A deployable mast concept which meets the weight, size and stability requirements for a feed support structure for offset antennas up to 100 meters in diameter is discussed. A triangulated truss configuration, the use of tapered tubes which exhibit a high strength-to-weight ratio, and low CTE graphite-epoxy material are seen to provide an efficient, lightweight and stable truss suitable for an antenna feed support. A low stowage ratio of 30:1 is achieved through a unique preloaded hinge located at the center of each longeron and an autonomous deployment cage with a drive mechanism. Initial analysis and proof of concept hardware validated the basic mechanism and design assumptions and provided a basis for further investigation. The concept can readily accept variations in member size and thus lends itself to optimization for other potential uses where a stiff, lightweight deployable truss is needed

Craighead, N. D.

Hult, T. D.

Preliasco, R. J.

English

NASA Technical Reports Server

DESIGN,DEVELOPMENT AND MECHANIZATION 
OF A PRECISION DEPLOYABLE TRUSS 
WITH OPTIMIZED STRUCTURAL EFFICIENCY 
FOR SPACEBORNE APPLICATIONS 
BY 
N. D. Craighead 
T. D. Hult 
R. J. Preliasco 
Lockheed Missiles and Space Company, Inc. 
ABSTRACT 
The demand for large space platforms and antenna systems has identified the 
need for extremely long, stiff deployable booms. This need has prompted 
Lockheed Missiles and Space Company, under contract to the Jet Propulsion 
Laboratory, to develop the technology which could provide a space mast 
structure with a length in excess of the height of the Washington Monument. 
This structure will be capable of repeatable precision deployments in the 
space environment without external aids. The lightweight truss structure 
which also functions as a precision mechanism can be stowed within the 
Space Shuttle cargo bay. This paper will discuss the design, predicted 
performance and hardware development of this structure. 
The structure, which will be described, is a unique application of a trian- 
gulated truss that successfully blends the mechanisms of a deployable struc- 
ture .with those of an efficient extended structure. Individual internal 
members are double tapered graphite-epoxy tubes for maximum strength/weight 
ratio and stowed efficiency. The longitudinal members are hinged at the 
mid point to achieve a simple mechanism joint which provides for a compact 
folding scheme. Precision alignment and alignment repeatability is pro- 
vided by a unique prestressed joint design and pretensioned diagonal members 
which eliminate deadband. These features integrate into a truss system 
that exhibits a stiffness-per-unit-length ratio and a stowage efficiency 
unsurpassed by any currently available design approaches. The design syn- 
thesis and detailed design work will be presented for the structure in toto 
as well as the joint details and deployment devices. Also to be summarized 
are the fabrication and tests on selected critical members and joints, the 
performance projections based on a mathematical model in the form of para- 
metric results, and a comparison of the theoretical and experimental com- 
ponent test results. 
315 
https://ntrs.nasa.gov/search.jsp?R=19820015491 2020-03-21T16:54:27+00:00Z
INTRODUCTLON 
The industrialization and colonization of space will require the establish- 
ment of truly large space structures in orbit. Though desirable, the on- 
orbit assembly and/or fabrication of &ch.structures encounters severe eco- 
nomic and technological limitations for the near future. 'Thus, the need 
exists to begin the examination of this new technology through the develop- 
ment of reduced scale systems. Self-contained, deployable structures serv- 
ing as an investigative foundation for later, more ambitious projects will 
fill this need. .As an economic incentive, these systems should be useful 
in their own right, while serving as stepping stones toward higher goals. 
The "focus mission" concept (Ref. 1) is an effort by NASA to sift the crit- 
ical technologies required for various mission objectives and combine them 
into a hypothetical mission which will support their development. One of 
these critical technologies is the development of a class of large aperture 
antenna systems (see Figure 1). That model requires reflector apertures 
ranging from 30 to 300 meters in diameter, operating at radio frequencies 
from TJHF to Ku-band. Lockheed Missiles and Space Company, Inc., under the 
sponsorship of the NASA/LSST Program, has initiated the development of a 
55-meter offset wrap rib antenna demonstration project, scalable to 100 
