Apollo Experience Report Earth Landing System
Apollo Experience Report Earth Landing System
Apollo Experience Report Earth Landing System
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CASE FILE
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by Robert B. West
Lyndon B. Johnson Space Center
Hozlston, Texas 77058
* Apollo Spacecraft
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CONTENTS
Section Page
SUMMARY ..................................... 1
i
INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
I
1
~
Command Module Weight Increase . . . . . . . . . . . . . . . . . . . . . . . 5
Main-Parachute-Cluster Interference . . . . . . . . . . . . . . . . . . . . . 8
iii
TABLES
Table Page
I MAIN-PARACHUTE DESIGN LOADS ................... 7
FIGURES
Figure Page
iv
Figure Page
A- 4 Block I pilot parachute .......................... A- 3
A- 17 Drogue-parachute r i s e r assembly . . . . . . . . . . . . . . . . . . . . A- 12
.
Figure Page
B- 6 Deployment envelope
(a) Drogue parachute .......................... B-14
(b) Main parachute . .......................... B- 14
c-11 Results of forward heat shield/suspension system impact test ..... C-22
vi
APOLLO EXPERl ENCE REPORT
EARTH LAND ING SYSTEM
By Robert B. West
Lyndon 6. Johnson Space Center
SUMMARY
The Apollo earth landing system operational requirements were defined through
detailed reviews of the total-mission environments associated with both normal atmos-
pheric entry and the various abort contingencies. These operational requirements and
the necessity for compatible interface with the command module dictated the basic de-
sign and performance requirements of the earth landing system. For example, the high
recovery weight of the command module ruled out the use of a single main-parachute
system because of the lack of experience with single parachutes i n a size needed to
recover the spacecraft within the limitations placed on the rate of descent. In addition,
the Apollo upper-deck structure presented formidable problems for packing and install-
ing a single parachute of the required size.
Although much was known relative to the system requirements during the initial
phases of the program, considerably more knowledge was gained during the course of
the development program. The more significant problems encountered during the de-
velopment of the Apollo earth landing system, the solutions, and the general knowledge
gained from having encountered these problems a r e discussed. A brief description of
the Block I, Block II, and Increased Capability Block II systems and a summary of the
test programs that were conducted are included.
INTRODUCTI ON
In January 1962, the original specifications were released for a parachute recov-
e r y system to be incorporated on the Apollo command module (CM). The original pro-
gram to develop and to provide this system was anticipated to extend over a period of
13 months; however, the final Apollo earth landing system (ELS)qualification test w a s
not completed until July 1968. Then the system was considered suitable for manned
lunar missions.
Early i n the Apollo Program, the fact was recognized that, to accomplish the
lunar-landing mission, certain major changes would have to be made to the initial CM
design, In consideration of program cost and schedular, a decision w a s made to con-
tinue the original CM configuration (then designated as Block I) through the initial
earth-orbital system-verification flights. At the same time, a Block 11 program was
initiated to implement the changes to the CM that were needed to accomplish the lunar-
landing mission.
The major differences between the Blocks I and I1 spacecraft affecting the ELS
were as follows.
1. The docking tunnel w a s shortened and the tunnel wall was tapered slightly in-
ward. This modification significantly changed the shape of the main-parachute stowage
compartment and necessitated complete redesign of the main-parachute deployment bag
and the retention system.
3. The pilot-parachute mortar mounts were moved from the deck of the forward
compartment and relocated on the side of the gussets.
The drogue, pilot, and main parachutes remained essentially unchanged f r o m the
Block I configuration except for a length of steel riser incorporated in the lower end of
the main-parachute riser for protection from abrasion. After the spacecraft 012 f i r e ,
many modifications were made to the Block I1 CM, resulting in a significant increase
in vehicle weight. Analysis indicated that the projected weight increase would result
in parachute loads greater than those that either the ELS or the CM structure was
capable of withstanding safely. Therefore, in mid-1967, a program was initiated to
increase the capability of the ELS and to reduce the main-parachute loads to acceptable
levels with the greater CM weight. The major changes made to the ELS during this
program were a s follows.
2. The diameter of the drogue parachutes was increased to reduce the dynamic-
p r e s s u r e conditions and to provide a more stable vehicle at the time of main-parachute
deployment.
