The Apollo 13 anomaly provided considerable impetus for a variety of types of cryogenic and ignition tests. The logic of the various test program designs, the test techniques, and their final impact upon the investigation findings are described. In addition, several test programs initiated to determine the thermal performance and general performance characteristics of the redesigned Apollo 14 cryogenic storage system are presented

Mcgee, J. M.

Propp, C. E.

English

NASA Technical Reports Server

A 7 -63 7qe
"A REVIEW OF CRYOGENIC TESTING PERFORMED BY
THE THERMOCHEMICAL TEST BRANCH, MANNED SPACECRAFT CENTER
IN SUPPORT OF APOLLO 13 AND 14"
By Clarence E. Propp and John M. McGee
NASA Manned Spacecraft Center
ABSTRACT
The Apollo 13 anomaly provided considerable impetus for a variety of
types of cryogenic and ignition tests. This presentation will treat
the logic of the various test program designs, the test techniques,
and their final impact upon the investigation findings. In addition,
the presentation will cover several test programs initiated to deter-
mine the thermal performance and general performance characteristics
of the redesigned Apollo 14 cryogenic storage system.
INTRODUCTION
This paper is a brief summary of the Apollo 13 and Apollo 14 oxygen
cryogenic storage system (CSS) tests performed at the Thermochemical
Test Area, Manned Spacecraft Center. This area includes five unique
facilities which provided the many specialized test environments and
fixtures necessary to accomplish the subject test programs. There were
25,000 manhours invested in 30 test programs of which 50% were com-
pleted between April 14 and May 22, 1970 in direct support of the
Apollo 13 anomaly and the requirements for redesign of the Apollo 14
CSS were made; however, this paper will treat only the test planning
logic and the highlights of the more informative tests. The Apollo
13 and Apollo 14 test programs were concerned with ignition and com-
bustion phenomena and the Apollo 14 tests were additionally concerned
with determining the operational characteristics of the redesigned CSS.
APOLLO 13 TESTS
The Apollo 13 Anomaly Investigation was a very intense series of test
programs and group discussions. The testing was controlled from a
single point for the purpose of avoiding redundant and unproductive
testing. The control of testing was effective and the pattern which
433
https://ntrs.nasa.gov/search.jsp?R=19720016149 2020-03-23T10:44:29+00:00Z
evolved was as described below:
Baseline Tests
These tests were performed for the purpose of basic orientation of
the Apollo 13 Review Board, the Observers, and many of the Panel
members on the combustion characteristics of the various CSS materials
and the operational characteristics of the CSS components and instru-
mentation.
The materials combustion tests were conducted in high pressure oxygen
gas, supercritical oxygen, and subcritical oxygen. There were several
informative movies taken during these tests which helped considerably
to approximate combustion rates and ignition energy values for teflon
wire insulation in 900 psi supercritical oxygen, as well as to estab-
lish a graphic understanding of this combustion process.
There were also several test programs conducted for the purpose of
characterizing the CSS components and instrumentation in order to
more beneficially interpret the data recorded pertinent to the Apollo
13 accident. The TTA test assignments were designed to evaluate the
Apollo CSS fluid flow dynamics and its associated valves, pressure
transducers, and flowmeter response times. The results of these test
programs confirmed that the analysis methods being used on the Apollo
13 data were correct.
At this point, it was evident that combustion of the teflon insu-
lation was a realisitc consideration and the energy released could
easily be sufficient to provide the pressure rise rates recorded on
the Apollo 13 data. Therefore, it was determined that while ignition
and combustion studies continued, an effort to determine the ignition
source would be initiated. This included the examination of all
aspects of the CSS handling prior to flight.