meters, that addresses these requirements (Ref. 2). 
Concurrent with this effort, LMSC is proceeding with the development of a 
deployable feed support boom. As the 55/100-meter antenna system project 
is a hypothetical mission, no specific mechanical requirements for the 
boom have been identified. General requirements, however, may be derived 
from the expressed desires to incorporate specific features into the 
mission, stowage and weight limitations of the STS and the functional 
requirements for an antenna system of this size and frequency range. The 
boom is thus designed to: function with an S-band offset fed reflector up 
to 100 meters in diameter (possessing booms with a total length of 208 
meters, and a positional accuracy of 6 mm between ends), be compatible with 
the shuttle envelope, and provide a re-stow capability. In addition, the 
boom will provide a definable stiffness at all phases of deployment and 
retraction, provide maximum stiffness within the allotted weight and envel- 
ope, and weigh no more than 227 kg (500 lbs). Its characteristics have the 
capability to be tailored to suit other potential uses such as ground-based 
mobile towers, solar panel support and solar sail booms, and as an ele- 
mental building block for large space platforms. 
A review of the design synthesis and accompanying analysis is given in this 
paper. 
316 
DESIGN DESCRI.PTION 
Current deployable truss designs suitable for on-orbit use sacrifice many 
of their structural characteristics to achieve their deployable function. 
Recognizing that this mast must function primarily as a stable structure, 
the design synthesis began by selecting an efficient truss; then a means 
of mechanization was conceived that did not compromise the primary objec- 
tive. The result is a feed support consisting of two dedicated subsystems: 
a folding mast structure and deployment cages containing the mechanisms 
which extend it. 
Structural elements and mast configuration were driven by the need for 
an efficient truss. The selected arrangement is shown in Figure 2. As 
designed, the mast is a simple, three-sided, completely triangulated and 
preloaded truss. This system is an efficient structure wherein all loads 
are carried in tension or compression along load paths intersecting at node 
points. Doubly tapered tubes are used for the longerons and battens to 
achieve a maximum stiffness-to-weight ratio. Simple small-diameter tension 
rods serve as diagonal members. Pretensioning of these diagonals elimin- 
ates clearance in the longeron pivot bearings which ensures structural con- 
tinuity and provides the majority of torsional stiffness. 
The final mast configuration compatible with 55-meter reflector structural 
requirements is a 122-meter-long, 40-bay mast composed of 3.05-meter-long 
graphite-epoxy tubes with a 7.62-2.54-cm-diameter taper and a .528-mm wall 
thickness. Possessing a mass of 172 kg, it has a first mode natural fre- 
quency of .105 Hz and stows to a height of 3.32 meters. 
TRUSS MECHANIZATION 
The design requirements for mechanization were as follows: 
F deployment must be automatic and reversible, 
l hinges, joints, etc. must have no detrimental effects on 
extended rigidity, 
a mast, reflector and support equipment must stow within STS 
cargo bay envelope 
These requirements were met by incorporating a center joint and pivoting 
end fittings for each longeron and flexible diagonal members. These fea- 
tures allow the mast to stow as shown in Figure 3. The longerons fold out- 
side the battens while the diagonals stow inside, providing a stowage 
ratio in excess of 3O:l ($5-meter design). 
317 
LONGERON CENIER JOINT 
The longeron center joint is designed to meet the requirements of low 
weight, zero deadband, and positive locking at full extension. In addition, 
it provides a force to aid the deployment of the longeron. 
The joint consists of a number of short links connected to fittings 
attached to the longeron halves. Figure 4 shows the joint in various 
stages of deployment. The center link pivots about the longeron center- 
line, facilitating the side-by-side folding of the longeron while avoiding 
the length extension impacts of a side hinge. As the two halves are 
allowed to rotate for mast extension, the four-bar linkage formed by the 
center link, control arm, and end fittings controls the motion of the 
joint; when fully open the two toggle links extend and lock the joint into 
position with sufficient preload to eliminate deadband for the orbital 
loadings. 
The locking force is provided by a torsion spring located at the pivot 
