A compact viscous rotary actuator/damper for use on the Mariner '71 and Viking Programs was developed. Several functions were combined into this single mechanism to control the deployment, latching, and damping of the solar panel arrays used on these space vehicles. The design, development, and testing of the actuator/damper are described, and major problems encountered are discussed

Harper, J. D.

English

NASA Technical Reports Server

6:,Z8291
VISCOUS ROTARY VANE ACTUATOR/DAMPER_'
By JACK D. HARPER
MARTIN MARIETTA CORP.
SUMMARY
3PL has developed a compact viscous rotary actuator/damper for
use on the Mariner '71 and Viking Programs. Several functions have been
combined into this single mechanism to control the deployn=ent, latching
and damping of the solar panel arrays used on these space vehicles. The
design, development and testing of the actuator/damper are described, and
major problems encountered are discussed.
INTRODUCTION
The Jet Propulsion Laboratory's Mariner Mars spacecraft, launched in
1971, and the Viking Orbiter spacecraft, launched in mid 1975, required the
use of panels of solar cell arrays for electrical power generation. Because
of the size of these solar panel arrays, they were folded parallel to the space-
craft's longitudinal axis to fit within the launch vehicle shroud. After launch
and shroud removal, they were rotated to their flight position. Deployment
of the solar panels from their launch position, latching of the solar panels
in their proper flight position and damping of solar panel perturbations
caused by the spacecraft engine firings were functionally controlled at the ,.
hinge line of the panels. These three main functions were incorporated into
a single unique fluid-filled viscous rotary vane actuator/damper. The con-
cept was originally developed for Mariner '71 and then modified and
improved for use on the larger solar panels on Viking Orbiter. This paper
will deal primarily with the design configuration flown on the Viking Orbiter
spacecraft.
DEPLOYMENT AND LATCHING
Deployment of the solar panels is accomplished by the release of
stored energy from a torque spring. This was a clock spring design on
Mariner '71 and later modified, because of weight and reliability consider°
I ations, to a pair of constant torque multi-leaf springs for Viking. The
torque level for the deployment springs was set by a design goal of main-
taining a torque margin of four times the maximum resistive torque.
i
This resistive torque comprises the friction of the hinge joint, the tare
torque of the deployment device, plus the items that bridge the hinge joint
such as electrical wire bundles, attitude control gas lines and the coaxial
#
This paper presents the results of one phase of research carried out at
• the Jet Propulsion Laboratory, California Institute of Technology, under
Contract No. NAS 7-I00, sponsored by the National Aeronautics and Space
Administration.
198 JPL Technical Memorandum 33-777
] 97602 ] ] 84-205
https://ntrs.nasa.gov/search.jsp?R=19760021203 2020-03-20T02:33:36+00:00Z
cable for the relay antenna. These torque values and the deployment torque
margin can be seen in Figure 1. The variations in available deployment
torque are from the mounting of the actuator; its crank and linkage to the
solar panels form a four-bar linkage that has an increasing mechanical
advantage that approaches one to one at the deployed position.
Under free running conditions, this excess torque margin would drive
the solar panel assemblies too fast and cause high shock loads to the
system at latching; therefore the springs are configured to drive the
panels in parallel with the fluid actuator. This causes compression of
the fluid between the vanes of the actuator pumping the fluid through the
orifice gap formed between the s,ationary vanes and the outer rotating
wall. This can be seen in the cros_ section of Figure 3. The actuator
in this case acts as a rotary rate-:imiting device that is velocity sensitive
and will, for a constant torque input, maintain a constant deployment
velocity. The device can be adjuster: to control the rate of deployment
independently of the input driving torq,le. The input torque is based on
the resistive torques and the required torque margin; then the fluid
viscosity and the orifice gap are selected to balance the input torque for
the deployment rate desired.
After deployment, the solar panels are latched in their flight position
by either or both redundant lock pins located in the rotating vanes. The
lock pins are spring loaded taper pins that engage conically reamed holes.
This configuration allows for a gradual entry of the lock pins into the
holes and near-zero backlash when engagement is complete. When coupled
with the controlled deployment velocity, this arrangement provides for small
or negligible impact loads at latching. ,_
To insure that the soiar panels are in the proper position at latching, •
the solar panels must be aligned to actuatorts locking pin locations.
This adjustment is made at the flight assembly by means of the turnbuckle
linkage assembly connecting the panels to the rotating portion of the
actuator. This linkage and its attach points are shown in Figure 2.
DAMPING
After deployment and latching, it was necessary to have a damperg-
_: in the system to minimize any interaction between solar panel perturbations
and the spacecraft's attitude control system (ACS). The requirements to _ o
decouple the solar panel assemblies _ resonant response from any known
pulses of the spacecraft ACS system were determined to be a minimum i
undamped natural frequency of .5 Hz and at least 30¢/0 damping. To meet the
