A unique mechanism developed for conducting friction and wear experiments in orbit is described. The device is capable of testing twelve material samples simultaneously. Parameters considered critical include: power, weight, volume, mounting, cleanliness, and thermal designs. The device performed flawlessly in orbit over an eighteen month period and demonstrated the usefulness of this design for future unmanned spacecraft or shuttle applications

Devine, E. J.

Evans, H. E.

English

NASA Technical Reports Server

N76-19178
By E. J. Devine and H. E. Evans
Goddard Space Flight Center
SUMMARY
A unique mechanism has been developed for conducting friction
and wear experiments in orbit. The device is capable of testing
twelve material samples simultaneously. Power, weight, volume,
mounting, cleanliness, and thermal designs were particularly
critical requirements that were successfully met. The device
performed flawlessly in orbit over an eighteen month period and
demonstrated the usefulness of this design for future unmanned
spacecraft or Shuttle applications.
INTRODUCTION
It is well recognized that the increased tendency of cold weld-
ing (friction) and wear in vacuum are potentially serious hazards
to mechanical devices operating in space. When exposed to the
space environment, conventional lubricants rapidly evaporate,
oxide films disappear once disrupted, and chemically clean metal
surfaces can come into i_timate contact. The result is dras-
tically increased friction and wear due to adhesion (cold weld-
ing). (Reference I) This cold-welding phenomenon in space is of
great interest to unmanned spacecraft instrument &_d equipment
deslg_ers and especially to the manned space effort because such
procedures as repeated orbital docking and assembly may be af-
fected by cold welding.
A very widespread and costly effort was undertaken to investi-
gate this problem under laboratory simulation of the space en-
vironment. At one time, a count showed that approximately 100
groups (both government and industry) had been or were studying
cold welding and wear. Most of these studies involve vacuum ef-
fects.
In contrast to the extensive laboratory investigations and In
view of the zo.cognlzed importance of the friction and wear
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problem, it was surprising that very little data had been, or
were planned to be, obtained in the actual space environment.
The only previously planned experiments were carried out by the
Jet Propulsion Laboratory (Reference 2) _z the early Ranger
flights and the Air Force Rocket Propulsion Laboratory using the
Experimental Research Satellite (Reference 3).
OBJECTIVE
The major objective of this program was to determine the effect
of the actual space environment on friction (cold welding) and
wear of widely used spacecraft materials. In addition to veri-
fying or disproving laboratory results, this experiment was
planned to establish a better definition of the degree of vac-
uumrequired for an adequate simulation of the space environment.
APPROACH
It was decided that the disk and rider technique would be used .-
for measuring cold-welding tendency and wear. Welding tendency
was determined by measuring the strain induced in a cantilever
beam supporting the rider. Semiconductor strain gages mounted
on the beam were selected for this measurement. Seven disks and
twelve riders were picked as a realistic sample size.
Cleanliness is mandatory when conducting experiments involving
material surface effects. (Reference 4) Great care was required
in designing the experiment to minimize the possibility of con-
taminating the material couples. Design of the drive train,
instrumentation, and material selection had to be made with the
objective of minimizing outgasslng. The drive train had to be
hermetlcally sealed and materlalv with vapor pressures less than
10 "13 tort had to be used.
To minimize contamination by outgassing from the satellite and at
the same time to give maximum exposure of the experiment to the
space environment, it was necessary to mount the experiment on
the outer surface of the spacecraft and to baffle it from any po-
tential outgassing source.
Most o_ the laboratory test work was done at pressures of 10 -8
to 10 -7 tort; therefore, it was zequired that the orbit chosen
for this experiment have an apogee that would expose the
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1976012084-074
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spacecraft to a pressure less than i0-I0 tort, preferably 10-12
torr. This condition would shed light on the degree of vacuum
required for adequate simulation.
The materials selected for study are commonly used in spacecraft
and also have properties that enable the effects of mutual solu-
bility and hardness to be studied. Materials were:
o Gold and silver
o 7075 anodized aluminum and 440C stainless steel (Re60)
o 440C stainless steel (Re60) and 440C stainless steel
(Re60)
o 440C stainless steel (Re60) and nitrided Nitralloy 135
mod steel
o 440C stainless steel (Re60) and 1020 carbon steel
o Be-Cu alloy 25 no. 190 heat treat and 440C stainless
steel (Re60)
Two samples of each of these combinations were tested with the
experiment module.
INSTRUMENTATION
The basic friction-test mechanism consisted of a hemispherical
rider sliding on the flat face of a disk rotated at constant ve-
locity. This geometry is simple and has been widely used in past
investigations of vacuum effects on friction phenomena.
