On the second flight of the INTELSAT V spacecraft the time required for successful deployment of the north solar array was longer than originally predicted. The south solar array deployed as predicted. As a result of the difference in deployment times a series of experiments was conducted to locate the cause of the difference. Deployment rate sensitivity to hinge friction and temperature levels was investigated. A digital computer simulation of the deployment was created to evaluate the effects of parameter changes on deployment. Hinge design was optimized for nominal solar array deployment time for future INTELSAT V satellites. The nominal deployment times of both solar arrays on the third flight of INTELSAT V confirms the validity of the simulation and design optimization

Gore, J. V.

Plescia, C. T.

Vorlicek, P. L.

English

NASA Technical Reports Server

DESIGN AND ANALYSIS CONSIDERATIONS 
FOR DEPLOYMENT MECRANIS!lS IN A SPACE ENVIRONMENT 
By: P.L. Vorlicek, J.V. GOre, C.T. Plescia 
Ford Aerospace & Communications Corporation 
ABSTRACT 
Cn the second flight of Ford Aerospace and Communication Corporation's 
INTELSAT V spacecraft the time required for successful deployment of the 
north solar array was longer than originally predicted. The south solar 
array deployed as predicted. As a result of the difference in deployment 
times a series of experiments was conducted to locate the cause of the 
difference. Specifically, deployment rate sensitivity to hinge friction and 
temperature levels was investigated. In conjunction with these experiments a 
digital computer simulation of the deployment was created to evaluate the 
effects of parameter changes on deployment. As a result of the experiments 
and simulation, hinge design was optimized for nominal solar array deploy- 
ment time for future INTELSAT V satellites. 'Ihe nominal deployment times 
of both solar arrays on the third flight of INTELSAT V confirms the validity 
of the simulation and design optimization. 
INTRODUCTION 
As satellites grow in size,the need for stowing the satellite within the 
dimensions of the launch vehicle fairing becomes a serious design con- 
straint. For this reason spacecraft are being built which are stowed in one 
configuration and then deployed into another configuration once in orbit. 
An example of such a satellite is the successful Ford Aerospace and Corrununi- 
cations Corporation (FACC) INTELSAT V spacecraft. INTELSAT V is powered by 
tm solar arrays that deploy once the satellite reaches geosynchronous orbit. 
CXI the second flight of INTELSAT V the time required for successful deploy- 
ment of the north solar array was longer than predicted. As a result of 
this, a series of ground-based experiments was conducted in order to locate 
the cause. Results of the experiments showed much higher friction levels on 
the flight hinge assemblies than had been originally predicted. In addi- 
tion, friction levels increased significantly at the low temperatures 
expected in orbit. Additional experiments and computer simulations gave 
additional insight into the solar array deployment mechanisms. 
As a result of the experiments and simulations, hinge design was optimized 
for nominal solar array deployment time for subsequent INTELSAT V satel- 
lites. lhe successful deploment of both solar arrays on the third flight 
of INTELSAT V confirms the validity of the simulation and design optimiza- 
tion. This paper will present a case study of the analysis and design 
changes that resulted from the deployment described above. Problems encoun- 
tered in the analysis of the solar array deployment will be discussed. This 
is intended to give some insight and guidelines for designers and analysts 
for use in design of similar mechanisms. 
211 
https://ntrs.nasa.gov/search.jsp?R=19820015484 2020-03-21T16:27:41+00:00Z
THE SYSTEM 
The Spacecraft 
The FACC INTELSAT V spacecraft is a communications satellite capable of trans- 
mitting 12,000 voice channels and 2 television channels. FACC has been con- 
tracted to build 15 INTELSAT V's. INTELSAT V M-2 was launched December 6, 
1980 and is currently operational over the Atlantic Ocean, supplying voice 
channels between North America and Europe. EM-l, launched May 21, 1981, is 
also operational over the Atlantic Ocean. The third INTELSAT V, FM-3, was 
successfully launched December 15, 1981 and is currently undergoing pre- 
operational testing. Figure 1 is an artist's rendition of the spacecraft. 