3. The size of the drogue-parachute mortar was increased to provide the addi-
tional volume required by the larger drogue parachute.
The various components of the Block I , Block 11, and Increased Capability
Block 11 systems a r e discussed i n detail in appendix A. In appendix B the test pro-
grams that were conducted to develop and to qualify the ELS are outlined.
2
FUNCTIONAL DESCRIPTION
The ELS consists of the various parachutes and related components necessary to
stabilize and to decelerate the CM to conditions that a r e safe for landing. The ELS was
designed to recover the CM after either a normal entry o r a launch abort. Nine para-
chutes are installed on the CM all of which function during the recovery sequence.
Three main parachutes, three pilot parachutes, two drogue parachutes, and a forward-
heat- shield-separation-augmentation parachute are included.
Drogue-
parachute
disconnect
24 O W f t
baroswitches
System A
*$I:
Crew switch
(main-parachute ~ ~ i ~ -
System E disconnect) parachute
disconnect
TD = t i m e-delay relay
11 am-n
baroswitches
3
heat shield from the CM, a small, 7.2-foot-
diameter parachute is mortar-deployed from
the forward heat shield. This parachute ex-
erts a force to extract the jettisoned heat
shield from the wake of the CM. Two
16.5-foot-diameter conical ribbon drogue
parachutes are mortar-deployed 1 . 6 sec-
onds after f orward-heat- shield jettison.
The drogue parachutes undergo a 10-second
reefed interval before disreefing to full
Splashdown velocitier
Three parachutes - 32 fps
Two parachutes - 36 fps
%@ open, and remain attached to the CM to an
altitude of approximately 11 000 feet. At
1. Forward heat shield jettisoned at 24 Mx) R drogue disconnect, three 7.2-foot-diameter
+0.4 sec
2. Drogue parachutes deployed reefed at
ringslot pilot parachutes are mortar-
24 OOO A +2 sec deployed. The pilot parachutes then pro-
3. Drogueparachute single-stage disreef,
10 sec
vide the force necessary to release the
4. Drogue parachutes jettisoned and main main-parachute retention system and to
parachutes deployed by pilot para-
chutes at 11 M)o A
extract the main-parachute -3ck assemblies.
5. Main-parachute initial inflation The main-parachute packs &rethen pulled
6. Main-parachute first-stage disreef,
6 sec
away from the CM and the three 83.5-foot-
7. vhf recoveryantennas and flashing diameter main parachutes are extracted
beacon deployed, 8 sec
8. Main-parachute second-stagedisreef,
from the deployment bags. Each main para-
10 sec, approximately 9 Mx) R chute then inflates through two reefing stages
9. Main parachutes released
to a full-open condition.
Figure 2. - Normal recovery sequence. If a launch abort should occur and
conditions are such that the maximum alti-
tude attained is below the opening altitude of the baroswitches, forward-heat-shield
jettison and deployment of the drogue and pilot parachutes occur on a timed sequence,
controlled by the time-delay relays. The events that occur during a launch abort are
illustrated i n figure 3.
4
~
system. Through the applied efforts of the various groups and agencies associated with
the system, these problems were resolved with minimum design changes and with the
least impact on the Apollo Program.
I
In addition to resolving difficult design problems, devising and optimizing compo-
nent manufacturing and assembling techniques were also necessary to ensure that each
part would function properly once it was assembled and installed on the spacecraft. On
none of the previous space programs w a s it necessary to contain the parachutes i n the ,
limited volumes and in the irregular shapes necessary in the Apollo Program. This
requirement necessitated the development of precise techniques for packing the para-
chutes at very high densities without inflicting damage that could propagate during de-
ployment. The incorporation of steel cable as an integral part of the parachute risers
required the development of stowage techniques that would provide assurance that the
cable deployed consistently and safely.