Parameters measured and telemetered were friction force, normal
force, and displacement of the rider because of wear. These were
measured _ , _rain-gage transducers.
_Cl_e _lde_- ._ _rt ° _ was designed with two flexible sections:
or,,J¢_si_ive _, " _ion force and the other sensitive to normal
_rce. "_ _ ,_ _ons were instrumented with epoxy-bonded_
d_L_ _ i_c_n strain gages. The gages were arranged in a half-
h_d_,_ configuration to increase the output signal and to cancel
t;Leapparent _traln due to thermal expansion mismatch of the
gage , td substrate.
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1976012084-075
Normal load was applied by means of an independent, adjustable
spring as shown in Figure I. Normal load for the gold versus
: silver samples was 4.45 N (I ib); all other me_al couples were
loaded to 8.90 N (2 Ib). The occurrence of a given amount of
wear (0.12 cm) of the rider would result in a corresponding mo-
tion of the load rod that would bring the rod in contact with the
wear transducer, giving a single-point measurement of wear.
Seven disks were stacked on a common drive shaft and 12 rider as-
semblies were equally spaced around the periphery, giving a capa-
bility for measuring 12 material combinations. Surface velocity
between disk and rider was approximately 0.i0 cm/s.
In the event that the friceion force on a particular couple
reached a level such that excessive power was required, the rider
: was removed from contact by means cf a sealed pyrotechnic actua-
tor. The electrical signal f6£ firing the actuator was generated
when the friction gage output exceeded a level corresponding to
a coefficient of 3.3.
The drive train for the device incorporated two state-of-the-art
devices. A brushless motor with solid-state commutation required
only 3.5W to drive the fully loaded device. The motor drove
an intermediate gearhead of 16.3:1. The motor and gearhead were
enclosed in a hermetically sealed package to avoid contamination
of tbe friction experiment and at the same time to maintain con-
ventional lubricants on the high-speed drive elements. Trans-
mittal of power through the hermetic package, together with an
: additional 72:1 gear reduction, was accomplished by a harmonic
drive mechanism.
Considerable attention was given to lubrication of the few slow-
moving parts exposed to the space v_cuum. The spline (gear) of
the harmonic drive mechanism was gold-plated and a light burnish
of aoJybdenum disulfide was applied, run in, and the excess re-
moved. Linear ba!l bushings, which floaL the transducer load
rods, were in-situ coated wit_ molybdenum disulfide. The drive
shaft support bearings were e_uipped with a self-lubricating
duroid retainer and shields.
The strain-gage trm_sducers were designed to meet the following
requirements:
Load range Normal force: 0 to 13.3 N (3 Ib)
FrlcLion force: 0 to 26 N (6 ib)
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1976012084-076
Maximum strain 5000 X 10 -6 cm/cm (limited by the bond
: strength)
Natural
frequency 2000 Hz
Sensitivity Normal force: 13.5 mV/N
Friction force: 6.8 mV/N
These requirements dictated the use of semiconductor strain-gages
for the required sensitivity. Literatl.re search (Reference 5)
indicated that radiation exposure in a one-year orbit would not
significantly affect the gage output. Silicon $ages of p-type
material are more radiation-resistant and were accordingly se-
lected. The gages had a nominal gage factor of 700.
Expezimental studies were conducted in a search for a bonding
technique thaz would eliminate organic materials. Although two
of these approaches (ceramic bonding and soft soldering) showed
promise, neither could be flight-quolified in the time avaZlable.
As a result, a high-temperature epoxy bond was selected for _his
application.
The transducer substrate chosen was 17-4 PH stainless steel. The
gages were used in a half-bridge configuration with one gage in
tension and the other in compression. The gage output as a func-
tion of load (strain) deviated only slightly from linearity with-
in the desired range of operation; therefore compensation for
this effect was not required.
The gages were compensated for temperature effects. First, the
two gages in a half-bridge were tested for gage-factor change
with temperature. Any mismatch was eliminate_ experimzntally by
a shunt resistor across the gage showing the larger change. Even
with the gages matched for gage-factor change with temperature,
the bridge sensitivity dropped with increasing temperature. This
effect was cqmpensated by a resistor i_, series with the voltage
supply to the bridge. At elevated temperatures, the bridge re-
sistance increased, causing the current and the voltage drop
acro_s the series resistor to decrease. As a result, the volt-
age across the bridge and the bridge output rose to compen-
sat_ for the loss of sensitivity. The compensated transducers
readily met the Jpecification of less than • 3 percent of full
output drift over a temperature range of -20 ° to 60°C. Typical
transducer 1Lnearity i_ shown by Figure 2.
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1976012084-077
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The relatively high output of the semiconductor strain-gage trans-
ducers made -Ceasible the use of individual linear integraeed-
circuit amplifiers. Thus the complexity of commutation was
avoiaed and redundancy was attained. Some penalty in power drain
was inherent in the individual amplifier _.pproach. It was desired