Figure 1. The INTEISATVSpacecraft 
The Solar Array 
The solar arrays on INTELSAT V consist of a yoke and three solar panels that 
deploy in an accordion-type manner (see Figure 2). Ihe array has two deploy- 
ment mechanisms: (1) torsion springs and (2) closed cable loops. The torsion 
springs provide the energy to deploy the array. The closed cable loops re- 
strain the deployment of transferring torques between the hinges, synchro- 
nizing the hinge deployment angles, and controlling the deployment rate no a 
point within the structural capability. Springs placed on the closed-cable- 
loop cables compensate for changes in cable length due to temperature 
variation. The springs also allow the hinge lines to be at somewhat different 
angles during deployment and add considerable complexity to the mathematical 
model of the solar array deployment. See Figure 2.for a description of the 
deployment mechanism and solar array parameters. Figure 2 also contains draw- 
ings of the hinge assemblies. 
212 
roRs1cN - 
jPRINC 
UmA 
'TORSION 
DRIVE ASSEl4BLY 
CLOSED CABLE 
LOOPS (CCL'S) 
MCKING BAR 
DEPLOYMENT 
CCL PULLEY 
Figure 2. The Deployment Mechanisms and Hinges 
THE SOLAR ARRAY DEPLOYMENT 
On the second flight the south solar array deployment time was 22 seconds. 
The north solar array almost completely deployed in 22 seconds, but space- 
craft roll rate data indicate the array was still moving until 32 seconds 
after release. Original predictions were for both solar arrays to deploy in 
approximately 13.5 seconds. 
Figure 3 shows the roll rate of the spacecraft during the deployment of 
both the north and south solar arrays on the second flight. As the arrays 
are deploying, the roll inertia of the spacecraft increases, which results 
in a decrease in the spacecraft roll rate. From 0 to 22 seconds the space- 
craft roll rate steadily decreases from .34 deg./set. to .12 deg./set. 
Accelerometer data indicated the south solar array locked into the 
deployed position at around 22 seconds. 
Further analysis of the data indicates there is a slow decrease in the 
roll rate from 22 to 32 seconds, which implies the north solar array was 
continuing to slowly deploy. The oscillation of the roll rate occurring after 
22 seconds is caused by the first bending mode vibration of the south array 
after it has locked up. This damped oscillation appears to be vibrating 
around a slowly decreasing roll rate, but due to the quantized nature of the 
roll rate data, no change in the roll rate is directly measured after the 
oscillation has damped out. 
213 
+ + + + + , 
2 -0.2: + + + + + + l + . + ? 3 
= -0.3. 
0 10 
TIME (S&NDS) 
30 40 50 
Figure 3. Spacecraft Roll Rate Data From FM-1 
EXPERIMENTS 
Two sets of ground-based experiments were conducted using an engineering 
model and flight hinge assemblies in order to locate the cause of the deploy- 
ment time difference. A series of deployment tests was made with an engineer- 
ing model of the solar array at FACC. These tests were intended to duplicate 
the flight experience, or to give insight into the possible causes. The sec- 
ond series of experiments was the measurement of resisting (friction) torques 
on flight hinge assemblies. 'These tests were intended to provide better 
values of the resisting torques in thermal vacuum environments, representative 
of expected orbital conditions. 
Tests on Engineering Model 
The deployment tests on the engineering model were divided into tm groups: 
1. Attempts to duplicate the flight data 
2. Deployment tests with simulated hot and cold closed cable 
loops and hinges, duplicating on-orbit conditions. 
A photograph of the test setup is seen in Figure 4. The array deploys hor- 
izontally and is supported at the panel centers by a sliding bar support 
device. The support rig was slightly inclined (2 mn/m)to counteract for the 
friction of the support device and air drag. At this inclination the array 
deployed in 22 seconds, the same as the south solar array on the second 
flight. 
