A simultaneous spin/eject mechanism was developed for aerospace applications requiring a compact, passive device which would accommodate payload support and controlled-release functions, and which would provide a highly accurate spin-ejection motion to the payload. The mechanism satisfied the requirements and is adaptable to other deployment applications

Banks, T. N.

Palmer, G. D.

English

NASA Technical Reports Server

N76-28288
A SIMULTANEOUS SPIN/EJECT MECHANISM
FOR AEROSPACE PAYLOADS ¢
By G. Dan Palmer, TRW Systems Group
and Tom N. Banks, AVCO Systems Division
ABS TRACT ; .
i
A simultaneous spin/eject mechanism was developed for aerospace
• applications requiring a compact, passive device which would accommodate
payload support and controlled-release functions, and which would provide a
highly accurate spin/ejection motion to the payload. The mechanism saris- i
fled tile requirements and is adaptable to other deployment applications.
IN TRODUC TION
The simultaneous spin/eject mechanism is an important element of a i
recently developed system for accurate, efficient, low-cost deployment of
aerospace payloads from their final booster stage (Figure t). Al_houghthis iis not the first program to employ a mechanical device to impart a simulta-
neous spin and eject function to a payload, the mechanism is unique because
it is extremely compact and lightweight, has a highly efficient energy
release, and has excellent, repeatable performance. These features have
been confirmed by extensive analytical simulations and ground tests under
t-g and zero-g environments.
For this application, ease of integration with the vehicle, the support-
ing spacer, and the booster configuration was important; in addition, payload
bumping potential and desired system accuracy called for l_w tipoff effects.
The mechanism can be easily adapted to other aerospace deployment applica-
tions. Basic design parameters are listed in Table 1.
The successful development of this mechanism involved _.he joint efforts 4"
of three separate contractors: AVCO, TRW, and General Electric. AVCO
was responsible for design, development, fabrication, and testing of the
mechanism as a subsystem; TRW supplied design analysis, system engineer-
ing, and technical support; and GE was responsible for the integrated system
(final stage booster with epin/eject mechanisms and vehicles).
fi
?This work was supported by U. S. Space and Missile Systems Organization
under Air Force Contract Numbers F04701-74-C-03_5, -0326, and -0328,
2
J
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3PL Technical Memorandum 33-777 165
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https://ntrs.nasa.gov/search.jsp?R=19760021200 2020-03-20T02:33:49+00:00Z
SPIN/EJEC T MECHANISM
The post-deployment configuration of tile mechanism, Figure 2, has
three basic -lements: (1) a spin/eject device which imparts the desired rela-
tive separation velocity and spin rate to the vehicle; (2) a central tie-down
system consisting of a structural spacer, a pyrotechnically actuated separa-
tion nut, and a high strength tie-down bolt; and (3)two in-flight disconnects
(IFD's) that provide the electrical interface between the vehicle and the
booster.
Spin/Eject Device
The spin/eject device consists of a spring-actuated guide housing with
helical grooves and helix guidepins, with the guidepins also functioning as
mechanical stops for the spring guide. The housing, which is fixed to _he
spacer hub, accommodates the two fixed and diametrically opposite guidepins
that ride in matching helix grooves. The force of the spring pushes the
spring guide against the guidepins causing simultaneous spin and longitudinal
ejection. The selected helix angle determines the amount of rotation in rela-
tion to the axial translation of the spring guide and the vehicle. Ejection and
spin forces are applied to the vehicle through its rear cover upon which the
spring guide bears while simultaneously engaging two torque pins.
The elongated slots at the ends of the helix grooves provide low rebound
and self-locking (Figure 2). During the development ejection tests, a change
from a completely round pin to a semi-fiat Din was made to eliminate local "
deformations (due to high bearing stresses) which occurred along the helix
groove surfaces. These local deformations had contributed to high friction
and posed a potential threat to test result repeatability. Friction was further
reduced by an application of Molycoat to the guidepin and grooved surfaces.