Simulations of CSS Preflight Servicing
The TTA tests were simulations of the Apollo 13 #2 oxygen tank special
detanking procedures which were used by KSC to cope with a detanking
problem. This was caused by a displaced inner tank coupling which
disallowed simple pressure application for liquid detanking and caused
them to detank principally by evaporation. It was suspected that the
prolonged and abnormal operation of the tank heaters during detanking
was directly related to the ultimate failure. As shown in figure 1,
this testing revealed that the heater thermostat switch contacts
welded in the closed position. Consequently, heaters were continuously
adding energy to the system during the simulated detanking operation
which allowed temperatures on the heater probe assembly to rise as
high as 10000 F (see figure 2), which is grossly in excess of the
+800 F thermal switch set point. The insulation of the destratifi-
cation fan motor wires located in this area was severely degraded as
shown in figures 3 and 4. This information was very significant for
incident reconstruction. As a result of this finding, it was reason-
able to postulate that ignition took place somewhere within the heater
probe assembly near the degraded fan motor wires after which com-
bustion continued until CSS failure.
Component Level Ignition Studies
This category of tests was designed to deliberately ignite the teflon
insulation in a simulated CSS to determine the potential CSS failure
modes and to gain a better understanding of system combustion propa-
gation and pressure rise rates. These tests indicated that the flight
structural failure occurred within the electrical conduit section be-
tween the pressure vessel and the outer shell of the accessory dome.
Figure 5 shows the results of these tests. Considerations were now
being given to performing a full simulation test using all the flight
type hardware possible. However, there were still many reservations
concerning the types and quantities of energy releases that could
take place after ignition of an "all up" configuration. Therefore,
additional tests were performed to determine if there would be gross
amounts of energy released from the combustion of metals within the
CSS. For these tests, TTA responsibility was primarily the CSS
heater/fan motor probe assembly. Every reasonable effort to ignite
the fan motors was attempted but no metal combustion was evidenced.
Figures 6 and 7 show the results of some of these tests.
Full Simulation 02 CSS Ignition Test
Subsequently, the "all up" test configuration was established and the
test objectives defined. The primary objective was to determine the
pressure and temperature time history of the combustion of the teflon
wire insulation and the extent of propagation to other materials when
the wire insulation was forcibly ignited near the lower fan motor.
The test article was a flight configured CSS tanked with approximately
900 psi, -190° F supercritical oxygen. The results of this test were
generally as expected. The conduit ruptured 1/4 inch from the
pressure vessel (see figure 8) within 58 seconds after ignition. The
combustion time history, shown in figure 9, was in agreement with the
results obtained from previous component level tests, and.was a re-
markable facsimile of the Apollo 13 accident time history which
occurred within a 75 second period. The difference between one "G"
and zero "G" burning rates for the teflon as well as the uncertainty
of the exact ignition location easily allows for the 17 seconds
difference. The relief valve functioned normally at 1006 psia and the
maximum pressure recorded was 1170 psia. As shown in figure 10, post-
test inspection of the CSS interior showed that essentially all teflon
435
burned except the lower capacitance probe insulator.
APOLLO 14
Material Ignition and Combustion Tests
During the Apollo 13 investigation, and throughout the resulting
Apollo 14 redesign period, the materials test programs continued to
provide useful data. The ignition and combustion tests were performed
at several MSC locations,* with the TTA responsibility being primarily
that of spark ignition studies. Several methods of controlling and
measuring spark energy releases were developed, and spark testing was
performed over a range of voltages, energies, and specimen types.
The test apparatus shown is shown in figure 11, a typical spark
ignition device schematic is shown in figure 12, and the results of
igniting an aluminum specimen are shown in figure 13. The test re-
sults may be generalized as follows:
1. Short term sparking is probable with wire breaks and inter-
mittent shorts and is very difficult to protect against. Normal
fusing techniques are too slow to provide adequate protection.
2. Lowering the voltage available at the gap makes it more
difficult to transfer spark energy at rates sufficient to cause com-
bustion.
3. Practical values for the spark ignition threshold of mate-
rials are very difficult to obtain because they are very dependent on
geometries from the viewpoints of ease of sparking, and heat retention
properties. That is, sharp points and/or low mass extrusions are more
easily ignited than blunt large mass objects.
4. Some ignition threshold values considered useful which were
developed by sparking from needle point geometries in an ambient
temperature high pressure 02 environment are 0.25 joules for teflon,
approximately 3 joules for poor conducting metals, and somewhat higher
energies for good conducting metals.