'Ihe attempts to duplicate the on-orbit results proved unsuccessful. Several 
test runs were made, varying the inclination of the support rig. No cases 
were recorded with the type of flight data experienced on-orbit. Additional 
tests were conducted in which the closed cable loop (CCL) was slipped off of 
the hinge pulleys. Again, no insight into the deployment time difference was 
gained. 
214 
Figure 4. The Experimental Model Test Setup 
The second series of tests on the engineering model was conducted to simu- 
late the on-orbit temperature conditions. The orbital configuration of the 
solar arrays during deployment is edge-on to the sun. This position results 
in a temperature difference between the sunward and shaded hinges. To evalu- 
ate the effects of hot and cold conditions a test was constructed where the 
upper hinges were heated to 85'C with lamps, and the lower hinges were 
cooled with gaseous nitrogen to -lOO°C. The deployment time in this confi- 
guration increased to 24.7 seconds. These tests indicated an increasing 
resistance torque level with decreasing hinge temperature. 
Additional tests ware conducted to examine the effect of the closed cable 
loop (CCL) temperatures on deployment times. Temperature changes on the 
CCL's affect the length of the cables, which will change the point at which 
the cable will go slack. When the effects of temperature on the CCL's were 
tested, significant variations in deployment time resulted. Under nominal 
on-orbit conditions the south array deployment time was three seconds faster 
than the north array during the tests on the experimental model. 0-1 the 
first flight the south solar array deployed 6.5 seconds faster than the 
north array. Evidently the temperature of the CCL's contributed to this 
difference. 
Beyond increasing the deployment time, the effect of the CICL temperature on 
the deployment did not reveal any insight into the deployment experienced 
on-orbit. Table 1 summarizes the four test runs described above. 
215 
I I I I 
60 90 120 150 
DEPLOYMENT ANGLE (Degrees) 
Figure 5. Hinge-Resisting Torque for Original Inter-Panel Hinges 
Table 1. Test Runs on Engineering Model 
Test Deployment Times (sec.) 
Number 1st Hinge Lockup Last Hinge Lockup Comments 
1 21.7 22.0 Rig inclination 2mn/m. 
Room temperature, baseline 
deployment 
2 24.0 24.7 Rig inclination 2mm/m. 
Nominal north panel on-orbit 
conditions. Upper hinges 
85OC. Lower hinges -lOO°C. 
3 21.7 22.9 Rig inclination 2mm/m. 
Nominal north panel taking 
into account CCL temp. effect. 
4 19.0 19.7 Rig inclination 2mm/m. 
Nominal south panel taking 
into account CCL temp. effect. 
Tests on Hinge Friction Levels 
The most important findings of the investigations were the measurements of 
the resisting torques on flight hinge assemblies. These measurements 
revealed that the resisting torques at low temperatures in vacuum were much 
larger than values obtained from measurements obtained on development model 
assemblies, and much larger than the values obtained in ambient conditions. 
216 
ORIGINAL HINGES 
P 
NEW HINGES 
DEPmYMENT ANGLE (Degrees) 
Figure 6. Hinge-Resisting Torques for Nodified Inter-Panel Hinges 
Figure 5 shows the resisting torque level variation with deployment angle at 
various temperature levels. The resisting torque does not include the tor- 
sion spring torques. The resisting torques at low temperatures are higher 
than initially estimated. The values are such that beyond a 100" deployment 
angle of the inter-panel hinges , inertia forces of the moving array are 
needed to assist the torsion spring torque in the deployment of the array. 
As a result of these findings all the hinges were given special lubrica- 
tion. In addition, the bearing tolerences were increased to allow for 
greater temperature variations. Results of the hinge friction measurements 
on the modified hinges are shown in Figure 6. The resisting torque level 
has been greatly reduced at low temperatures. Further tests showed that the 
friction level on the modified hinges is not as sensitive to temperature 
variation as the original hinges. 