The design change and the coating permitted a reduction of friction forces by
a factor of 2, resulting in a realization of 88 percent of the theoretically
available energy. The redesigned pin also permitted re-use of ejector hard-
ware without refurbishment for successive ejection tests.
Spin torque is transmitted from the moving spring guide to the vehicle
through two torque pins extending aft from the vehicle rear cover to slotted
grooves in the spring guide forward face (Figure 2). This interface is (:riti-
cal at the instant of vehicle-to-spring guide disengagement because it not only
provides the torque transfer, but also supports and stabilizes the vehicle dur-
ing the ejection stroke. The latter condition is facilitated by both the flat
area and the conical frustum extension on the forward face of the spring guide
which matches and centers the mechanism with respect to the flat area and
mackined indentation in the vehicle rear cover. All ejection tests demon-
strated a clean, non-disturbing separation at this interface.
1
166 JPL Technical Memorandum 33-777
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1976021184-173
Central Tie-Down System
A central tie-down system was developed that utilizes a mechanical
spring deployaaaent device to impart a simultaneous spin/eject motion to the
vehicle upon activation of an explosive put. The structural spacer provides
the interface transition between booster and vehicle while accommodating
powered flight loads. The tie-down and release device consists of a high-
strength separation bolt, under a preload of 37.8 kN (8500 lb), with an ord-
nance activated separation nut (SOS114196) located within the vehicle.
The preload magnitude is based en static and dynamic loads and vibra-
tion environments assoc;.ated with the boost phase, and is designed to pre-
clude vehicle-to-spacer "gapping." A simple assembly and pa-eload
procedure was developed for integrating the spin/eject mechanism with the
spacer and vehicle.
The tapered seat at the aft _.;nd of the spring guide housing prevents
rebound of the ejected separatio'_ bolt. The bolt is trapped in its rearward
position prior to spring guide movement providing a protuberance-free
interface with the vehicle.
Inflight Disconnects
Electrical requirements for the subsystem were met by providing two
IFD's (18-pin standard Bendix connector) at the rear cover/spacer interface
(Figure 2). These connections provide checkout, monitoring, and control ""
functions from the launch facility, as well as transmission of the activation
current on in/light command to fire the pyrotechnics of the separation nut.
To preclude pin binding during extraction, AVCO designed and devel -
oped a connector support device. This device allows three-dimensional
adjustment during assembly and adequate freedom for rotation which permits
the pin connector to be separated with little or no side load upon preload
release and during the initial spin/eject motion. This "floating nut-plate"
(Figure 3) prohibits relativ- motion prior to the separation event and accom-
modates the rotation required up to the time of LFD physical separation.
SPIN/EJEC T REAC TIONS
Booster control system capability limits dictate the maximum axial and
angular (roll) impulses to the booster at 75, 6 Ns (17 Ib-sec) and 13.6 Nms
(10 ft-lb-sec). Booster component responses to the separation nut ordnance
shock and to off-center deploya'nents were of concern; however, ground test
results verified that the induced environments were within acceptable limits.
Although the deploying vehicle is supported at the forward face of the
spring guide, very little transverse moment capability exists at this inter-
face. Consequently, the vehicle tends to hinge, or tip, relative to the mov-
ing spring guide ifthe reaction moments on the booster stage yield
!
JPl, Technical Memorandum 33-777 167[
1976021184-174
overturning moments to the vehicle greater than tile restoring moment
capability. Off-axis deple_Tnents, i.e., where the ejection force line of
action does not coincide with the combined booster-vehicle e.g., are sus-
ceptible to this tendency. The degree of hinging is a function of the deploy-
ment reaction moments and the booster-vehicle mass properties.
Because the existing design has limited moment capability at the
vehicle-spring guide interface, and because the booster is relatively massive
for the initial vehicle deployments, only the last few vehicles in multiple
deployments experience significant hinging.