m
I tained on two of the three tests and because
of the first announcement of a CM weight
I
creased the strength of the structural
1 Figure 4 ' - main-parachute load members of the parachute. These changes
compared with CM weight. caused a significant increase i n weight and
I
5
5 11-
- 10--
'ij
II
14 x 10
= 1312 t-
3
- 13 OOO
Increased Capability
created new problems because of limited
stowage volume. Shortly after the start
of main-parachute-cluster tests , modifica-
tions had to be made to the main parachutes
to change their opening characteristics to
.E
g- 8
9
8 150 Block II qualification achieve more evenly balanced load sharing
Block I
Block II
qualification
among the parachutes , thereby reducing the
% 7-
'
u
5
-
-
qualification
First ELS aerial
drop test
peak opening loads,
Year
pleted qualification, each system of the
spacecraft had progressed to the point at
Figure 5. - Increases i n CM recovery which accurate total-weight estimates were
weight. available. Although the maximum projected
weight for a Block II spacecraft was above
the specification values, the overweight
condition was not sufficient to justify major design changes in the ELS. Therefore,
the Block II program was pursued as a minimum-change effort.
During the months immediately following the spacecraft 012 accident, numerous
modifications were made to the CM. By mid-April 1967, weight estimates indicated
that the projected weight of the spacecraft had increased to a value greater than that at
which the ELS could recover the CM with an acceptable safety margin. The CM recov-
ery weight possibly could increase to as much as 13 000 pounds, which, i n turn, would
increase the parachute loads to levels unacceptable for the parachutes and the CM
structure.
The implemented solution consisted of increasing the size of the drogue para-
chutes and of providing the existing main parachutes with an additional reefing stage.
The larger drogue parachutes and the additional main-parachute reefing stage were
necessary to ensure adequate safety factors for the parachutes and the CM structure at
the 13 000-pound recovery weight. Larger drogue parachutes on the heavier CM re-
duced the dynamic pressure at drogue disconnect/pilot mortar fire to a level near that
of the smaller drogue parachutes on the lighter spacecraft. The additional reefing
stage i n the main parachutes reduced the individual and total main-parachute loads to
values no greater than the design loads for an 11 000-pound CM.
On July 31, 1967, a program was approved to implement these changes to the
ELS to become effective on the first Block II spacecraft (CSM 101). At the beginning
of the improvement program, very few data were available to provide a basis f o r a
detailed analysis of two-stage main-parachute reefing. Therefore, aerial drop tests
had to be conducted early in the program. These tests generated the preliminary per-
formance data that were used to establish reefing parameters and t o support an evalua-
tion Of the adequacy of the existing main-parachute design. From these initial tests,
two-stage reefing w a s proved feasible and the existing main-parachute design w a s
proved structurally adequate with the higher CM e i g h t . Reefed intervals of 6 seconds
in the first stage and an additional 4 seconds i n the second stage of inflation were se-
lected. These reefed intervals provided nearly uniform loading i n each stage and added
only 2 seconds to the total inflation time of the earlier single-stage reefing system.
The new reefing-time-delay requirement was then factored into the development of a
6
reefing cutter. In conjunction with the main-parachute-design changes, definition of the
maximum volume available for larger drogue-parachute mortars became necessary be-
cause the size of the drogue parachute w a s limited primarily by the size of the mortar
tubes that could be fitted to the spacecraft. Volume w a s made available for mortars
accommodating 16.5-foot-diameter drogue parachutes.
Design loads, lb
(a
Parachute configurations
Original Block 11 ELS Final Block I1 ELS
with 11 000-lb CM with 13 000-lb CM
Individual
a
These values should not be used for direct comparison.
7
Main-Parachute-Cluster Interference
On November 27, 1962, the first aerial drop test was made using a parachute test
vehicle (PTV) to investigate the performance characteristics of a cluster of three inde-
pendently deployed 88.1-foot-diameter ringsail main parachutes. For the initial test,
the vehicle w a s ballasted to only 4750 pounds, one-half the weight of the CM. The con-
figuration of the parachute test specimens and the related components represented the
then-current spacecraft design. Deployment of the main parachutes was achieved by
simultaneously mor t a r - deployed pi lot parachutes .
The results of this test indicated a significant problem relative to the deployment
behavior of clustered ringsail parachutes and eventually led to some unique design fea-
tures i n the main parachutes. The three ringsail parachutes inflated in a nonsynchro-
nous manner, that is, one canopy inflated rapidly and inhibited the filling of the lagging
parachutes. This behavior was most pronounced during the inflation following disreef .