THEORETICAL ANALYSIS 
In conjunction with the laboratory experiments, a theoretical model and 
digital computer simulation of the solar array deployment was developed. The 
purpose of the model was to: (1) recreate the on-orbit deployment results, 
(2) give additional insight into the deployment mechanism dynamics, and (3) 
provide a tool whereby the data from the friction tests could be evaluated 
in respect to deployment dynamics. In the simulation the panels and yoke are 
modeled as rigid elements interconnected with flexible hinges and extendible 
closed cable loops with accurately modeled temperature compensating springs. 
The torques acting on the hinges in the simulation include: 
1. The torsion springs at the hinges 
2. Torques from the closed cable loops 
3. Stick-slip coulomb friction torques (stiction). 
Stiction is a resisting force that always oFposes velocity and accounts for 
the fact that a finite force is needed to start a body moving. Accurate 
simulation of stiction is important for both the designer and analyst to 
consider, as this was eventually found to account for the differences in the 
north and south panel deployment times. 
217 
DEPLOYMENT SPRING 
TORQUE, ORIGINAL HINGES 
RESISTING TORQUE 
D 45 90 135 180 
DEPLOYNENT ANGLE (DEGREES) 
Figure 7. Spring and Resisting Torques Used in the Simulation on 
the Inter-Panel Hinges 
Examination of the results of the hinge friction tests shows the resisting 
torque to be angle dependent. Since the stiction torque is angle dependent, 
the torque was represented in the simulation as shown in Figure 7. Figure 7 
shows the torsion spring torques and the nominal resisting torques used for 
the inter-panel hinges in the nominal-case computer simulation. 
Several simulation runs were made, varying the driving torques and 
stiction level at each of the hinges. Results of the simulation are the 
deployment angles plotted against time. From these curves the deploment 
time of the panel can be determined. See Figure 1 for a definition of the 
deployment angles. Figure 8 is the result of a simulation of the deploy- 
ment of the north solar array with the nominally measured friction levels 
shown in Figure 7. The deployment time is approximately 18.5 seconds. 
Additional runs were made attempting to duplicate a deployment similar to 
that experienced on orbit. The runs showed that the deployment rate was most 
sensitive to the friction level on the yoke/solar-array-drive-assembly 
(SADA) hinge. Raising the stiction on the yoke/SADA hinge 50% and the 
yoke/inboard-panel hinge 50% yielded a deplopent similar to that experi- 
enced on orbit. Figure 9 shows these results. The array deploys for approx- 
imately 26 seconds at wfiich point it nearly stops, but velocity data indi- 
cate the array slowly moves for another 8 seconds, at which point the 
deployment stops with the array in a partially deployed position. 
218 
90 
180 
180 
c 
g 180 
t 
P 
2 
2 
* 0 
:: 
P 
8 0 
L3 
0 
0 
T 
5 10 15 
DEPLOYMENT TIrn (seconds) 
30 
90 
LAO 
180 
C 
-80 : 
p 180 
P 
Figure 8. Simulation Results of 
the Deployment of the 
North Solar Array Under 
Nominal Conditions 
1 
DEPIDYWNT TIME (Seconds) 
Figure 9. Deployment Similar to That 
Experienced on Orbit for 
the North Solar Array 
CHANGES TO THE HINGES 
As a result of the findings of the simulation runsand the experiments, 
several changes to the hinges were made. Among the changes are: 
1. Special application of lubricant to all moving parts 
in the hinges 
2. Increased bearing tolerances to allow greater variation 
in temperature 
3. Increased polish on locking bars (see Figure 2) 
4. Increased torsion spring pretorque level (see Figure 7). 
Using the stiction torque level on the modified hinges as given in Figure 5, 
the predicted deployment time decreased to about 12 to 14 seconds for the 
south and north solar arrays respectively. Figure 10 shows the simulation 
results with the modified hinges on the north solar array. 