DESIGN EVALUATION
The mechanism's flight performance was verified through extensive
analytic simulations and ground tests. Functional performance specifications
required a separation velocity of 1 + 0. t m/see (3.24 4- 0.3 ft/sec) and a vehi-
cle spin rate of 450 "72 deg/sec with a tipoff rate less than 3 deg/sec, while
not exceeding the booster reaction impulse limits. Further, contact with
adjacent equipment or vehicles during boost and deployment was to be avoided.
Analytical Simulation
Both closed-form and computerized solutions were utilized in the func-
tional performance analyses. These simulations involved single-body ejec-
tions (the ejector spacer cantilevered from an infinite mass) and two-body
ejections (the ejector spacer rigidly attached to the booster). Closed-form
terms satisfactorily describe the kinematics of the mechanism because the
helix proportions the ejection energy between spin and axial translation. The
closed-form solution was used in the mechanism parameter/performance
tradeoffs and for support of the early development tests.
To include booster control system and ejector flexibility effects (e. g.,
torque pin/rear cover and guide pin/groove compliances), however, a more
extensive computerized simulation was developed. The simulation was used
in establishing the pre-test predictions and post-test evaluations for both the
system separation tests (zero-g drop) and actual flight tests. Critical areas
of deployment clearance, optimum deployment order and sequence for multi-
ple deployments, and trajectory inputs (targeting biases) were evaluated
based on the computerized simulation. This simulation included two six-
degree-of-freedom bodies (one with a complete control system and logic),
clearance envelope geometry, flexible ejector effects, and provisinns for an
approximation to the hinging effect.
Ground Test
Development and verification of the simultaneous spin/eject me, a-
nism, both as a subsystem and as part of a missile system, comprised a
¢
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lb8 JPL Technical Memorandum 33-777
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1976021184-175
major portion of the ove,'all program testing activity. The ground ejection
testing history is briefly summarized in ;-'igure 4.
Ejection Tests (l-g)
Early design verification of the mechanism covcept was obtained from
a series of fLxed-base vehicle downward ejection tests. Several tests were;
run using a conservative simulation of the hinging environment; a 10-degree
vertical tilt induced a transverse moment on the vehicle; during the ejection
stroke caused by the 1-g gravity vector. Several test configurations under-
went preconditioning (shock and vibration} prior to testing. The vehicle was
a mass simulated model with actual rear cover hardware; the spacer and the
mechanism were full-scale development, or flight-type, hardware
configurctions.
AVCO developed and employed a computerized data reduction technique
for use with the high-speed (400 frames/see) film coverage. Markings were
made on the ,ehicle model, and the relative positions of these markings and
changes in dir_ensions were measured, frame i_y frame, using a Vanguard
film anal_.zer. Appropriate geometric relationships and camera optics coded
into a computer program were input with measured data from the Vanguard
analvzJr; velocity, spin rate, and tipoffdata were output.
For this test phase, eleven subsystem ejection tests were performed:
three checkout tests of early design hardware; five development tests, one
with the semi-flat pin reference design; and three engineering demonstration
tests, all with design reference hardware (Figure 4).
In addition to verifying the simultaneous spin/eject concept and design
approach, the tests were used to: (1) demonstrate functional operational and
performance capabilities, (2) determine repeatability, and (3) establish final
design performance parameters and compliance with design requirements.
Separation Tests {Zero-g)
The system-level separation tests utilized the mechanism tt, .:,'t_; • t
"_, vehicle from a booster model under a zero-g environment. The z .,"
environment was obtained by simultaneously dropping the vehicle ./r.,., '_
booster model prior to initiating the ejection event, and continuinf. _., .tall
beyond end of stroke (Figure 4). As part of the flight-proof testing, an addi-
tional test conducted with the vehicle and spin/separation flight-proof
was
eject subsystem in the zero-g environment.
The configurations tested represented different mash properties of the
booster, as well as different offsets for the ejection impulse reaction with
respect to the system e.g. Corresponding results were predicted for each
I test from analytic simulatio.-_s that accounted for specific mass and offsetparameters. T}'._ col-rc._pondeace was excellent between predictions and
rlleasureme,its.
