Deployment mechanisms were developed to position scientific instruments during probe descent into the Venus atmosphere. Each mechanism includes a provision for pyrotechnic release of the enclosure door, negator springs for positive deployment torque, and an active damper using a shunted dc motor. The deployment time requirement is under 2 seconds, and the deployment shock must be less than 100 g's. The mechanism is completely dry lubricated and constructed mainly of titanium for high strength and high temperature stability. The mechanism was qualified for descent decelerations up to 565 g's and for instrument alignment up to 940 F. The mechanism requirements, the hardware design details, the analytical simulations, and the qualification testing are described

Meadows, F. R.

Miyakawa, R. H.

Townsend, W. L.

English

NASA Technical Reports Server

DEPLOYMENT MECHANISMS ON PIONEER VENUS PROBESX 
By W. L. Townsend, R. H. Miyakawa, 
and F. R. Meadows 
Hughes Aircraft Company 
ABSTRACT 
Deployment mechanisms were developed to position scientific instruments 
during probe descent into the Venus atmosphere. Each mechanism includes a pro- 
vision for pyrotechnic release of the enclosure door, negator springs for posi- 
tive deployment torque, and an active damper using a shunted d-c motor. The 
deployment time requirement is under 2 seconds, and the deployment shock must 
be less than 100 g's. The mechanism is completely dry lubricated and con- 
structed mainly of titanium for high strength and high temperature stability. 
The mechanism has been qualified fgr descent decelerations up to 565 g's and 
for instrument alignment up to 940 F. The paper describes the mechanism 
requirements, the hardware design details, the analytical simulations, and the 
qualification testing. 
INTRODUCTION 
The Pioneer Venus mission includes a Probe Bus that carries three small 
probes and one large probe to Venus for release and descent into the Venus 
atmosphere. Each of the three small probes, as shown on Figure 1, has two 
deployment mechanisms designed to stow scientific measurement instruments dur- 
ing the 565-g deceleration of descent. When the Probe has reached 65 Km 
altitude from the surface, the enclosure doors are pyrotechnically released to 
allow rapid deployment of the Small Probe Net flux Radiometer (SNFR) and the 
Small Probe Atmospheric Structure Experiment (SAS) instruments. Atmospheric 
data are taken for the last 65 Km of descent until the Probe lands. The Venus 
atmosphere is very dense, reaching 1400 psia at the surface, and aerodynamic 
heating causes a maximum temperature of 940°F on the mechanism. 
The generation of the mechanism design requirements has covered many 
variables such as entry angles, spin speed and direction, Probe nutation and 
aerodynamic turbulance torques. For instance, the lumped environmental torques 
could either aid the SAS deployment as much as 3.0 in-lbs or possibly retard 
SAS deployment as much as 8.5 in-lbs. This extreme variation in environmentally 
applied torques meant that a large spring force must be applied to ensure 
deployment, but damping provisions must also be made to prevent instrument 
damage should these environmental extreme forces not be present. 
*This work was conducted under NASA-Ames Contract 2-8300. 
143 
https://ntrs.nasa.gov/search.jsp?R=19790013194 2020-03-20T18:32:55+00:00Z
In addition to the deployment requirements, the instruments have to be 
protected structurally and thermally during the deceleration prior to deploy- 
ment. The SAS and SNFR instruments are stowed aft of the conical heat shield 
to avoid temperatures of several thousand degrees during initial entry, A full 
enclosure, which is covered with silicone rubber ablative material, is still 
required for mechanism protection. 
The six mechanisms are now installed on the small probes and are under- 
going system testing. The launch of the Pioneer Venus multiprobe mission is 
planned for August 1978. 
REQUIREMENTS 
The mechanism design requirements can be divided into three basic modes 
of operation as listed below. Table 1 lists the specific parameters required 
of each mode. 
1. 
2. 
3. 
Stowed Mode. The instruments are kept folded into a 
retracted position by the enclosure door. While in the 
folded condition, the mechanism is subjected to vibration 
and entry deceleration simulations up to 706 g's for 
qualification. 
Deployment Mode. The enclosure door is pyrotechnically 
released by a bolt cutter. As the door is sprung open, 
the instruments must be deployed in less than 2.0 seconds. 
However, if the deployment is too rapid, the stopping shock 
must not exceed 100 g's at the instrument tip. 
Descent Mode. The mechanism must withstand the high pressure 