219 
DEPLOYMENT TIME (SW.) 
Figure 10. Deployment Simulation With the Modified Hinges on the 
North Solar Array 
RESULTS 
The hinge assemblies on the third flight of INTELSAT V incorporated the 
modifications previously described. The satellite was successfully launched 
December 15, 1981. Data from the onboard accelerometer, shown in Figure 11, 
indicate that the first hinge of the south array locked up at 11.8 seconds, 
and the first hinge of the north solar array began lock up at about 13.8 
seconds. 
The accelerometer is located on the spacecraft body, near the north solar 
array. 'Ihe proximity of the accelerometer. to the north solar array makes it 
more sensitive to accelerations of the north solar array. Therefore, 
the small initial acceleration disturbance occurring at 11.76 seconds is 
assumed to be the. south solar array locking up, and the larger disturbance 
at 13.64 seconds is assuned to be the north solar array locking into posi- 
tion. These results are in excellent agreement with the simulation results 
mentioned in the previous section. 
220 
LOCK UP OF 
NORTH ARRAY 
LOCK UP OF 
SOUTH ARMY 
\ 
- 
I 
I 
I  
RELEASE OF I I  
SOLAR ARRAYS OLAR RRAYS I  
I  I' ’ 
r  I  
I  I  
I  I  
I  I  
I  I '  ’
I  I  
1  I  I  I  I  
0  5  10 0 11.76 13.64 
DEPLOYMENT IME (SECONDS) 
Figure 11. Accelerometer Data From FM-3 
:5 
CONCLUSIONS 
The most probable cause of the deployment time difference has been identi- 
'fied as the variation in the friction level in the hinge assemblies at low 
temperatures. Some of the changes to the hinges included special lubrica- 
tion of the guides and bearings and additional polishing of the locking 
bars. Experiments on the modified hinges showed a reduction of the resist- 
ing torque on the hinges by about 50% at low temperatures. When the new 
friction values were put into the computer simulation, the deployment times 
of the south and north solar arrays decreased to 11.9 and 13.7 seconds, re- 
spectively. The deployment times of the south and north solar arrays on the 
third flight of INTELSAT V were 11.8 and 13.6 seconds, respectively, in 
excellent agreement with the theoretical results. These data validate the 
analysis used in the simulation technique described in this paper. Using 
techniques similar to those described herein will provide analysts and 
designers with more accurate simulations and a better basis to evaluate 
potential problems associated with deployment mechanisms. Points to be 
stressed are: (1) adequate theoretical analysis of a mechanism should be 
undertaken, (2) mechanisms should be tested under conditions which duplicate 
the range of expected orbital environments in order to identify sensitive 
conditions, and (3) during the ground testing of very large complete assem- 
blies it is very difficult if not impossible to adequately duplicate the 
orbital conditions. 
* The authors would like to thank Dr. A.K. Panerjee and Dr. P.M. Sarba for 
their contributions to the theoretical simulation. 
** The mrk described herein is based, in part, upon work sponsored by 
INTELSAT. Any views expressed herein are not necessarily those of INTEL- 
SAT. 
221 
P. L. Vorlicek 
Ford Aerospace and Communication Company 
3939 Fabian Way 
Palo Alto, California 94303 
Mr. Vorlicek is currently an engineer in the Systems Analysis Department of 
the Dynamics and Control Section at Ford Aerospace and Communication Company. 
He has received both his B.S. degree, in 1979, and his M.S. degree, in 1981, 
in Aerospace Engineering from Massachusetts Institute of Technology. 
Co-authors of this paper are Mr. J. V. Gore and Mr. C. T. Plescia who are also 
with Ford Aerospace and Communication Company. 
222 


Design and analysis considerations for deployment mechanisms in a space environment

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