damping requirements, the actuator/damper had to have a dual spring rate,
one for deployment and one, much higher, for system frequency control.
This was accomplished in the device by the action of the latching pins. As
the pins stop the system deployment, they lock out the deployment springs
and couple the actuator/darnperts outer rotating body to the inner stationary
base through the center shaft, which is a tuned torsion bar. The torsion bar
had to be configured stiff enough to meet the frequency requirement but flex-
ible enough to allow sufficient panel rotation at engine firings and shut-off
for the damper to be effective. For the Viking solar panel assemblies, the
3PL Technical Memorandum 33°777 199
1976021184-206
I _ 1 ¢
torsion bar provided an undamped natural frequency of ,87 Hz. The flexing
of the torsion bar allows relative motion between the damper vanes, displac-
ing fluid through the orifice gap. It is the work required for this i luid dis-
placement that damps the solar panel vibrations.
The damping of a rotary device like this is dependent vpon the
geometry of the vane configurations, the working fluid viscosity and the
orifice size and shape. The parameters easiest to adjust to meet the damp-
ing requirements are the fluid viscosity and the orifice size. To meet the
clamping requirements of the Viking spacecraft with its larger, heavier solar
panels as compared to the solar panels oa Mariner w71, it was necessary to
increase the working fluid viscosity and decrease the orifice gap. The com-
bination finally selected was a dimethyl silicone oil, because oe its relative
flat viscosity vs. temperature curve, with a kinematic viscosit V of lv?,,000
centistokes and an orifice gap of 0.015 in. With this combination, the
damper was ab!e to provide the required 30% damping over the predicted
temperature range of -10°F to +l15°F, with damping ranging from 30% to
53%. Although this fluid and orifice size did provide the needed damping,
they presented a problem to the deployment function of the actuator/damper.
With the input torque of the deployment springs set and this high fluid viscos-
ity and small orifice, the deployment time became unacceptably long. To
shorten the deployment time, a step, shown in Figure 3, was added to the
inside wall of the body to increase the orifice gap. This gives a faster rate
at the beginning of the deployment and a gradual decrease to the slower rate
about halfway through the deployment.
Another problem that had to be considered is the fluid expansion and
contraction with temperature. At high temperatures, the fluid e..pansion ."_,
causes high internal pressure, increasing the possibility of leakage or a
structural failure of the pressure vessel. At low temperatures, the fluid
contraction can cause a void within the fluid. This vacuum "bubble" allows
: relative vane motion without forcing the fluid to be pumped through the
orifice and greatly reduces the damping capabilities of the device. To pre-
vent these problems from occurring, a temperature compensator was used
to provide an additional spring-pressurized reservoir that supplies the
expansion volume needed at high temperatures and the make-up fluid needed :
at low temperatures. The temperature compensator, shown in Figure 3, ,_
consists of a sealed piston, a compression spring of conical washers and a
small orifice connection to the working fluid chamber.
?
:
DEPLOYMENT TESTING
For the deployment testing, the solar panel assemblies were position- "_
,: ed with the hinge axis vertical to minimize the gravitational effects. All
i of the development testing was performed using a pair of simulated solar :
panels shown in Figure 4. Deployment tests were conducted to check
the actuatorts deploying, rate limiting and latching functions. Because
: the Viking system was a double folded panel pair, with the outboard panel
i deployed and latched during the deployment of the inboard panel, the effect _
of this two-panel deployment was investigated during the deployment and
latching tests.
200 SPL Technical Memorandum 33-777
1976021184-207
1
I J !
The flight solar panel structure was not capable of supporting the
cantilevered weight of *he outboard panel, so the panels could not be free-
deployed in the flight-like manner. To minimize the complexity and cost of
the testing equipment and the risk to flight hardware, it was decided to
deploy the flight panels in teps using a mid-h;'-z _.o_ support tor the deploy-
ment of the outboard ,* _'r.an,.,. Testing of the flight panels is shown in Fig-
ure 5. By running tests with both the simulated panels and the flight panels,
a collation of the deployment times was assembled for evaluation of pre-
- flight checks and a prediction of flight deployment times.
DAMPING TESTING
Damping tests were performed on the test setup shown in Figure 6. i :
It consisted of a simulated solar panel assembly, the actuator/damper in
an insulated box shroud and a liquid nitrogen/hot air temperature control _
system. To simulate the maximum .16 g force on the solar panels from a _
main engine firing, the simulated panel and actuator/damper are assembled ]
in the deployed position (latch pins engaged) and mounted at an angle of 9 ° ! _
from vertical When released from this position, the horizontal component _
of the weight will approximate the load the torsio_l bar and damper will see _ •
in flight.
All dampers were tested in this manner at various temperatures _ :
from -19°F to 124°F, with the damping ranging from 30% to 53% for all
fligh_ units. A typical damping curve taken at room temperature is