This crowding effect and nonsynchronous inflation, often referred to as cluster inter-
ference, was not a new phenomenon but was unusually pronounced with the ringsail de-
sign. This uneven load sharing resulted in abnormally high opening loads on the
leading parachute i n the cluster.
The approach taken to correct this condition was to explore modifications that
would reduce the rate of inflation of the parachute. Much of this rapid inflation of the
ringsail parachute w a s attributed to a characteristic of the canopy to continue to fill
during the reefed interval and thus produce a large reefed shape with internal pressures
throughout a large portion of the canopy. This characteristic, i n turn, contributed to a
rapid full-open inflation following disreef. The flow of air around the large bulbous
shape of a rapidly developed leading parachute w a s also noted to have a distorting ef-
fect on the adjacent lagging parachutes and to greatly inhibit the inflation in the reefed
condition.
The modificatioh demonstrating the most favorable effect in reducing the cluster
interference was the removal of 75 percent of the material from the fifth ring of the
canopy, thereby forming an open ring around the periphery of the crown. This open
ring limited canopy growth in the reefed condition to a more cigar-like profile and pro-
duced near-uniform growth of each of the three main parachutes during reefed and dis-
reefed inflation. A second change, greatly improving the shape of the lower skirt area
and allowing a more efficient inflow of air into all three parachutes, was the relocation
of the reefing rings on the skirt band from the radial seam-attachment point to a point
on the skirt band in the middle of the gore (referred to as midgore reefing).
Based on test results obtained during this s a m e effort, a decision was also made
to remove four complete gores from the main-parachute canopy to reduce weight. This
modification produced the basic 68-gore-configuration main parachute that would even-
tually be flown on Apollo spacecraft.
The combination of reduced peak opening loads and the removal of material re-
sulted i n a net allowable weight saving of approximately 45 pounds i n the main para-
chute and harness assembly without exceeding the specified maximum r a t e of descent
of 33 f p s at 5000-foot pressure altitude. The open-ring-configuration main parachutes
also reduced the system oscillations of the two-parachute cluster configuration f r o m
approximately t 20" to t So, causing a reduction i n landing hazards.
8
Parachute Riser Abrasion
On September 6, 1962, during the fifth total-system developmental test, using the
boilerplate (BP) 3 test vehicle, a s e r i e s of events occurred that resulted in separation
of the main parachutes and loss of the vehicle. The BP failed to stabilize during the
drogue-parachute interval, and the pilot mortars fired with the vehicle in an apex-
forward (approximately 60 ") flight attitude. The existing main-parachute r i s e r s , which
were joined at a confluence above the vehicle, hung under the airlock with the vehicle
i n this flight attitude. This anomaly prevented full deployment of the harness until the
vehicle rotated to a more favorable attitude. The parachute opening loads were trans-
mitted through the unprotected textile harness legs directly into the gussets and the air-
lock structure, promptly severing the harness legs.
The incorporation of the flowerpot parachute attachment and the disconnect fitting
on the Block I1 spacecraft made redesigning the main-parachute risers necessary
(fig. 6). The flowerpot concept brought the two drogue-parachute r i s e r s and the three
main-parachute risers into a common fitting. This concept reduced the available
volume and eliminated the possibility of using the bulky Block I-type protected-fabric
riser on the Block 11 spacecraft. Stowage tests on a Block I1 upper-deck mockup indi-
cated that steel-cable main-parachute r i s e r s could be incorporated and stowed i n a
manner assuring orderly deployment. Because the steel-cable main-parachute riser
provided a solution to the Block 11 volume problem and afforded the necessary resist-
ance to abrasion, the decision was made to incorporate it i n the Block 11 system.
9
Drogue parachute
1 In several of the Block I1 aerial drop
tests incorporating the flowerpot parachute
attachment fitting, numerous cable strands
were damaged on the drogue risers. This
damage occurred in the portion of the cable
riser that contacted the lip of the flowerpot
fitting while the vehicle w a s oscillating and
the parachute loads were high.
A 4-inch-wide band of lead tape, wrapped around each of the drogue riser cables
where they crossed the lip of the flowerpot, provided excellent protection from abra-
sion and proved to be a simple and an effective solution to the problem. In the Blocks I
and 11 systems, the steel-cable parachute risers proved an acceptable design and func-
tioned correctly on all tests and flights.