that the amplifier stability and temperature drift performar, ce
w_ld at least equal that of the transducers. This was, in fact,
achieved by careful burn-in and selection. A sample of 500 com-
mercial pA709 amplifiers was burned in over a temperature range
of -40° to 100°C; 250 were selected because of mi_.imum change in
offset cLrrent.
The schematic for the strain-gage half-bridge, the s.-4 .-shunt
resistor-type temperature compensation, and the integrated-cir-
cuit a_plitier is given by Figure 3.
Ancillary circuits are required for the following:
o Convert_-r regulators for supplying -3 V (4-0.5 percent)
for the strain-bridge excitatio,L, ± 9 _ for unregulated
amplifier pc_er, and 5 (_ 0.5 percent) V for temperature
sensors
o '£hermistor cemuerature sen_or_ "-r . ._flight temperature
mortitor._ng
o Shaft ,:otati_ sensor for status monitoring of the drive
trai_
o limer for turning the experiment on and off at a I0 per-
cent duty cycle to ex_end the lifetime of the material
couples
The exper_rent gs _ired with solid-conductor, teflon-insulated
wire to n.inhrize trepped gass_s and to provide a control of the
outgassed materials (Figure 4). Teflon and the minute a_ount of
epoxy bondlnZ meterial for the gage_ are the onli organic mate-
rlals permitte _ in the friction-measurement envir-.ms.
TEST PROGR_4
Th_ £nstzoment was qualified for launch vibration a_ tl_efollo_-
trig _evels :
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1976012084-078
0"
E
Prototype Flight Qualification
SLle vibration 8g vector 5g vector
Random vibration 6g 6g
The prototype disclosed large displacements of the normal load
rods when the vibration was parallel to the rod (thrust axis).
As a result, vibration stops were incorporated to limit the ex-
cursion of the rod and thus prevent overstressing of the normal
load strai_l-gages. Two units were subsequently qualified at
flight levels with no significant shifts in the straln-gage
transducers.
_e experiment was qualified for thermal vacuuw over a tempera-
ture range from 0° to 60% and vacuum from I0-_ to i0-7 tort.
These tests were conducted in an oil-free, ionizatlon-pumped
system and are the source of tilevacuum-friction data discussed
in the next section.
The instrument was integrated and tested on the OV-I-13 Sstellite
with the following results:
o Electrical, RF, and magnetic interfe:ence tests showed
no _nterference with the satellite systems or the other
sensitive experiments.
o Ther_l-vacuum (including solar-simulation) tests estab-
lished the worst-case condition as the cold temperature
and demonstrated the capability of the passive ue_@era-
ture control to maintain a l_edicted range of 0 ° _o 15°C.
o Satellite vibration test with a Ig vector input resulted
in no degradation. The vibration test demonstrated that
the strain-gage transd,,eersmaintain remarkable stabil-
ity under mechanical abuse. The other unknown factor
was the stability with time. Table i gives stability
data for a complete set of transducers -_asured at the
beginning and end of an 8-month interval. The drift
figures shown include both the transducer and integrated-
circuit amplificr circuit. In only one case did the
drift exceed the design specification, the drift in this
instance being in the amplifier circult.
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v
1976012084-079
0
(,
IN-ORBIT OPERATION
The friction-and-wear device _as operated in orbit throughout
the 18-month life of the OV-I-!3 Satellite. The experiment was
turned on and data recorded on an average of one orbit per week.
Because the duty cycle was designed for 3 minutes of operation
out of 30, this schpdule resulted in an average of 6.5 turn-on
intervals per week. This totals approximately 500 intervals of
3-minute duration for a total running time of about 25 hours.
Instrument performance was flawless throughout the mission.
REFERENCES
I. F. P. Bowden and D. Tabor, The Friction and Lubrication of
Solids, Oxford, 1950.
2, J. B. Rittenhouse, L. D. Jaffe, R. G. Nagler, and H. E.
Martens, "Friction Measurements on a Low Earth Satellite,"
Report 32-402, Jet Propulsion Laboratory.
3. '_ Study of Space Propulsion System Experiments Utilizing
Environmental Research Satellites," AFRPF-TR-65-73, Air
Force Rocket Pro_:Ision Laboratory.
4. D. R. Milner and G.W. Ro_e, "Fundamentals of Solid Phase
Welding," Met. Fev. 7 (28), 1962.
5. J. F. Kircher anJ R. E. Bc_uan, Effects of Rad_ation on
Materials and Com_,onents, Reinhold Pub. Corp. i964.
16
1976012084-080
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Figure1 Sclmmmti++diq_wn offriction-testm chanismshowingIX_i_onofspringforapplicationof
normall,xd.
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lrllIII _ _. UI
Figure2 Transduceroutputasa functionof load. Transducer_ul_trate: 17.4 PH_I
;reel; 3.00-V excitation; SN - normal load; SF = friction load. _
w
; ORIGINALPAGE IB 77 :
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1976012084-081
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Figure 3. Schematic diagrarr of the half-bridge, the series-shunt resistor type i
ten,perature compensation, and the Lntegrated-cireuit amplifier of a
strain-gage transducer.
__ e .............
Figure4. Fric1_,on-and-wearmodule.
_ 78
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1976012084-082
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1976012084-083


In-flight friction and wear mechanism

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