point of one of the toggle links. Due to its placement and the kinematics 
of the joint, the direction of the resultant force changes during opera- 
tion. Figure 5 graphs typical values for the moment about the longeron 
center during joint operation and compares them to predicted values. The 
positive values denote a moment acting to extend the longeron, while neg- 
ative values denote a force acting to stow the longeron. The motion is 
limited by a stop slightly before the toggle links are allowed to reach 
full extension. This forces them into compression, which in turn causes a 
tension load in the two center links. This preload eliminates backlash 
caused by bearing clearances. 
A model of the joint is shown in Figure 6. Aluminum was used for ease of 
manufacture, availability and cost. For flight use titanium or chopped 
graphite fiber-reinforced plastic materials will be used. 
END JOINTS 
The longeron end joints are simple clevis connections as shown in Figure 7. 
The bearing clearances in this joint are eliminated by the tension in the 
diagonals. 
DIAGONALS 
The diagonal tension members are designed as several layers of unidirec- 
tional graphite/epoxy fibers having a rectangular cross section, typically 
1.5 x 5 mm. They are formed with a curve along their length which, when 
the mast is stowed, causes the diagonals to lie inside the truss, avoiding 
potential interference with the longerons or deployment mechanism. 
318 
EXTENSION/RETRACTION MECHANISM 
The mast extension sequence is illustrated in Figure 8. The stowed mast 
is held in the deployment cages which are functionally divided into two 
compartments, one for handling the stowed mast and one for extending the 
bays. The stowed mast is slowly raised toward the forming compartment; 
the mechanism in the forming compartment lifts a single batten assembly 
and extends the longerons and diagonals of one bay until the bay is fully 
formed. The process is repeated at the rate of approximately 1 bay every 
2 minutes until the complete mast is fully formed. These steps are 
reversed to retract the mast. A simplified schematic of the device is pic- 
tured in Figure 9 which illustrates the high/low speed drive arrangement 
and the function of the gear boxes, driver and belts. Not shown are torque 
tubes interconnecting the drive motors to enhance redundancy and assure 
speed synchronization. A unique longeron forms the right angle bend in the 
mast. A significant design feature of the deployment mechanization is that, 
during deployment, loads are transmitted through the deployed mast sections 
into the upper deployment cage and around the deploying sections. This 
feature assures a predictable structural stiffness throughout the entire 
deployment operation. 
Additional considerations were given to the transfer of ascent loads and 
subsystem interfacing. The cages function as chassis to which the various 
spacecraft subsystems are mounted and serve to transfer spacecraft loads 
to the boost vehicle. Loads to the orbiter are transferred through an 
adaptor ring. Figure 10 shows a complete spacecraft (.55-meter configura- 
tion) stowed in the orbiter bay. 
ADDLTIONAL MODELING AND DEVELOPMENT 
A full-scale demonstration longeron was constructed of T-300 autoclave- 
cured graphite epoxy on an aluminum mandrel. Fiber layup of the tube shown 
in Figure 11 is O/45/45/0 degrees. This layup represents a compromise 
between the desires for a low thermal expansion coefficient (CTE = 4.0 x 
10m6/"C along the longitudinal axis), adequate torsional stiffness (G = 
11 GPA), and a reduction in the probability of microscopic crack propaga- 
tion due to differences in the ply orientation angle. End fittings and 
the center joint were bonded in place on a fixture that ensured a precision 
alignment between ends and a dimensional exactness of the overall part. 
As a complement to the ongoing design layout work, a mathematical model of 
the truss was produced to predict the structural and mass properties of the 
truss. This model can also be used for both parametric and point design 
case studies. Functioning as a design tool, such analysis allows member 
sizing and optimization consistent with design goals and physical constraints. 
319 
Figures 12 and 13 are examples of the parametric data generated. They 
represent the cantilever bending stiffness and mast weight, respectively, 
as a function of the mast geometry and material thickness. 
CONCLUSIONS 
Preliminary design and development work at Lockheed Missiles and Space 
Company, Inc. has resulted in a deployable mast concept which meets the 
