Canister Subsystem Requirements
.
House and support all elements of experiment
.
Basic load-carrying structure
Structural design loads from Spartan
Deployable
Provide ejector for inflatable structure
Provide interface for struts
Interface with Spartan
High design margins
and functional elements of the SAMS, (c) the geometry
of the interface with the Spartan and the limitations
imposed by the dynamic envelope of the STS, (d)
mechanization to accommodate release of the stowed
inflatable structure in a controlled manner, (e) providing
the mounting points for the three strut structures, and (f,)
taking advantage of the weight available for the
experiment to develop large, structural design margins
and a structure with a first natural mode above 35 Hz to
preclude a requirement for a modal survey test.
These requirements resulted in a canister
structure that is 2.0 meters long, 1.1 meters wide, and
0.46 meters high. Four spring-loaded doors open to
allow deployment of the inflatable antenna structure.
Part of the SAMS light panel is attached to the inside
surface of the top cover, and the three side doors contain
“pods” for storage of the inflatable struts. All doors are
held in place by pin-puller latches. A large springloaded
plate is located in the floor of the canister for
purposes of ejecting the inflatable structure at the
beginning of the deployment sequence. This plate also
supports the rest of the light panels.
The structural design of the canister was based
on quasi-static acceleration loads defined at the Spartan
center of mass. A standard finite-element code for the
determination of panel stresses and loads was used with
hand analyses for fittings and mechanisms stress, and
margins. Because of (a) the complexity of the
honeycomb construction for modeling, (b) the geometry
of the canister, (c) the mechanisms tie points needed to
accommodate articulation of the panels, (d) and an
ejection plate for pushing the stowed inflatable structure
away from the canister, the final model was 23,000
d.o.f. This model, afier experimental verification of the
fundamental structural modes below 50 Hz, was used
with the Spartan analytical model for the coupled-loads
analysis for the combined loads of the Spartan/IAE on
the shuttle.
Special functional tests for the canister included
(a) panel deployment, (b) ejection-panel spring
calibration, (c) pyro/pin-puller release, and (d) structural
natural frequency identification. The till-scale
engineering model and hardware used for the tests are
shown in Figure 6. The panel tests consisted of repeated
articulations with adjustments of spring stiffness and
damping of the actuator to ensure timely and low shock
deployment. The ejection-panel tests were based on
repeated deployments of a mass simulation of the
inflatable structure to accommodate evaluation and
7
Fig. 6. Canister Structure @us)
adjustment of the spring cluster to achieve the proper
ejection velocity. Pyro pin-puller release tests were
done to demonstrate a “clean” release and functional
performance of the “initial motion” kick-off springs.
Forced vibration tests of the full-scale engineering test
unit, which simulate the full-up canister system, were
conducted at the Goddard Space Flight Center (GSFC)
to determine all the significant structural modes below
?O Hz which turned out to be 35 Hz.. The results were
.ien used for test/analysis correlation for validation of
the structural model that was used for the
Spartan/IAE/STS-coupled loads analysis. A highly nonlinear
dynamic response of the stowed intlatable
structure resulted in the need to analytically account for
the change in the response frequency of two significant
modes in the linear analytical model. This was
necessary to account for the actual frequency shift of
*several Hz in the structure when under high-level
dynamic loading, since simulation of the non-linear
characteristics is not practical.
Surface Accuracv Measurement Svstem
The design drivers for the surface measurement
subsystem include (a) remote measurement of the
retlector surface on orbit and in the presence of near
direct sunlight with a resolution of &-0.1 mm rms, (b)
coverage of at least 90% of the surface, (c) a
measurement cycle of no more than 40 seconds, and (d)
a development and flight hardware cost of under $lM
(Table 3). A number of systems were identified for
qossible application to IAE. However, only one system
as even close to being affordable for the IAE. That
system is based on a Digital Imaging Radiometer (DIR)
developed by McDonnell Douglas for measurement of
slope errors on ground-based solar concentrators.6*7*8
The concept is based on using a number of discrete light
sources, located near the center of curvature of a
surface, and then photographing the reflected rays.
Surface deviations from a perfect surface will result in
shading patterns as seen by the camera. The magnitude
and distribution of such shading patterns are used to
determine the slope error distribution of the antenna
aperture. The components needed to implement this
approach include (a) a number of discrete light sources
mounted on panels, (b) high resolution video cameras,
(c) video records, and (d) electronic circuits to sequence
and control the triggering of a large number of shortinterval
light bursts.
The initial step of the design was to analytically
characterize the system performance parametrically, with
Table 3. SAMS Subsystem Requirements
Remote measurement of reflector surface
Concept based on MDAC DIR
Measurement data recorded on VCR
Sample data transmitted to STS
Measurement accuracy &-0.1 mm rms
Surface Coverage 290 percent
Surface measurement cycle 540 seconds
Development and flight hardware under $lM
1
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the resulting information used to determine ,(a)I
the
,- number of light sources? (b) the light panel size and
shape, (c) the characteristics of the light sources, such as
wavelength and luminous intensity, (d) the camera
characteristics that include pixel resolution, flux
sensitivity, band pass filter and its center, and (e) the
electrical and software requirements for operating the
cameras and timing for the light panels.
