A three segment, 15-foot boom mechanism was developed to deploy magnetometers from the Pioneer Venus orbiter spinning shelf. The stowage mechanism is designed to contain the magnetometers during launch and to deploy these instruments by centrifugal force upon pyrotechnic release. Unique graphite-epoxy boom segments are used for a lightweight design with sufficient strength to withstand a 7.5 g orbit insertion force while extended. The detailed design is described, along with the test methods developed for qualification in a one-g field

Townsend, W. L.

English

NASA Technical Reports Server

N 79 -21 377
MAGNETOMETER DEPLOYMENT MECHANISM FOR PIONEER VENUS*
By William L. Townsend
Hughes Aircraft Company
ABSTRACT
A three segment, 15-foot boom mechanism was developed to deploy magnetometers
from the Pioneer Venus orbiter spinning shelf. The stowage mechanism is designed
to contain the magnetometers during launch and to deploy these instruments by
centrifugal force upon pyrotechnic release. Unique graphite-epoxy boom segments
are used for a lightweight design with sufficient strength to withstand a 7.5 g
orbit insertion force while extended. The detailed design is described along
with the test methods developed for qualification in a one-g field.
INTRODUCTION
The Pioneer Venus 1978 mission includes magnetometers on the spinning orbiter,
Figure I, to map the magnetic fields around Venus. These measurements require
a 15-foot separation from the edge of the spinning shelf in order to minimize
spacecraft magnetic effects. A 2-axis magnetometer is placed at the 15-foot
position, and a single axis magnetometer is located at a lO-foot position for a
magnetic gradient measurement.
The Magnetometer Deployment Mechanism (MDM) supports the magnetometers in a
three-segment folded condition for launch, spacecraft separation and spin-up.
The mechanism is pyrotechnically released to deploy the magnetometers approximately
four hours after launch. Figure 2 shows the MDM in a partially deployed position.
Early deployment is necessary for mass balance of the spinning shelf and to permit
magnetometer calibration while still in the known magnetic field of the earth.
The MDM reliability must be emphasized since failure to deploy will leave the
spacecraft with a wobble angle unacceptable to most of the scientific instruments.
When the spacecraft arrives at Venus, a 7.5 g orbit insertion maneuver is
required while the MDM is locked in the extended position. In addition, the spin
stabilization during orbit insertion requires that the MDM deflection be limited
so that spacecraft wobble angles are not excessive. This combination of strength
and deflection requirements dictates a boom designed mainly for stiffness rather
than material strength.
Once the spacecraft is in Venus orbit, the MDM must maintain magnetometer axes
alignment within I° of the prelaunch condition. The error must be allocated to
include thermally induced distortion, latch repeatability, initial alignment
uncertainty and elastic deformation due to centrifugal forces.
* This work was conducted under NASA contract NAS 2-8300
23
https://ntrs.nasa.gov/search.jsp?R=19790013206 2020-03-20T18:55:47+00:00Z
DESIGNREQUIREMENTS
Themission profile places design requirements for the MagnetometerDeployment
Mechanism(MDM)into the following four distinct modes. Values are listed in
Table i.
I) Launch- The MDMmust be folded and stowed to fit on the edge of the
spacecraft shelf and to be within the iii inch shroud diameter. The stowage
system must limit launch accelerations on the magnetometerinstruments to 22.5 g's.
2) Deployment - The MDM stowage mechanism must be designed with redundant
pyrotechnic release for reliability. The MDM must be centrifugally deployed from
the shelf while spinning at approximately 6 rpm. The mechanism must be designed
with high reliability to latch after extension and with sufficient strength margin
to absorb excess latching energy.
3) Orbit Insertion - The extended MDM must withstand 150% of the 7.5 g
orbit insertion forces without damage and must be stiff enough to limit tip
deflection to 10.5 inches when subjected to 7.5 g force.
4) Orbit Operation - The MDM must be designed using magnetically clean
materials to limit magnetic effects to 0.04 gamma at the tip. The magnetometer
alignment must remain within I° of initial installation.