weight, size and stability requirements for a feed support structure for 
offset antennas up to 100 meters in diameter. A triangulated truss config- 
uration, the use of tapered tubes which exhibit a high strength-to-weight 
ratio, and low CTE graphite-epoxy material provide an efficient, light- 
weight and stable truss suitable for an antenna feed support. A low stow- 
age ratio of 3O:l is achieved through a unique preloaded hinge located at 
the center of each longeron and an autonomous deployment cage with a drive 
mechanism. Initial analysis and proof of concept hardware have validated 
the basic mechanism and design assumptions and provided a basis for further 
investigation. The concept can readily accept variations in member size 
and thus lends itself to optimization for other potential uses where a 
stiff, lightweight deployable truss is needed. 
REFERENCES 
(1) Freeland, R. E. and Campbell, T. G.; Deployable Antenna Technology 
Development for the Large Space Systems Technology Program; AIAA/NASA 
Conference on Advanced Technology for Future Space Systems, Hampton, 
VA, May 8 - 10, 1979, NASA CP-2118. 
(2) Wade, W. D. and McKean, V. C.; The Technology Development Methodology 
for a Class of Large Diameter Spaceborne Deployable Antennas; 15th 
Aerospace Mechanisms Symposium, Huntsville, AL, May 14 - 15, 1981, 
NASA CP-2181. 
320 
FIGURE 
DEPLOYMENT 
CAGE 
LONCERON JOINT 
L TILT PLATFORM 
INTERFACE 
FIGURE 2. PARTIALLY DEPLOYED MAST 
(DEPLOYMENT CAGES ASSURE A DEFINABLE LOAD PATH. DURING MAST DEPLOYMENT) 
321 
I 
FIGURE 3. STRUCTURAL ELEMENT STACKING CONFIGURATION 
322 
CONTROL LINK 
CENTER LINK 
HALF DEPLOYED 
, /-PAD (COMPRESSION) 
CENTER LINK (TENSION) 
TOGGLE (COMPRESSION) 
L LINK (TENSION) 
SECTION A-A (PARTING LINE) DEPLOYED AND LOCKED 
FIGURE 4. LONGERON CENTER JOINT SCHEMATIC 
1 - 
0.9. 
0.8- 
0.7- 
0.6- 
-,‘IM’ 
DEPLOYMENT ANCLE.DECREE 
- PREDICTED 
- - - TESTED 
FIGURE 5. TORQUE.ABOUT HINGE CENTER VS DEPLOYMENT ANGLE 
323 
FIGURE 6. LONGERON JOINT-A PRELOADED, OVER-THE-CENTER LATCH THAT IS 
SELF-LOCKING AND EXHIBITS NO DEADWD 
6-7 
LONCERON PIVOT BEARING (2) 
BATTEN SEGMENT 
L EXTENSION RACK (2) 
FIGURE 7. TRUSS CORNER-JOINT 
(ALL LOADS INTERSECT AT CORNER NODE POINTS.) 
324 
---r--t STOWED 
BATTENS CONTAINER 
STOWEQ PARTIALLY FORMED 
MAST SECOND BAY 
FIGTJRE 8. EXTENSION/RETRACTION PRINCIPLE 
EXTENSION DRIVE3 _ 
GEAR BOX 
AND II‘ 
TORQUE 7000 RPM, 
TUBE 
TAKE-OFF - 
1000: 1 
(GEAR HD+ 
BULL AND 
POSITION - 
DRIVE 
FIGURE 9. E~ENSION/RETIUCTION DRIVE 
SECOND BAY 
FORMED 
POSITION 
18 cmlMlN 
STOWED BATTEN 
MECHANISM 
325 
ARRAY DEPLOYMENT 
DEVICE 
REFLECTOR 
SOLAR ARRAY 
ANTENNA 
PALLETIZED 
ATTACHMENTS 
FIGURE 10. SPACECRAFT IN STOWEil CONFIGURATION 
FIGURE 11. TAPEREDGRAPHITE TUBES OFFERI~~GHSTRENGTKANDLOW~TGEET 
326 
8.3 1950 
‘a 
x 1800 
P 
z 
6 - 1650 
5; 
i=,z 1500 
f” 
Oli nv 1200 
Rz 
ia 
<iTI 900 
zm 
P 
600 
300 
0 
\I 3 IN - 1 IN B LONCERONS 
!i 
12 FT BATTENS 
THICKNESS KEY 
I’ I ‘I II I I\ 
\I 
THORNEL - 50 GrlE 
:’ 
\ \ = 0.0208 IN \ ------- = 0.0832 IN   .  I  
\ \ --- = 0.1664 IN   .  I  
\ 
1 \ 
\ 
\ \ 
\ \ 
TO OBTAIN THE OVERALL 
EQUIVALENT BENDING 
STIFFNESS, DIVIDE THE 
NORMALIZED STIFFNESS 
BY THE DESIRED NUMBER 
OF BAYS 
0.66 1.0 1.33 1.66 2.0 
BAY ASPECT RATIO (L;N;;;;N) 
FIGURE 12. EFFECT OF MATERIAL THICKNESS AND ASPECT RATIO ON NORMALIZED 
BENDING STIFFNESS (NORMALIZED ~0 BAY LENGTH) 
3 IN -1 IN ‘9 
12 FT BATTENS 
1 I 1 THICKNESS KEY 
: 0.208 
m-w--- z 0.832 
-w- = 0.1664 
LONCERON ASPECT RATIO ( B TTEN ) 
FIGURE 13. EFFECT OF MATERIAL THICKNESS AND ASPECT RATIO ON MAST 
STRUCTURAL WEIGHT 
327 
Timothy D. Hult 
Lockheed Missiles and Space Company 
P.O. Box 504 
Sunnyvale, California 94086 
Mr. Hult is presently an associate design engineer in the BP/Antenna Systems 
Laboratory involved in the engineering of large space deployable reflector 
systems. He received his B.S. degree in Mechanical Engineering from Clarkson 
College in 1980. 
RLchard J. Preliasco 
Lockheed Missiles and Space Company 
P.O. Box 504 
Sunnyvale, California 94086 
Mr. Preliasco is currently assigned to the BP/Antenna Systems Laboratory work- 
ing as an associate design engineer on large space deployable reflector sys- 
tems. He received his B.S. degree in Mechanical Engineering from Worcester 
Polytechnic Institute in 1980. 
Co-author of this paper is Mr. N. D. Craighead who is also with Lockheed in 
Sunnyvale. 
328 


Design, development and mechanization of a precision deployable truss with optimized structural efficiency for spaceborne applications

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