Because of the inherent inflation characteristics of the main parachute, the con-
ditions of this test produced much higher local-stress levels i n the canopy skirt than
were experienced with conditions at a much higher axial load on previous tests. This
10
deficiency was corrected by the addition of circumferential reinforcing tapes to
strengthen the canopy (fig. 7). This test failure clearly illustrated the necessity for
demonstrating the inflation characteristics and the strength of the parachutes over the
entire range of possible operating conditions.
11
Because of the damage being sustained, a concerted effort w a s made to design
handling equipment and to establish procedures minimizing packing damage. During
this effort, approximately 70 trial parachute-packing operations were performed, in
addition to the parachute packing that was being accomplished to support the develop-
mental aerial drop tests. Steps taken to correct this problem were as follows.
3. The design of the packing pressure feet and protective pads was modified
many times.
7. Various lubricants (Teflon spray, Tef lon-coated cloth, et cetera) were used
on the walls of the packing container to aid in packing under pressure.
11. Padded reefing-cutter packets were developed and incorporated into the
parachutes.
Because of the high densities achieved and because of the tendency for the para-
chute packs to expand when removed f r o m the packing container, form blocks and
vacuum packaging were necessary for storage of the parachute packs to ensure a proper
fit of the parachute pack during installation on the spacecraft.
12
The high-density parachute-packing equipment and the procedures developed and
refined throughout the program allowed for maximum use of the available stowage
volume on the spacecraft. Although minor packing damage (small cuts and burns) is
still present, it is reduced to a level no longer considered to be a hazard o r problem
to the parachutes.
The Block I1 forward heat shield w a s captured in the wake of the vehicle, and the
r e v e r s e flow, which w a s stronger than anticipated, forced it back toward the parachute
compartment. Because this condition represented a serious hazard to the ELS, appro-
priate corrective action was again taken through the addition of a Block 11 forward-heat-
shield-separation-augmentation parachute system (fig. 8).
Because this condition was observed on the final qualification drop test for the
original Block II system, and because no unmanned Block 11flights were scheduled,
i concern was focused on a method to demonstrate the fix. This problem was resolved
when the Increased Capability ELS program provided a means of evaluating +heper-
f ormance of the forward-heat-shield- separation- augmentation system during total-
system aerial drop tests.
i
i
13
Parachute pack Forward-heat-sh ield
jettison system
Mai n-Parachute Oxidizer
Redundant Two cartridges B u r n Damage
One Block I AS 202 main parachute,
retrieved following the spacecraft landing,
revealed many small burn holes throughout
the canopy and suspension lines. Labora-
tory analysis revealed that the parachute
had been damaged by nitrogen tetroxide
(N204)expelled from the CM reaction con-
t r o l system (RCS). The postflight investi-
gation disclosed that the ratio of fuel and
oxidizer carried i n the CM was such that,
-Y ’ L L a n d y a r d system
Lanyard
during the burn and dump modes (used to
purge the RCS after main-parachute deploy-
Electrical harness
Disconnect ment), the fuel was depleted before the oxi-
Redundant dizer, causing raw oxidizer to be dumped
Figure 8. - Block II forward-heat- into the air stream. The N204, which is
shield parachute system. very damaging to nylon parachute material,
was then sufficiently concentrated on the
parachute to burn many small holes. This
condition was corrected on subsequent missions by controlling the- ratio of fuel and oxi-
dizer loaded on the CM to ensure that the oxidizer would be depleted before the fuel
during the burn mode. Thus, only the fuel remained, and it does not react chemically
so as to degrade the strength of the nylon.
An anomaly occurring during the recovery of the Apollo 15 CM caused one of the
three main parachutes to collapse during the final descent. The posff light investigation
revealed three potential causes for the anomaly: (1)a possible collision of the jetti-
soned forward heat shield with the main parachute, (2) a failure of the suspension line
to the r i s e r connector links, and (3) RCS fuel burning the fabric riser o r suspension
lines. This investigation was seriously hampered because only one of the three main
parachutes was recovered following the landing, and the recovered parachute was not
the parachute that collapsed. Secondly, an onboard camera that clearly showed the
sequence of the parachute deployment was turned off just before the failure occurred.