TABLE i
REQUIREMENTS SUMMARY
Launch Mode (Stowed)
Sine Vibration Thrust
Lateral
Random Vibration
Natural Frequency
Magnetometer Max. Acceleration
Deployment Mode
Spin Speed
Release
Deployment Shock
Temperature
Orbit Insertion (Deployed)
Spin Speed
Insertion Acceleration (-/- to boom_)
Natural Frequency
Deflection Maximum
Temperature
Orbital Operation (Deployed)
Spin Speed
Magnetics
Temperature
Mechanism Boom Length
Alignment Accuracy
Weight Deployable Portion
Non-Deployable Portion
5 to 2000 Hz
5 to 2000 Hz
20 to 2000 Hz
Total
11.5 g Max.
5.5 g Max.
11.5 g - RMS
> 4.0 Hz
< 22.5 g's
6.0 to 7.0 rpm
Redundant pyrotechnic devices
< 8 g's
0 ° to 100°F
30 rpm
7.5 g's(ll.25 g design)
> 1.0 Hz
10.5 inches at tip
+10OOF to + 150°F
up to 65 rpm
< 0.04 gamma
-200°F to + 150°F
> 15.7 ft
+ 1.0 degree
I
10.36 ibs
6.08 ibs
16.44 Ibs
24
In addition to the mission requirements, a ground test requirement was
imposedto permit deploymentdemonstration during system test. Although the
ground test imposesgravity induced frictional forces and aerodynamicforces not
encountered in orbit, it does provide a set of data for analytical comparisonto
orbital conditions.
BOOMDESIGN
As can be seen on Figure 3, the boomsegmentsare tapered tubular membersfor
lightness, high stiffness and strength. The design loads including the 50%margin
above7.5 g orbit insertion modeare as follows.
a) Inboard Hinge 892 ft-lb
b) Center Hinge 380 ft-lb
c) OutboardHinge 94 ft-lb
In addition to orbit insertion loads, the boomsegmentsmust be designed
to absorb the residual energy for the worst case latching condition. For example,
the middle segmentmust absorb 7 foot-pounds of excessenergy. There is also a
torsional strength requirement becauseof massasymmetryof the hinges and
magnetometers.
Graphite-epoxy material was selected for the boomsegmentsbecauseof its
low density, high modulusof elasticity, magnetic cleanliness and ease of
fabrication in a tapered tubular form. Beryllium was seriously considered but
could not be easily obtained with sufficient magnetic cleanliness. Fiberglass-
epoxy and aluminumtubing have structural properties significantly inferior to
those of graphite-epoxy. The extent of the taper is dramatic with a 4.125 inch
diameter at the root and 1.5 inch diameter at the tip. The tube ends are rein-
forced by bonding into the sockets of the titanium hinge fittings. Intermediate
load points are reinforced by fiberglass bulkheadsbondedinside the segmentsand
fiberglass rings bondedoutside the segments.
HINGEANDLATCHDESIGN
The engineering modelhinge and latch are shownon Figure 4. Thehinge pin,
a beryllium copper cylinder, forms a journal bearing with the 6AL-4Vtitanium
hinges. Redundantjournals are provided by a sliding fit on each hinge half.
There are redundant beryllium copper locking pins spring loaded against the
titanium semi-circular tracks. The half-hinge and track are machinedfrom one
6AL-4V titanium piece. In order to meetredundancyrequirements, each locking
pin has sufficient strength to withstand all loading conditions.
The preloaded locking pins have the disadvantage of friction drag along the
tracks throughout deployment. Onthe other hand, this design insures that the
pins remain cockeduntil the proper latch position is reached. The drag friction
torque is constant throughout deploymentand is predictable using phenolic bonded
MoS2 dry lubrication. Wearduring ground test wasshownto be of little concern
since an excess of i00 deploymentcycles causedno significant changein friction
torque.