The next phase of system development involved
identitication of flight-qualified components for the
system. The biggest challenge, of course, was the video
cameras. Fortunately, a Videospection CCD camera that
had been previously flown on the STS was identified and
found to meet the functional performance requirements
(Figure 7). The next challenge was to identify light
sources (such as high-output LEDs) that were available
in large numbers at low cost, since the final design
utilizes 512 clusters of 36 LEDs (Figure 7). A relatively
new product by Rohm Corp. was located, but was not
tlight-qualified. However, since the flight system will
operate with up to 20% LED failure, the usual flight
certification was not required. The mounting of the
individual LEDs on the supporting panels utilized
standard techniques for mounting electronic components
on printed wiring boards.
The special tests required for development of the
Uight hardware included (a) reflectivity of the aluminized
mylar to LED illumination, (b) camera aperture
evaluation and calibration, (c) light-source intensity
evaluation, (d) system characterization, using a scale-size
calibration mirror, and (e) full-scale calibration, using a
3-meter-inflatable section of the fi&size reflector
structure. The retlectivity of the illuminized mylar was
obtained by mounting the actual aluminized mylar
membrane on flat plates, illuminating the reflection
surface with the flight-type LEDs, and measuring the
return signal. This reflectivity data was used in the
design of a 0.30 meter-diameter glass mirror to be used
VIDEO CAMERA VCR
LIGHT PANEL SET
Fig. 7. Surface Accuracy Measurement System
9
_ for subsequent system calibration. The camera-aperture
,alibration was accomplished by tests in direct sunlight
and in the dark with variable LED intensity. The lightsource
intensity was determined by direct measurement
as a function of applied vo!tage. The complete system
was evaluated for the first time by using the 0.30 meterdiameter
mirror, which had the same reflectivity as that
of the membrane and a 28-meter radius of curvature.
eis test established the required flux density of the
return signal, the required camera-aperture opening, and
the adequacy of the LED’s output. The full-scale test,
using a 3-meter section of the 14-meter reflector offered
the first opportunity to evaluate the system in a realistic
manner. By rotating the camera so that its field or view
moves across the surface of the inflatable reflector, a
simulation of an on-orbit measurement has been
achieved. The results of this test demonstrated that (a)
the measurement resolution of the system was on the
order of 0.1 to 0.2 mm rms, (b) the final cameraaperture
settings were established, (c) the time required
for multiple measurements was verified, and (d) the final
configuration for the light-panel performance was
determined.
.- The demonstrated performance of the surface-
.lleasurement system effectively satisfied the subsystem
design requirements. In fact, a number of performance
results exceeded expectation. The total development and
hardware cost was under $lM; measurements were
successfully made in near direct sunlight; and the flightqualified
cameras were more than adequate for the
system.
.-Inflation Subsvstem
The key design drivers for the inflation
subsystem included (a) high-pressure nitrogen gas
storage for the inflatable.structure, (b) sensors, valves,
and regulators 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 (Table 4).
The functional design of the 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
?quirements. Component selection was based on
tireviously qualified hardware used for the Spartan
attitude-control, cold-gas system and on previous
L’Garde, Inc., flight systems. The supporting structure
Table 4. Intlation Subsystem Requirements
Active pressure control system
Provide pressure vessels, regulators, sensors,
and valves to supply 52930 cc NZ at 20.68
MPa
Control inflation pressure to the required level
- Deployment, design flow rate 53 %
- Deployed, support structure 0.02 MPa
&-3% static
- Deployed lenticular structure 2.07~10~~
MPa &3%
Maximize use of Spartan components
Maximize use of previous successful L’Garde,
Inc., designs
used for mounting the tanks, plumbing, and components
is an aluminum honeycomb panel similar to that used for
the canister. This panel also contains the electronics,
control boxes, and SAMS video cameras and is
permanently attached to the Spartan for return to earth
after completion of the experiment. The two large
structural composite gas tanks utilize the same mounting
configuration as that for Spartan to minimize
requalification costs. The component mounting and
tubing are similar to previously used designs by
L’Garde, Inc. (Figure 8). Design factors of 2.5 over
operating pressure were used. No attempt was made to
develop a light-weight, highly compact inflation system
for this experiment because of cost limitations.
Specialized testing included leak, proof-pressure,
and functional performance validation. Simulation of
Fig. 8. Inflation System
10
fl orbital inflation was done by using two large tanks which
were evacuated and then filled with gas using the sensing
and gas-flow control techniques proposed for the
experiment. Refinements of the design were based on
these test results.
The inflation subsystem development was
relatively straightforward, and its functional performance
easily satisfied the subsystem requirements.