JPL Technical Memorandum 33-777 Ib9
d
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1976021184-176
Deployment data for the separation tests from the vehicle and booster
model rate gyros and accelerometers were hard-wired to the data acqui,_ition
system. For the flight-proof test, the dynamics portion of the flight instru-
mentation system (rate gyros, accelerometers, nd PGM telemetry) was
utilized.
Results of the four demonstration separation tests and the flight-proof
separation test data, including the effects of new hardware _nd environmental
exposure, indicate oustanding agreement with predictions and an excellent
degree of repeatability (Table 2).
GONG LUSION
The final design hardware and functional performance of the simulta-
neous spin/eject mechanism fully meets the subsystem and system require-
ments as verified by analyses, ground tests, and successful flight tests.
Specifically,
• The electrical inflight disconnect design, which accommodates
simultaneous rotation and translatiorJ, is functionally adequate;
no adverse effects were observed.
• The ordnance-activated separation nut is a reliable device for
initiating the spin/eject process. The shock environments
produced by separation nut activation and spin/eject process
are within acceptable limits. -'
• The m ,:hanism provides predictable, acceptable, and repeata-
ble performance; tipoff effects were consistently less than the
3 deg/sec requirement.
• Deployment reactions generated by the operation of the mechan-
ism were accommodated by the booster without detrimental
effects,
The design latitude of this mechanism can readily accommodate a prac-
tical range of vehicle mass properties by modification to the helix an_,le,
ejector spring rate, and/or stroke. The design ensure," adaptability to future
aerospace vehicle or spacecraft deployment subsystem ,spplications.
AC KNOW LEDGEMENT
The author• gratefully acknowledge the contributions and assistance of
R. A. Winje of TRW Systems Group and S. J. Terani. Jr. of AVCO Systems
Division.
i 170 $PL Technical Memorandum 33-777
1976021184-177
Table 1. Mechanism basic design data
Parameter Norrtlnal Magnitude
"Jechanism
Ejection spring rate 26. 62 kN/m (1824 lb/ft)
Spring free length 16.6 cm (6. 55 in. )
Spring compressed length 9. 1 cm {3. 5g in. ) -.
Active stroke 5. I cm (1. 9_ in. ) , -
Initial spring force 2, 13 kN (480 lb)
I
Torque radius 2.4 cm (0. 9488 in. )
• Helix angle 9. 965 deg
Displacement ratio, angular to linear 4. 18 deg/crn (10.608 deg/in. )
_'riction coefficient 0. 12
Exterior size 6.6 cm {2.6 in. )diameter. 14.0 crn
{5. 5 in. } height
Weight (including spri-,,) 1.45 I,g (]. t9 lb)
Spacer
Spacer size 25.4 cm (10. 0 in. )diana ter, 20. 3 cm
{8. 0 it,. ) height
Spacer weight {including separation bolt, 2.97 kg (6.54 Ib)
nuts, and cable assemblies)
Table 2. Zero-g separation results
Parameters Test Designation
1 2 3 4 Flight- Proof
Axial Velocity, m/see
measured 0.98 I. 07 1.04 0.94 0.8_
predicted 0.95 I. 01 I. 01 0. 88 0. 82
Axial Velocity, ft/sec
measured 3.2 _, 5 3.4 3. 1 2. 8
predicted 3. I 3. a _. 3 2.9 2 7
Angular Rates, deglsec
measured pitch -5. 0 -0. 1 0. _ -3.0 -_.
predicted pitch -3.6 0. I 1.8 -5.2 -5. 2
r, teasured yaw 3.2 0. 2 - I. 1 3.2 2. 8
predicted yaw _. I O. 2 -0. S 3.0 2.4
measured roll 478 483 472 475 472
predicted r_ll 465 453 456 469 458
Tipoff Rates7 deg/see 1.4 0.2 i. 4 2. Z 1.9
fMechanism induced, based on vector difference of test and predicted transve,_e rates
"" .IPL Technical Memorandum 33-777 171
l
1976021184-178
X 6 DOF BOOSTERWITH
/'-'- CONTROLSYSTEM
Y _J" BOOSTERMASSTO +
SPIN-EJECT | ) "_ _ | VEHICLE MASSRATIO
FORCESfrORQUES - ,_ 10:!