and temperatures up to 940°F while maintaining alignment 
of the instruments within ,+ 1 degree. 
MECHANISM TRADEOFFS 
The deployment time requirement of less than 2.0 seconds along with the 
deployment shock restriction made the possible mechanism approaches very 
limited. An undamped spring deployment design showed impact shocks above 400 to 
500 g's. In addition, the rebound problems with the undamped spring approach 
could only be solved with a deployed latch which had difficult dynamic and 
strength requirements. 
An active d-c brush motor could be sized to drive the mechanism but 
requires active rate feedback to fall within the impact requirement. The rate 
feedback closed-loop electronics for each of six mechanisms makes the cost and 
weight penalties much higher than the baseline. The more simplified stepper 
motor driver electronics are more attractive for rate control than the d-c 
144 
motor system, but the stepper motor would be much heavier. A heavier stepper 
motor is required because the 2.0 second deployment time means a small gear 
ratio and high motor torque for control, 
The damped spring approach could use viscous fluid, rubbing friction, or 
electromagnetic damping. The nearby science payloads along with the wide 
temperature excursion ruled out the viscous damper because of concern for out- 
gassing. The electromagnetic damper was selected over the friction damper due 
to the fine tuning capability and consistency of the d-c motor. The d-c motor 
can be fine tuned using a shunt resistor to allow for changes in requirements or 
variations from unit to unit. 
DESIGN DESCRIPTION 
The deployment mechanisms and protective housings are designed to accommo- 
date both the SNFR and SAS instruments interchangeably. There are differences 
in instrument attachment, interconnections, and instrument aerodynamic shapes. 
The mechanism has different deployment springs, different stop angles, and damper 
settings for each type of instrument. The diagram of Figure 2 describes the 
mechanism with the SAS in the deployed position. 
The mechanism includes a machined structural base that supports both the 
mechanism and the protective housing. The protective cover is deployed 90 
'degrees by a torsional spring after the tie-down bolt is cut by a pyrotechnically 
driven device. The squibs and squib drivers are redundant. The instruments 
are deployed by a negator spring once the door motion has started. The door is 
deployed in 0.1 to 0.2 seconds which is more rapid than the instrument motion. 
As a backup, there is a mechanical finger on the instrument platform to safely 
push the cover ahead in case of interference. The cover support is captured by 
the wedging action of a C-shaped clamp to absorb excess energy and to prevent 
rebound. 
The instrument deployment is driven by a 301 CRES negator spring selected 
for its constant torque properties. The damping is provided by a d-c brush 
motor geared to a higher speed by a 38.3:1 ratio. The motor is mounted on ball 
bearings to preserve the air gap, but dry-lubricated journal bearings are used 
elsewhere. The titanium gears have Vitrolube 1220 MoS dry lubricant which 
was selected for the binder cure temperature of 950°F $0 minimize further out- 
gassing during descent. There is no latch in the deployed position. A canti- 
levered beam provides a spring action stop that is rigid enough to maintain 
alignment after a short settling time. 
The photographs of Figures 3 and 4 show the mechanisms in the deployed 
condition with the housings removed. The instrument deployment negator spring 
and the cantilevered beam stop are depicted in Figure 3. The base structure 
is machined 6AUtV titanium because of the high strength requirements during 
entry and thermal expansion compatibility with the mating probe structure. 
Beryllium copper is the shaft material selected to act as a journal bearing 
and for shock absorbing properties, since it has a low modulus of elasticity 
145 
compared to steel. The mechanisms' physical characteristics are shown in 
Table 2. 
ANALYTICAL MODEL 
The flight loading environment at the time of deployment, as depicted in 
Figure 1, is severe and difficult to predict. Trajectory dependent and hence 
time varying aerodynamic conditions are primarily within the transonic region 
characterized by bow shocks and complex local flow f.ields. Attitude motion of 
the probe generates time varying inertial loads in addition to axial decelera- 
tion. Asynnnetric ablation on the aeroshell can induce vehicle spin of up to 
100 rpm resulting in centrifugal force fields. In addition, the mechanism per- 
formance is dependent on design parameters such as friction, spring tolerances, 