Electronic Subsvstem
The design driver for the electronic subsystem is
the initiation, sequencing, and control of all IAE
functions that include (a) pyrotechnic release devices, (b)
pyrotechnic valves, (c) synchronization/control of the
video cameras, VCRs, and light panels, (d) multiplexing
of engineering data, (e) logic and control of the
inflatable pressures, and (f) interface with the Spartan
(Table 5). The electronics are designed around the Intel
87Cl96K-MOS processor.
The subsystem design was based on conventional
electrical-circuit analysis and hardware packaging
techniques successfully used on previous L’Garde, Inc.,
flight hardware. The electrical components were
selected from JPL parts lists that identify sources for
high-quality and reliable hardware. The components are
mounted on four-layer circuit boards using standard
approaches. The circuit boards are integrated by
insertion into the motherboard (Figure 9). The
completed motherboard/circuit board assembly is housed
in an aluminum enclosure which is mounted on the
. equipment assembly structure. Cabling used in the
electronic subsystem utilizes standard nickel-plated-type
connectors.
Functional tests were based on component
evaluation, circuit characterization, assembly
performance, and overall subsystem capability.
Table 5. Electronic Subsystem Requirements
l Experiment sequence and timing
l Deployment control and inflation control
l Instrumentation control
l SAMS control
l Health and status monitoring
0 MIL components to maximum extent
possible
PROTOTYPE ELECTRONICS
Fig. 9. Electronic Subsystem
Environmental Interactions
The experiment structure/environmental
interactions include the effects of (a) atmospheric drag
on the stability of the structure, (b) orbital thermal
environment and atomic oxygen on the thin-film
materials, (c) UV radiation on the thin-film materials,
(d) low-orbit radiation on electronic components and,
(e) impact of space debris on the structure.
The atmospheric drag on the 14-meter solidsurface
antenna reflector results in a significant relative
separation of the Spartan/IAE from the Orbiter. This
separation (without STS station keeping) is 8 km, 20-km,
and 50 km on the first, second, and third orbit,
respectively. Consequently, the experiment was
designed to be completed during just one orbit to
minimize STS propellant. Additionally, analysis has
shown that these same drag forces tend to stabilize the
deployed structure with its longitudinal axis parallel to
the ram direction. This means that minimal control
authority is required.
Consideration of the orbital thermal environment
is included in the design of the inflatable support
structure by the addition of ML1 blankets to the torus
and struts to maintain dimensional stability. However,
the lenticular structure, which consists of the transparent
canopy and the aluminized mylar reflector, was not
treated in any way for the control of temperature during
the one-orbit experiment. Consequently, the internal
pressure will decrease significantly when the structure
falls in the earth’s shadow; but the inflation system will
compensate for this. However, for a long-term
application of this concept, different and improved
flexible materials and thermal control coatings will be
11
employed to maintain dimensional stability of the
Y reflector.
Atomic oxygen is not considered a problem for
this experiment because of its one-orbit duration. The
estimated degradation of the reflector membrane for one
orbit is 7.62 x lo-’ mm. However, for real applications,
special materials and/or surface coating will have to be
utilized.
The effects of UV radiation on the thin-film
materials and the South Atlantic Anomaly on the IAE
electronics were addressed. There are a number of
candidate membrane materials with real potential for
being radiation resistant, as compared to those used for
the experiments. For long-term applications, such
materials would have to be used. Cost constraints for
the experiment precluded the use radiation-hardened
electronic components. Therefore, to significantly lower
the probability of a single-event upset in the 296 km and
39’ inclination orbit, STS operations specify that the
SpartamIAE will not be put into orbit near the South
Atlantic Anomaly.
The impact of the IAE colliding with sizeable
ipace debris is not considered a problem for the
experiment because of adequate make-up gas. Even in
a long-term application, this is still not a major problem,
since the torus and struts would be rigidized, and the
operating pressure in the lenticular structure would be
two orders of magnitude below that of the IAE,
requiring only a small amount of make-up gas.
Conclusions
,*
Significant accomplishments at this time include
(a) the fabrication of a large, flight-quality-deployable
space structure for under $lM, (b) demonstration of
mechanical packaging of a 14- by 28-meter space
structure in a container the size of a large oftice desk,
(c) manufacture of a 14-meter-diameter reflector
membrane that has a surface precision on the order of 1
to 2 mm rms, (d) development of a space-qualified,
surface-accuracy measurement system that operates in
the presence of near direct sunlight for well under $lM,
and (e) the experimental determination of the
torus/canopy interface on a large, inflatable torus by
simulating O-g in a full-scale, neutral-buoyancy trough.
The remaining experiment objectives to be
accomplished on orbit include (a) validation and
characterization of the deployment sequence, (b)
determination of the reflector-surface precision and its
thermal stability in a realistic operational environment.
Acknowledgments
The research 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. Development of
the flight experiment is being done by L’Garde, Inc.,
under contract to JPL.
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