represented in Figure 7. _, .'
Certain problem areas were identified from the development and
acceptance tests conducted. Early damping testa showed that the tempera- _ ,.
ture compensator must be decoupled from the damper or the pressure on i _,
the fluid would force it into the compensator and not past the damper vanes. ._
This was ¢.orr_cted by making the fluid passages between the compensator i :
and darr,/_e.- a pair of .006-in diameter holes. A suddenly applied load 4
would not drive fluid through this small orifice, but a gradual change in
temperature would allow fluid to flow between the damper and compensator,
Also the device had to be pressure filled to a level that would remove any
axial play within the umt and completely seat the "0" ring seals so that they i
would not move or "breathe" when a load was applied. _ ;
_ VIKING FLIGHT PERFORMANCE _
_ During the launch and early flight of the Viking missions, data was _ i
_ transmitted that allowed evaluation of the flight performance of theactu,ttor/damper assemblies, With four assvmhlies per spacecraft and _ i
two successful launches, the data received gives a good indication of the _
_ actuator/damperWs flight characteristics. The data gave direct readings
<: for deployment times and latching but not for the damping function; this
, _ was interpreted from data of other spacecraft subsystems.
_'!__ 3"PL Technical Memorandum 33-777 ZO 1
1976021184-208
Solar panel deployment times were estimated, based on pre-flight
testing and the predicted flight temperature range, to be between Z minutes
to 2 1/2 minutes, with the bay 1 panel being the slowest because of increased
inertia due to a relay antenna attached to the outboard panel.
The last ground tested deployments and flight actuals are shown
below:
TABLE I. VIKING SOLAR PANEL DEPLOYMENT TIMES
Launch A Launch B
Location Last Preflight Time +6.7 sec Time +6.7 sec
Test -0.0 sec -0.0 sec
|
Bay 1 2 rain Z. 8 sec 2 rain 11 sec
Bay 5 I rain 39 sec 2 min 2 sec
Bay 9 1 rain 43 sec I min 55 sec
Bay 13 1 rain 39.7 sec I rain 55 sec
Bay 1 1 min 49 sec Z rain 10 sec
Bay 5 I rain 48 sec Z rain I sec
B_y 9 1 rain 41 sec I rain54 sec
Bay 13 1 rain 39.5 sec 1 rain 47 sec
It can be seen from the flight deployments that all actuators function-
ed quite well, giving a uniform deployment within a small time variation.
Actual flight deployment times were slightly faster than predicted because
of the fluid's sensitivity to temperature. The flight devices were close
to nominal temperature, and the predicted times were based on the low
tempe rature predict s.
Based on the ACS gas usage, the actuatcr/dampers functioned
properly, preventing any adverse interaction between the solar panel
natural frequency and the ACS gas jet firings. Confirmation of the
latching and damping functions of the devices was given by the midcourse
engine firing in that there was no discernible c, g. shift, indicating a free _
or non-latched panel; in addition, the ACS system showed normal operation ;
after the engine firing, i ;
{
/
w
Z0Z 3PL Technical Memorandum 33-777 t,
1976021184-209
!CONCLUDING REMARKS
This viscous rotary damping device has been developed, tested and
successfully proven on two spacecraft programs. ']:he major unexpected
phenomena encountere, _. with the design and their _olutions have been
discussed; however, there has been no attempt to describe all of the
problems, pitfalls and agonies encountered with a design effort of this
nature from concept to a proven flight-ready mechanism. Because of
the nature and complexities of this type of fluid actuator/damper, the
design does not come only from careful analysis but must be gained from
• testing and experimentation. This device has the advantage of being
designed and developed for one program, then re-evaluated and expanded
to requirements greatr-r than the original design goals. This additional
work and testing have increased the knowledge of the deviceWs performance
characteristics and capabilities.
Because of the devicets versatility, it has been considered for many
applications where panels, booms or other appendages have to be deployed
and/or damped in a space environment. It is currently being used on 5PLVs
Mariner J'upiter/Saturn Program and Rockwell InternationalVs Global
Positioning Satellite. It has also been included by Martin Marietta and
General Electric Space Division in proposals for other spacecraft
and satellites.
i k;
i i_ JPL Technical Memorandum 33-777 203 )
1976021184-210
80 I I
TORQUE AVAILTORQUE FACTOR =
MIN TORQUE TORQUE REQ'D
AVAILABLE 60
60-_ T.F. _ =3.3417.988
AS DESIGNED 2-SPRING SETS
T.F. 30 = 1.67
17.988 --
L FAILURE MODE I-SPRING SET
,c_
_T 4O
:3 MIN TORQUE
AVAILABLE
, O I-SPRING SET_
TORQUE REQ'D
I
oi. I I
0 45 9O
ANGULAR POSITIONe cleg
Figure 1. Input Torque Available and Resistive Torques
Compared to Panel Angular Position
_wrl|w "
; _ Y,N(.,?'-
.i
1
1Figure 2. Actuator/Damuer. Turnbuckle Linkage and
Solar panel Attach Bracket !
1
Z04 JPL Technical Memorandum 33-777
!
1976021184-211
"TPL Technical Memo1"andurn 33-777
Figure 4. Deployment Test Setup With Simulated Solar Panels
p*
J
Figure 5. Final Testing of Flight Solar Panels and Actuator/Damper i i
t
206 JPL Technical Memorandum 33-777RI_0DUCIBILrrYOFTHE
ORI0iNAL PAGE 18 POOR I
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1976021184-213
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JPL Teehnleal Memorandum 33-777 Z07 ]
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1976021184-214


Viscous rotary vane actuator/damper

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