cable bending effects, and motor damping which are all in varying degrees 
environmentally dependent. 
The time sequencing and complex interaction of both environmental effects 
and dispersed design parameters necessitates the development of analytical 
models in order to assess mechanism performance and structural integrity. A 
computer program was written incorporating both probe and mechanism dynamics 
for time simulation of the rigid body aspects of the deployment process. A 
finite element model and standard modal analysis techniques are employed to 
determine structural response at impact. Parametric studies are utilized to . 
establish the combination of extreme conditions which result in worst case fast 
and slow performance as summarized for the SAS in Table 3. 
The level of torque for the negator deployment spring is selected to 
ensure positive torque margin throughout deployment without benefit of momentum. 
In other words, should the instrument momentarily stall at any intermediate 
angle, the negator must have sufficient torque to restart. The significant 
torque variables are displayed on Figure 5 for conditions influencing the SAS 
deployment. The static torque margin shown on Figure 5 never falls below zero. 
The resulting fast and slow deployment times are graphically shown on Figure 6. 
The computer results are shown in Table 4 depicting the wide spread in all 
parameters between fast and slow deployment conditions, 
The cantilevered beam used as a spring stop is necessary because the tip 
accelerations during stopping exceed 100 g's even with an active damper. A key 
contributor to the stopping acceleration is the shock produced by ongoing motor 
rotor energy after the instrument reaches the end of travel. Therefore, the use 
of higher damping coefficients cannot directly solve the problem without far 
exceeding the 2.0 second limit in slow deployment cases. Figure 7 describes the 
relationship between stop spring stiffness and tip acceleration for the fast 
deployment case. The figure also shows the overtravel tradeoff which is an 
important consideration for enclosure clearance. 
146 
TEST PROGRAM 
A comprehensive test program was conducted on a prototype unit to cover 
all aspects of environments, including high-g, vibration, deployments under 
many environmental simulations, and descent temperatures. Two of the six flight 
units were tested to qualification levels of acceleration, temperature and 
vibration. The remaining four units were subjected to acceptance levels of 
acceleration, temperature, and vibration. The acceptance criteria were in 
terms of deployment time, deployment shock, and deployed alignment measurements. 
The pyrotechnic bolt cutter release was done once in development and four times 
in qualification to ensure compatibility with the mechanism, In order to save 
on the cost of the non-reusable bolt cutters, manual releases were done for 
most deployment tests. 
The high-g testing turned out to be most revealing in discovering design 
deficiencies. Several cases occurred where structural distortion at 706 g's 
caused mechanical contacting of delicate parts not designed to carry loads. 
These areas had to be reinforced with structural stiffeners along with better 
supports for electrical wiring. During the first high-g test, the free end of 
the negator spring was driven off its support post so that subsequent deployment 
could not occur. A special hook was designed to prevent this unacceptable 
negator motion during high-g forces and to spring away during normal deployment. 
The probe descent simulation of the high pressure and 940°F temperature 
was conducted on the development model only because of permanent damage to some 
parts of the system. Since it would be very difficult to measure instrument 
angle change during exposure, measurements were made on a fixture before and 
after exposure. Unfortunately, the instrument angle changed about 1.5 degrees 
as a result of this test. Subsequent testing was conducted to isolate the 
problem to permanent set of the beryllium copper cantilever stop spring shown 
on Figure 3. It was discovered that beryllium copper creeps to a new permanent 
set when exposed to 940°F under the residual preload of the negator deployment 
spring (approximately 15 in-lbs). The cantilever stop spring was then changed 
to a 17-4 PH CRES design with the same spring rate. The mechanism exposure was 
rerun with a small acceptable change in alignment of only 0.035 degree. 
The deployment tests were conducted at high and low temperatures, but most 
data were collected at room ambient conditions. A loading fixture was designed 
to apply either aiding or restricting torques to simulate the extremes of the 