25
The locking pin is designedwith a tip radius to ride on the track, with a
lead-in tapered section, and with a final cylindrical section to absorb stopping
forces without being driven back out of engagement. The tapered section provides
a 4 degree pull-in range to encompassa possible offset angle should the hinge
stop rotating prior to full closure. Whenthe locking pin is fully engaged, it is
spring preloaded into a tapered socket for a repeatable latch position. Latch
reliability is also enhancedby using a preload to impart locking pin acceleration
of 50 to i00 g's. Even in the highest conceivable hinge closing rate of 300
degrees/second, the locking pin fully seats prior to overshooting the socket
centerline.
The hinge shownon Figure 4 is typical of the middle and outer hinge designs.
The root hinge as seen in Figure 3 has a cantelevered hinge pin to be compatible
with the shelf structural load path. The locking pins are designed as described
aboveexcept the latching action is parallel to the hinge.
ELECTRICALCABLEDESIGN
The electrical cable is a special design for this application to achieve
specific electrical performancefor low level signals and to achieve low repeatable
bending torque over the deploymenttemperature range of 0° to 100°F.
The conductor complemento meet the magnetometerexcitation and signal feed-
back requirements is listed below. The cable bending torques were measuredto be
as high as 18 in-oz at low temperature and less than i in-oz at high temperature.
Silicone encapsulation is moldedaround the conductors to form the necessary flat
cross-section. In order to saveweight, the silicone encapsulation provides the
necessary insulation resistance for the single conductors and braided shields
without separate teflon jackets.
Quantity T e
Electrical Cable Conductors 3 Twist_ielded Pairs
2 Twis ted- Pair s
6 Shielded Singles
2 Sing le s
MDM STOWAGE SYSTEM
The MDM is folded for launch in the manner shown on Figure 2. The three
segments are supported by two cradles placed close to the mass concentrations of
the hinges. Both cradles have hinged segments that are released by redundant
pyrotechnic pin pullers. The outboard cradle (more distant from the root hinge)
contains a spring-loaded plunger to push the three segments clear of the cradle
despite any misalignment forces induced by distortion. The plunger also breaks
lose any static friction in the cradle linkage to insure a predictable deployment
sequence. The cradle linkage and plunger are shown on Figure 5.
The cradle structure is made of aluminum with MoS 2 dry lubrication on the
hinge pins. Two links are hinged and contain torsion springs for deployment to
provide redundant mechanisms for clearing the path for boom motion. Each boom
segment has fiberglass support spacers to contact the aluminum cradle structure.
These spacers are bonded to the boom segments so that the aluminum support brackets
provide a clean unobstructed track-like surface for the different shapes that
must pass during deployment.
26
Since both cradles must be pyrotechnically released for successful MDM
deployment,both contain redundant pin pullers to eliminate single point mechanical
failures. This redundancyconcept is shownon Figure 6. At each cradle both pin ......
pullers are retracted. The intermediate link permits release even if any one of
the 2 pins per cradle remainsengaged.
DEPLOYMENTSEQUENCE
Thedeploymentsequencewasmodeledon a nine-degree-of-freedom dynamic
simulation to include spacecraft dynamicsas well as boomdeployment. Thehinge
and cable friction torques were included with cases to cover a range of coefficient
of friction from zero to 0.5. Although the spin speedwill be selected for the
deployment, the setting tolerance is included in the simulation to combinelow
friction with high spin speedfor maximumresidual energy. On the other hand,
high friction is combinedwith low spin speedas a check on latch reliability.
Thedeploymentsequenceis displayed in Figure 7. Thedeploymenttime is
9.75 secondsfor a nominal case from electrical initiation of the squib to full
extension. Themiddle hinge latches first with the outer hinge latching only
about 0.i secondslater. At this point the root hinge actually reverses direction
momentarily and then resumesthe deployment.
Whenthe root hinge finally latches at 9.75 seconds, the outer hinge receives
its maximumbending moment. Table 2 displays an analytical comparisonof the
levels of energy, bending momentand rates encountered.
TABLE2
MECHANISMLATCHCHARACTERISTICSIN ORBIT
Order of Lock
Closing Rate (deg./sec.)