After a thorough analysis of the existing data, and after considerable testing, the
following conclusions were reached,
1. Although it passed very close to the descending spacecraft, the forward heat
shield did not contact the Apollo 15 main parachute.
2 . Although faulty connector links were found in the recovered main parachute,
failure of connector links did not cause the main parachute to collapse.
3. The most probable cause of the anomaly was burning monomethyl hydrazine
(expelled from the C M RCS) coming in contact with the main-parachute fabric riser.
A complete description of this anomaly and the subsequent findings and corrective ac-
tions a r e presented as appendix C.
14
CONCLUDING REMARKS
The performance of the ELS during many rigorous tests, and the performance of
each component of the ELS during the spacecraft flights proved that the program objec-
tives had been met. A parachute-recovery system was provided that satisfied the mis-
sion requirements, was compatible with the physical characteristics of the vehicle, and
had a high degree of reliability. The increasingly severe demands placed on the ELS
i n t e r m s of recovery weight, limited stowage volume, and a wide range of initial con-
ditions were met and resolved without undue impact on the overall Apollo Program.
15
APPENDIX A
The basic system concept of the ELS changed little throughout the program from
that described in the original statement of work. However, the individual components
of the system underwent numerous changes as the program progressed through the
Block I and 11phases. The following is a brief description of the major components of
the Block I, Block 11, and Increased Capability Block II ELS.
BLOCK I ELS
Drogue-Parachute Systems
The drogue parachute, a 13.7-foot-diameter conical ribbon type, was actively
reefed t o 39.3 percent of its nominal diameter by redundant reefing lines with two
8-second reefing cutters per line. This parachute is illustrated in figure A-2. The
drogue-parachute mortar assembly is illustrated in figure A-3. An electric current ~
deployment bag; the steel-cable portion of the riser, contained in a polyurethane foam
ring, breaks out of the foamed ring- and uncoils. The parachute then goes through a
n o r i a l inflation process.
i
I
-Cover I
Deployment bag, Ring, foamed rise;- 1
t t m-drogue parachute
i
i
A- 2
P ilot- Pa rac hUte System
The pilot parachute, which extracts and deploys the main parachute, is illustrated
i n figure A-4. This was a 7.2-foot-diameter ringslot design permanently attached to
the main parachute and main-parachute deployment bag. The pilot-parachute mortar,
pack assembly, and related components are illustrated in figure A-5. As with the
drogue system, the pilot parachute was deployed through the action of expanding gas in
the mortar tube. The pilot parachute then
provided the force required to release the
main-parachute retention system and ex-
tract the main-parachute pack from the CM.
Ring, foamed riser - \\ -h
Pack assembly
Sabot assembly
Breech
1 Tebu-J assembly
A- 3
- Deployment bag,
main parachute
7.25 Ib I 1 -
125 0 Ib
1
parachute released a chain lace securing the
retention-assembly center panel to the two
side flaps and to the upper flap. Release of
the chain lace allowed the center flap to open
outward and release the main-parachute pack.
No. of gores: 68 A s the pilot parachute lifted the main-
Calculated total porosity: 12 percent
Reefing: 7 . 9 R DR (9.5 percent midgore) parachute pack away from the vehicle, the
Reefing duration: 8 sec main parachute was extracted from the de-
Drag area reefed: 353 ftZ ployment bag in an orderly manner beginning
Drag area disreefd: 4MxJ ft2
with the connector links, followed by the
suspension lines, and finally by the canopy.
Figure A-6.- Block I main parachute.
A- 4
v/
Riser to main-oarachute n
pack, +Z bay Main -Pa rac h Ute Attach ment Fitting
Riser to main-
parachute pack, The two main-parachute attachment
fittings were located on the upper deck of
the C M at the base of longerons 3 and 4.
Each of the two fittings was capable of with-
standing the total opening loads generated
by the three main parachutes. A main-
parachute-harness disconnect wasincorpo-
rated into each of these attachment fittings.
This unit included a pyrotechnic pressure
cartridge which, when initiated, drove a
sharp blade into the harness retaining pin,
severing that leg of the harness and releas-
ing it from the vehicle. The main-parachute
attachment fitting and harness disconnect
a r e illustrated in figure A-9. As a safety
feature, the two disconnects received their
current from two separate electrical
Figure A-8. - Block I main-parachute sources; thus, an inadvertent premature
harness assembly. signal would disconnect only one leg of the
harness, and the main parachutes would
remain attached to the CM through the other
leg.