environment. Provision was made to add a shunt resistor to the motor to trim 
the extent of damping of each unit. As it turns out, a 5 ohm resistor was 
suitable for all units. The results are summarized in Table 5. It should be 
noted that the data show results that are less than one-half the allowable 
specification limits in time and g-loading. The analytical results showed a 
much wider dispersion predicted at Venus because the aerodynamic induced 
torques as shown on Figure 5 have a more severe effect than the linear simu- 
lation used in test. The computer simulations were rerun using the laboratory 
induced torques and verified this difference. 
147 
CONCLUDING REMARKS 
The development of these mechanisms shows a high degree of sophistication 
and confidence in performance predictions compared to just a few years ago. For 
instance, the use of dry lubrication on gears, ball bearings, journal bearings, 
and motor brushes is now well enough defined to predict friction and life 
instead of just survival in space. The use of the computer simulations of 
mechanism behavior now allows all variables to be studied so that the testing 
matrix can be very much reduced. However, it continues to be important to 
touch on each test environment to check the level of accuracy of the computer 
simulations and be sure that some critical condition has not been overlooked. 
Table 1. Requirements summary for qualification 
Launch Mode (Stowed) 
Sine vibration - lateral 
thrust 
Random vibration 
Acceleration 
Temperature 
Deployment Mode 
Deployed angle - SAS 
SNFR 
Peak tip acceleration 
Deployment time (including door opening) 
Overtravel allowance 
Temperature 
Acceleration 
External torques - SAS 
(lumped) SNFR 
Deployed Mode 
Deployed angle error 
Temperature 
45 g 45-60 Hz 
30 g 25.5-100 Hz 
12 g rms overall 
706i 
-98 F to +170'F 
160° + lo 
120° T lo 
100 g- 
2 0 
56 
seconds 
-98OF to +123'F 
5.5 g 
13.0 in-lbs peak 
8.5 in-lbs peak 
,+ lo 
-98'F to +940°F 
148 
Table 2. Mechanism physical characteristics 
Enclosure structure 6AL 4V titanium with silicone 
rubber ablative material 
Enclosure size (inches) 4.2 W x 9.4 L x 5.8 H 
Weight summary (lbs) 
Enclosure 
Base structure 
Mechanism parts 
Instruments 
Total 
SAS SNFR 
2.01 2.01 
1.04 1.04 
1.71 1.74 
0.29 0.80 
5.05 5.59 
Instrument connections 
SAS 6 electrical wires 
1 bellows 
SNFR 9 electrical wires 
Damper motor 
Weight (lbs) 0.15 
SLaLLc LUGS) n 96 
Damping (in-lb-sec/rad) 1.10 
Motor resistance (ohms) 6.5 
Negator spring torque 
SAS (in-lbs) 19 2 1 
SNFR (in-lbs) 16 2 1 
149 
Table 3. Definition of worst-case deployment conditions 
Trajectory and 
aerodynamics 
Angle of attack 
Attitude dynamics 
Negator spring 
Motor damping 
Temperature 
FAST 
SAS 
yE 
= -2(-p (1) 
Maximum aiding aero. 
-12O 
Outboard = +0.3 g 
Inboard = -1 g 
20 in-lbs 
1.5 in-lb-sec/rad 
123'F 
yE 
= -750 
Maximum retarding aero. 
+12O 
Outboard = -0.3 g 
Inboard = -1 g 
18 in-lbs 
5.2 in-lb-seclrad 
-103'F 
(1) yE is the probe trajectory entry angle relative to Tenith. 
(2) Aiding and retarding aerodynamic torques have 30% dispersion margin. 
Table 4. Performance results of the computer simulation 
DEPLOYMENT PARAMETERS 
Peak Elapsed Tip Alignment 
Accel. Time Overtravel Error 
CONDITIONS (g's) (set> Cd%> Cd%) 
* Fast 51 0.24 2.8 -0.005 
Slow 6 1.52 0.3 +0.14 
SNFR Fast 75 0.23 3.0 -0.05 
Slow 8 1.29 0.3 $0.1 
Specification maximum 100 2.0 5.0 1.0 
150 
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PROBE ATT1 
DYNAMICS 
FORCES 
SAS 
TUDE CENTRIFUGAL 
FORCES 
DECELERATION 
FORCE FIELD 
ANGLE OF ATTACK 
PROBE VELOCITY 
Figure 1. Pioneer Venus Small Probe 
PYROTECHNIC 
jBOL T CUTTER 
BASE STRUCTURE 
DC MOTOR 
DAMPER 
ENTRY PROTECTIVE 
SPRING 
COVER DEPLOYMENT 
ENTRY 
COVER 
GEARING 
DEPLOYMENT 
NEGATOR 
PROTECTIVE 
Figure 2. SAS Deployment Mechanism and Cover 
152 
Figure 3. SAS Deployment Mechanism 
., 
,~:(~ ,’ cinMP ~-0 C~TURE “, ;CdVER 
DEPLOYED COVER /SUPPORT ” 
’ sUppoRT \ 
b 
, ” NiGAT6k 
DEPLOYMENT 
BUMPER 
‘-/ 
&STRUCTURE,. -A 
‘N’ET FLilX * 
RADIOMETER ,. 
‘PROBE 
Figure 4. SNFR Deployment Mechanism 
153 
I NEGATOR SPRING 
15 
t. 
c----------- 
I 
-*, STATIC 
I TORQUE / 
lo i 
t 
I 
I 
-15- DYNAMICS 
-201 1 I I I I 
0 30 60 90 120 
DEPLOYEMENT ANGLE, DEG 
Figure 5. SAS Minimum Deployment Torques 
90 - 
60 - 
30- 
o- 
I 
30, 
0 
MAXIMUM 
DEPLOYMENT 
CONDITION 
DEPLOYMENT 
CONDITION 
“..A I .Y 
TIME FROM PYRO INITIATION,SEC 
Figure 6. SAS Deployment Time History 
154 
5’ 
4’ 
*BASELINE = 4500 IN-LB/RAD 
Sk., 
I 
- 
-- 
-9 
I 
--- 
4”“” 
STOP STIFFNESS, IN-LB/RAD 
Figure 7. SAS Impact and Overtravel Response 
155 


Deployment mechanisms on Pioneer Venus probes

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