" 6 DOF VEHICLE \\ TIPOFF IMPULSE
AT ENDOF STROKE
(RANDUM)
Figure I. Schematic representation of vehicle deployment
;+
,0 VEHICLE _) VEHICLE END :i
SEfARATION SPRIN( I GROOVE +
NUT CABLE U _
"_ IFD _+
CONNECTOR ,_
+ 12) ++ +_
p EJECTOR ++
i TORQUE ; GUIDE
_ PIN (2) HOUSING
RETAINING FD _'JlOEFIN (21 +RECEPTACLE
RETAINER 12)
INTERFACE KEY i
_ARINO ii_
SEPARATIONBOLT
ADJUSTINGNUT FIN LkFETY RING
• BOOSTER • BOOSTER END
Figure Z. Simultaneous spin/eject mechanism, post-deployed
I?Z ,_ IPL Technical Memorandum 33-777
1976021184-179
O ASSEMBLED VEHICLE AND SPACER
GROUNDING FINGERS\ _ / rCONNECTOR[ _1
D,.EO',_"OND.E.,EC"lO,',
AND PIN SEALS--_---_
_ HOUSING
: VEHICLE SIDE _ _ SPACEFU'BOOSTER SIDE
<
O DISENGAGED VEHICLE AND SPACER ,_?"
.o
° _
LAR DIPLACEMENT _ _
NEEOED FOR 0.762 CM
_,.,.w0iDISENGAGEMENT(0,300 IN) OF CONNECTORSVEHICLE SIDE _ SPACERIBOOSTER SIDE
Flgu_'e 3. Basic design, Infllght disconnect
3PL Technical Memorandum 33-777 173 I
t
i i
19760211R4-1RF_
• EJECTOR TEST SETUP, t-g
___ OVERHEADSUPPO.T• TESTW_"
_! R_ed base, downwardejection ;n
SPACER/EJECTOR -- "_.<f_- ]'g field
J_=-_'_ -_- VEHICLE MODEL • NUMBER OF T_-SfS AND MECHANISM
, (DESCENDING) HARDWARE CONFIGURATION: 8
-ri.ONGEES--- i _ 3C_,ecko_,(,orlydesign.ardware)
:) Development (1 with final heJ'ix
" = / _ (_ SIDE CAMERA pin design)
._, ( AV, SPIN)ARRESTING - _ 3 Engineering demonstration (design
COLLAR .._/_,. reference hardwo,e)
°
i[ P' , BOTTOM CAMERA
-" (TtPOFF. AV, SPIN) RESULTS
• lo Successful deployments; res,Jlts
and repeatability allowed reduction
of planned test program
• Hinging effects predictaL, le and
.° acceptable
1 • SYSTEM SEPARATION TEST SETUP, 0-g • Friction determination: ,.=
p = O.124(7 tests)
• Tipoff rate:BOOSTER MODEL
_//(DESCENDING)
< 3 deg/sec (10 tests)
,_'.....- BOOSTER ARRESTING • Spin-rate and velocity predictions
__ _ BLOCKS confirmed: (50-g tests)
f .
SPACER/EJECTOR _ - VEHICLE MODEL
(DESCENDING)
_V • TESTTYPE:
. .i_ ,_ _ VEHICLE Twobody separation 0 free-fall in
_ _ ARRESTING 0-g field
_i _;;_ OEV,CE" • NUMBEROF TESTSAND MECHANISM :
• HARDWARECON;,IGURATION: :
_ 4 Engineeringdemonstration(designreference hardware)
• -_ I Flight proof (flight configuration)
i i .Figure 4. Ejector and system separation tests ;
174 JPL Technical h4omorandum 33-777[
1976021184-181


A simultaneous spin/eject mechanism for aerospace payloads

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