5. Each design configuration must be independently evaluated to
determine its vulnerability to spark ignition.
436
*Flammability Testing Conducted in Support of Apollo 13, by
L. J. Leger, Materials Technology Branch, MSC, and R. W. Bricker,
Structural Test Branch, MSC.
The Apollo 14 redesigned CSS excluded the destratification fan motors
thereby eliminating a potential high energy spark source. In addition,
all teflon wire insulation was deleted which eliminated most of the
easily ignitable combustible material. The new design incorporated
extensive use of stainless steel sheathed, silicone dioxide insulated
conductors which provided considerably less vulnerability to damage
and combustion than the Apollo 13 installation. All of the elements
incorporated in the redesign which would be exposed to oxygen were
spark ignition tested under many failure mode simulations. The re-
sults were that the Apollo 14 redesign spark tests produced no ig-
nitions when a circuit with resistances and fuse protection like that
in the spacecraft was included.
Alternate Destratification Methods
Following the elimination of the destratification fans, two courses of
investigation were followed: one, providing a substitute for the loss
of the fans, and two, investigating the problems associated with
having no destratification devices.
One of the more promising substitutes considered was an external re-
circulation loop concept. A test was performed to determine the
feasibility of using a modified version of the Apollo Environmental
Control System (ECS) coolant pump to circulate the fluid to destratify
the oxygen in the CSS. Figure 14 shows the disassembled pump. The
pump electrical parts were excluded from contact with the fluid by use
of a magnetically coupled fluid impeller. In the test setup, the pump
was installed such that the oxygen would flow from the vent to the
fill line of an Apollo cryogenic storage vessel. For this test an un-
modified coolant pump was used, and nitrogen was substituted for
oxygen. Various gaseous and supercritical fluid conditions were
tested to simulate the full mass quantity spectrum. A normal usage
outflow of 1.5 pounds per hour was provided during the test. The
feasibility was clearly demonstrated in a low density or 5 pounds per
cubic foot test, in which purposely induced stratification was elimin-
ated by operation of the external loop.
To aid in determining the effects of having no destratification devices
in the cryogenic storage vessel, tests were performed in which KSC pre-
launch tanking procedures were duplicated both with and without heater
operation, without the use of destratification fans, and with the
vessel in a normal vertical configuration. Following each test, the
vessel was agitated to simulate launch vibration; this resulted in a
maximum of 10 pounds per square inch decrease from 900 pounds per
square inch absolute pressure. These pressure decreases were accom-
panied by 3° F decreases in indicated temperature and a 2% decrease
in indicated quantity.
Additionally, an investigation of stratification at low fill densities
was conducted by depleting the cryogenic storage vessel to 35%, estab-
lishing normal usage flow and maintaining pressure for 24 hours. The
vessel was then agitated to simulate engine firing accelerations and
vibrations and no appreciable pressure decay was observed.
The final test in this series was to determine the operational time
available when depleting at the maximum expected flowrate to 150 pounds
per square inch absolute without the use of heaters and starting at a
low fill density (20% full).
The results indicate a time to minimum pressure of 49 hours with the
vessel located in a 600 F environment. Even though two phase fluid
existed somewhere between 554 and 501 pounds per square inch absolute
pressure, it did not cause any operational problems.
Flight Support Tests
To meet a specific Apollo 14 in-flight test objective, it became
necessary to determine the best method for filling the newly con-
figured cryogenic storage vessel to approximately 60%. In the first
method tested, the vessel was filled with liquid oxygen, dumped to
61% using gaseous oxygen pressure, fully pressurized with ambient
temperature gaseous oxygen (quantity increased to 74%), and depleted
at one pound per hour for 40.5 hours. During the depletion phase,
the pressure decreased below the minimum operational limit and re-
quired 5.5 hours to recover although the heaters were on. Figure 15
is a plot of the pressure versus time for this period. The vessel was
subsequently agitated but no pressure decay was experienced.
In the second method attempted, the vessel was filled with liquid
oxygen, pressurized with heaters, depleted at 1.55 pounds per hour
for 63 hours, and further depleted for 44 hours with equilibrium flow
to maintain normal pressure. No abnormal pressure decay resulted from
depletion flows or from agitation of the vessel.