Locking Energy (ft-lb)
FundamentalFreq. (Hz)
Transverse Moment(ft-lb)
MomentCapacity (ft-lb)
High Energy
High rpm-Low Friction
Low Energy
Low rpm-High Friction
Root Middle Outer Root Middle Outer
Hinge Hinge Hinge Hinge Hinge Hinge
3 i 2 2 I 3
22 274 6.9 9.6 i04 72
0.53 4.78 0.48 0.23 0.73 0.ii
3.2 17.0 6.9 3.2 17.4 6.4
165 230 45 105 77 30
234 554 106 234 554 106
TEST VERIFICATION
The MDM has undergone a complete qualification at the unit level in-so-far-as
practical without the exact spacecraft interface. The qualification will not be
fully complete until the MDM can be integrated with the spacecraft. The two most
important aspects are vibration with true spacecraft structural coupling and thermal-
vacuum with realistic thermal inputs. However, the following tests are appropriate
for unit level design verification. The data will be discussed in regard to only
the first three items since the last two are purely survival exposures or a pass-
fail criterion.
i) Release and deployment in room ambient conditions
2) Static loads and deflection
3) Alignment repeatability
4) Thermal cycling with high and low temperature release
5) Workmanship vibration
27
DEPLOYMENTDEMONSTRATION
Roomambient release and deploymentwereconducted with a rate controlled spin
table to set the proper initial spin speed. Film coverage was used to visually
record the latch sequence, and potentiometer readings were recorded from each
hinge to verify latching velocity and exact timing. Strain gages were used to
record boom strain at each hinge during latch-up to verify actual bending
moments resulting from excess latch energy.
Initial testing was planned in room ambient conditions despite the aerodynamic
drag and increased friction due to i g forces. A simplified math model showed
that a slight increase in spin speed over the planned six rpm would provide
sufficient energy for a deployment demonstration. However, even at speeds up to
9 rpm in air, the root hinge would not deploy to the latch position. A subsequent
more detailed math model verified that most of the initial excess momentum was
dissipated at latch of the middle hinge and that aerodynamic drag prevented a
continuation of the deployment to the latch position of the root hinge. The
test was then repeated in a 40 foot diameter tent filled with helium, 1/7 the
density of air. The deployment was successful in helium, and the closure times
and rates closely compared with the math model as shown on Table 3. A graphic
plot of displacement versus time is shown on Figure 8.
TABLE 3
DEPLOYMENT TEST RESULTS IN 95% He ATMOSPHERE
9.0 RPM SPIN
TEST RESULT
7.5 RPM SPIN
SIMULATION TEST RESULT SIMULATION
LATCH TIMES (sec.)
Root 8.4 7.7 ii.I 10.3
Middle 2.4 2.5 3.2 3.3
Outer 2.9 3.0 3.6 4.3
VELOCITY AT LATCH (deg/sec.)
Root 28 32 13 18
Middle 236 235 140 150
Outer 83 83 50 40
PEAK BENDING MOMENT (ft-lb)
Root 201 191 132 122
Middle 237 227 131 127
Outer 52 52 16 34
ALIGNMENT REPEATABILITY
Most of the factors contributing to alignment error are small compared to
the 1.0 degree requirement and compared to the repeatability of the latch mechanism.
The effects of temperature gradient, for instance, are minimized by the low
coefficient of expansion of graphite-epoxy. Initial alignment uncertainty is
minimized by the use of optical surfaces on the spacecraft, the MDM mounting points
and on the magnetometers. These optical references are calibrated on the MDM
while suspended with the root hinge fixed to the ceiling. The one-g acting along
the axis of the suspended boom approximates the effect of centrifugal force in
orbit. This arrangement places an approximate elastic strain on the hinges as
in the spinning orbital condition. Three-axis alignment data are then used to
align instruments on the spacecraft at a later stage of integration.
28
Repeatability data were collected while the MDMwasvertically suspendedto
verify the alignment error for this predominantfactor. Ten readings were taken
in each direction of both lateral axes using an 18-ounceforce at the tip prior
to each reading. Thealignment in the boost direction was the morerepeatable
with an uncertainty of __+0. I inches at the tip. Thedeploymentaxis alignment
includes the hysteresis of the locking pin engagementwhich showeda tip
repeatability of _+ I.I inches. Although I.I inches at the tip seems excessive,
the angular error is only 0.35 degrees of the 1.0 degree error allowed.