A third method was implemented whereby the vessel was filled with
liquid oxygen, dumped to 51%, pressurized to normal pressure with
gaseous oxygen, and maintained at normal pressure by inflow of gaseous
oxygen for 10 hours. Agitation produced a pressure decrease of 249
pounds per square inch and subsequent repressurizations with heaters
followed by agitations continued to show pressure decreases.
Of the three methods, the second method was chosen since it fit with-
in launch pad operational constraints and produced no instability in
the vessel.
438
CONCLUDING REMARKS
The test programs conducted at the TTA, MSC, in support of the Apollo
13 investigation provided very fast, accurate insight into the cause
of the flight anomaly, as well as important guidelines for the Apollo
14 CSS redesign. These test programs also provided the operational
information necessary to gain a high confidence level in the rede-
signed Apollo 14 CSS.
Some specific conclusions from these tests, which are of interest to
cryogenic systems designers, are as follows:
1. Electrical circuits included in oxygen storage vessels may
be hazardous unless special engineering precautions are taken to
limit the availability of electrical energy and/or to provide geom-
etries demonstrated to be safe.
2. Potentially detrimental fluid stratification effects may be
minimized by using external loop circulation techniques and/or by
special filling and operational methods such as those developed for
Apollo 14.
439
•70-40850 
-r 
o 
Figure 1 . - Welded heater thermostat switch contact 
T1.44-JAu1.40)
I~~~~~~~~~~~~~~~~~~~~~~~~~~
/ 
1-4
I~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~--
jr~~~~~~~~~~~~~~~~~~~~~~~~~a
/~
~
~
~
~
~
~
~
U
fl±N
J.3dW
I£ 
A
lIW
IW
$$ 
ilL¥3V
H
0
$H
N
I~I 3A
31 
ain
o
ill
I~
 
o
,
C
D
~
~
b
00 
El
0
I 000 
I·
441
S-70-40291 
Figure 3.- Most severe degradation to fan motor wire insula 
S-70-40287 
Figure 4.- Average degradation to fan motor wire insulation 
•p-
•F" 
-P-
Figure 5.- Electrical conduit after rupture 
S-70-40036 
•p-
•p-
• M H p 
Figure 6.- Fan motor wire and terminal insulation combustion from u 
S-70-43325 
Figure 7.- Fan motor charred stator winding insulation 
from upward burning 
kk6 
S-70-41292 
-3 
Figure 8.- Electrical conduit rupture after "all up" tes 
zw
a
M
-a,-,frntf- t±T :7r IF T
e!F..T4 I I
L10
OOC
90C
r
M ff-,IZ; v~ =:...,.. I
70(
I-T !- H I X .1 1:- I:: , I-: ,
I _:1 I . i::. _ . :i.
3I FLIGHT RELIEF VALUE PRESSURE
A': "-- . . . ... _ _:. :::: ..  . :
-- 1f-
- I-:
0IT
T1
i ..... .. ~-,-1 .....t
TIm? FROm -0 8 =.16
FIRE PULSE 
IGNITION 0 5 
I ' 1'1
IX P. lOI 5!::
S . -7-
-t 12
Ti i z.zl iizl !
5 24 32 40 48 56 64
-'T [,j:- .... l::: : 1'::..-16 .. ~J :::..l: ....... 2 : W2 ....
16 24 32 kO0 48 56 64
:11 :: ^tl:: l::::_ .. 1:11 n.. __.. :_ _. ~ _.[. --: --. ----|---|I
1·.-.- ._
..- .... _ ..'-  ,, -.. ..,. : . . . .. .:. i "--1-- I i
1E1 .,; . _ · -1 f:- -. I::::i.:. 1
;: : ::: :: :.. . !::: ...... · ,'. . ',. : ... -11TI
. . : :.. : ....' :: :, I. _', .. I ' . .' ' ::
~0-- GSE TA NK PREsSiURE ]:'-.. +:'.
:::
' :: . ' ; " " i I 110W
__ ' ...... ' ' ' . ' . ' . . ~ ~ ~ ~ ~ ~ ~ ~ ' 
~~~~~~~~~~I.. . .I,.[I _8 
.. I
·-200
:1
Figure 9.- "All up" test time history
448
F
'" :::" ' .1 I : I I 1 :;l :A
:: :LJ A . It, f ,::£-Li_
:-!I1
Tie Trr Time fry. i'en Ifr iTe Tire
Event fire pu.s.. iitio n Even ire puLse, igntiton ,
oe s conds second. n. er seco. : Iseconds
Fire pulse applled to ignit- 0.0 -1.99 Wire loop tear r.no... orff. 51.92 19.93|
er on lover fn motor ires scaleIQ ver fr n motor uoses blo. 1.99 0.0 () Fr ht positin tep. ensor | 2.03 50.04
0 ottcale
Upper rfn motor .uses bIovn 26.70 24.71
GSE reliefr vlv. o!-eno t 54.1L 52.15
Wire loop tep. se nsor 1095 PlaI
indicated .1o5 tp. rise 143.33 41. 34
I_ IOSE -li.~ nia_ oprn- S t
Flight positlon tp. sensor 47.93 45.94 1129 psi. 55.14 53.15G | Indicated slov teap. rise L_ l
1 1- am tank pressur- 57.03 55.04O light relief cvlve opens 50.86 18.87 achieved (1197 rpsi)
st 1006 psla
e i-. l w thru flight r.liet 59.02 57.03
alver - I00. no./hr.O .. nk rupture at 1134 p,;. 5999 58.00
. .. - .. - -
. ...... aa- - ;. - wo- 1 ..... ..... . ..... . 11 . = A ... = . .1 1. . I .T
I-
Miaa i';l::~
..... .1 .-X.
h: i:.
-; ': 1i 1~ 1i " 1 1 -44H
LIGHr IANIK ILIETILPERAIURE,
: -11' :::: -1:. :1' -': :
: ' I
=- i
R.-F :..-Fl.l:l r.l>l - -..+ H --,; :. J-i- +. +:::i' I : 3 iiiil_ t l; : l aitw i t 1I Wi;;l ;1 ' - ' 1 -1j'-1 1 1 ' t --
I6
1'
At,
Pmid~
IE
iI
U 
E
7
:T::
I 
:-.4Hi:-: : ..:-
'. r -!iI
-~--:. J:1-: I 1 ,-L Ir
Iw
It
Iw
1
rF IrI r TANu TrrMnrAlTl]r: T.i: iT.11
1:
-1-r!E
::lE:L.: -I r r'·: ·: 
gm
IC l0iV > OT:- s7 !!:-- I1
S-70-41327 
Figure 10.- Top of heater and capacitance probe assembly 
after "all up" test 
kk9 
S-70-40406 
Figure 11.- Supercritical oxygen spark ignition test apparatus 
450 
U0 dJ
U
a
£0a)4J U,to
s.,40S.
.d0 *,40. .dI(4,0)wr:3
451
+
aLc%
El 8
Z4
E
-4
 D-;
I u 
S-70-43575 
/' 
Figure 13.- Results of aluminum burn after spark ignition in a stainless s 
S-70-47003 
VJl 
Figure 14.- Disassembled recirculation pump 
la0o
0 
co
'-4'4 4 7-4
to 
ul
7~~~~~~~~~~~~~~~~~~~~~~~~~~..
/,
Lf)
C4 
(U~~~~r
$4- :1
O
 
~
 
~
~
~
~
~
~
~
~
 
O)
O
 
O
-r4
J~
~
~
~
~
~
~
0
0
~
~
~
~
~
~
~
~
~
~
~
~
c
:'~
~
~
~
~
~
~
~
~
~
~
~
~
c
v
 
_s 
_
 
_
 
_
 
_
n
s
s
C
u
 
c
I_________ I 
s
U,7-
o
~
c
 0o
 
o
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
a
~jsd--ozss~ 
s::
454
000'


A review of cryogenic testing performed by the thermochemical test branch, Manned Spacecraft Center in support of Apollo 13 and14

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server