STATIC LOADS
Two static load tests were conducted using sand bags for incremental loading
at the middle hinge, at the outer hinge and at the tip. Deflection measurements
were made at each loading point after each weight was added and after each weight
was removed. Strain gages were monitored to verify material stress st each _
loading condition up to the 9.0 g qualification load. The resulting data _ after
interpolation to the 7.5 g condition are shown on Table 4.
TABLE 4
STATIC LOADS TEST RESULTS
Middle Outer
A. Orbit Insertion 7.5 g's
Deflection (inches) 1.80 6.77
B. Deployment Loading
Equilva lent Hinge
Rotation (radians)
14.59
Roo t Mid d Ie Ou t e r
Hinge Hinge H ing e
0.026 0.052 0.054
Equilvalent Moment
(ft-lb) 638 270 67
Equilva lent Stiffness
(ft-lb/rad.) 24,560 5,240 1,240
In the 7.5 g limit load of orbit insertion the tip deflection was 14.59 inches,
38% greater than the design analysis. Although _ the deflection was greater than
planned, the boom met the strength margins according to the strain gage data and
the increased wobble of I.I degrees during orbit insertion is acceptable.
CONCLUSIONS
The Magnetometer Deployment Boom has been successfully developed to meet
the objectives of the Pioneer Venus mission. The development was conducted without
benefit of a full development model. Only one of the three hinges was machined
from aluminum to verify fits prior to commitment of the titanium flight hardware.
Aside from static deflection and deployment in air, the analytical simulations
were sufficiently successful to eliminate the need for extensive ground testing.
As a low cost approach the only MDM constructed was used for both qualification
and flight.
29
Figure I. Pioneer Venus
Orbiter
Figure 2. Magnetometer
Deployment Mechanism
with Therma i Control
Finishes
i •
Figure 3. Magnetometer
Deployment Mechanisms
without Thermal Control
Wraps
30
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Figure 4. Hinge and Latch
Assembly, Engineering
Model with Partial Boom
Segments
KICKOFF
PLUNGER
INBOARD
BOOM
O
ELECTRICAL
CABLE
MIDDLE
BOOM
0
ALUMINUM m
EPOXY FIBERGLASS LEASE m
BOOM SUPPORTS LINKS
\
SHELF
ATT AC H M E
FIGURE 5. OUTBOARD BOOM SUPPORT
PYROTECHNIC PIN
PULLER SQUIB
SH E LF
ATTACHMENT
31
RELEASE
F RONT V l EW
AND
TORSION
SP RING
DI RECTION I BOOM SEGMENT
/i
i
INTERMEDIATE i REE PIVOT
** l.., REDUNDANT /_ _ / PYROTECHNIC
_.s :___ / ,_PO,LE_S
FIGURE 6. REDUNDANT PIN PULLERS
Q
6.0 sec
4.0 sec
2.7 5 sec
O. 0 sec
SPACECRAFT
SPIN DIRECTION
-- STOWED POSITION
-- DEPLOYED POSITION 9.0 sec
sec
ROOT
HINGE
M I D D L E
HINGE
9.7 5 sec
OUTBOARD
HINGE
FIGURE 7. TYPICAL DEPLOYMENT SEQUENCE
32
200,
160
120
80
rc
(0
40
d
uJ
.J
Z
< 0
<
_ -40
Z
"r
-80
-120
-160
TEST DATA
[_ _- ROOT HINGE
MIDDLE HINGE
OUTER HINGE
-- _Ar/,. M I DD L E HINGE
I LATCHED ANGLE
,
-- _.,,,II,--OUTER HINGE /
I /"_1'_- R O OT HINGE
J/
#
&
-200 I I I I
0 2 4 6 8 10
TIME, SECONDS
FIGURE 8. SIMULATION OF 9.0 RPM DEPLOYMENT IN 95% HELIUM
33
.o;


Magnetometer deployment mechanism for Pioneer Venus

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