\
Disconnect
cartridge
pressure A Reefing Cutters
The 8-second reefing-line cutters
used i n the Block I drogue and main para-
chutes were identical and interchangeable.
vehicle The reefing cutters were used to sever the
harness
reefing line, which limited the inflated di-
ameter of the parachute for a predeter-
mined time, thus reducing the parachute
opening loads.
A- 5
r M o u n t i n g ring
end of the time-delay burn, a pyrotechnic
charge ignited and caused a cutter blade to
sever the reefing line. This same type of
reefing-line cutter was also used to release
Time-delay compound
(three to four pressings)
the VHF recovery antennas and the flashing
light, allowing them to deploy 8 seconds
Expansion chamber
after the main parachutes were deployed.
BLOCK II ELS
The general arrangement of the Block I1 ELS installed on the upper deck is de-
picted in figure A- ll. The Block I1 installation used the available volume in the forward
compartment more efficiently than did the Block I system. The incorporation of the
flowerpot parachute attachment fitting elim-
inated the large, bulky, two-leg, main-
parachute harness assembly and the
-Sea recovery confluence fitting and provided for much
better stowage of both the main- and drogue-
parachute r i s e r s .
A- 6
The Block I main-parachute retention flaps were replaced by a series of fabric-
covered, spring steel straps attached to the upper-deck structure. These straps were
then chain-laced by an interlocking length of cord to loops sewed to the face of the main-
parachute deployment bag (fig. A-12). Because the Block I retention flaps also pro-
vided environmental/thermal protection for the main parachutes, eliminating these
flaps necessitated the incorporation of these protective features into the Block II de-
ployment bag. This task was accomplished by adding a layer of 1/4-inch Dacron felt
to the walls of the deployment bag.
A- 7
Block I 1 Parachute Riser Assemblies
The Block II main parachutes incorporated r i s e r assemblies composed of steel
cable and fabric that attached each of the three parachutes directly to the CM. The
steel-cable portion was 80 inches long and was composed of six 0.28-inch-diameter
cables, assembled with swaged-steel end fittings. These end fittings were designed to
attach one end of the cable riser to the flowerpot fitting on the CM and to attach the
other end to the main-parachute fabric r i s e r . The entire cable riser, except the end
fittings, was encased in a thermally fitted, polyolefin sleeve to maintain the six indi-
vidual wire cables in their relative positions during stowage and deployment. The
fabric portion of the main-parachute r i s e r assembly remained generally the same as
in the Block I configuration, except for a slight reduction in length.
A- a
Two- Stage Main -Pa rachUte Reefing
The significant difference between the Reefin
original Block II main parachute and the In- line B Reefing
cutter
creased capability Block II main parachute Reefing cutter Reefing
line A
was the incorporation of two-stage reefing.
The two-stage reefing system, illustrated
in figure A- 13, reduced the peak opening First stage Second stage
loads generated by each of the main para-
chutes. The incorporation of the additional
reefing stage reduced the initial drag area
present at high-dynamic pressures and then Line A
allowed the parachute to inflate to a larger - Shorter length (264 in.)
intermediate drag area before inflating to 6-second cutters
A-9
Although single-stage reefing of parachutes has become a rather common practice
f o r reducing opening loads, the use of two-stage reefing was virtually untried before its
incorporation in the Block 11 system. This system proved to be effective and reliable ,
i n reducing the parachute opening load below what otherwise would have been encoun- I
tered because of the increased weight of the CM. The final Block 11 main-parachute
configuration is illustrated in figure A- 15.
A- 10
Deployment bag,
drogue parachute continuous ribbons with a single-lap splice
eliminated the added material previously
required to make multiple-gore splices.
A-11
Because of an increase in the drogue-parachute design loads from 1 2 600 to
17 200 pounds, the steel-cable portion of the drogue r i s e r had to be strengthened. This
strengthening was accomplished by replacing the three strands of 0.28-inch cable by
four strands of 0.25-inch cable, and by incorporating the necessary changes to the r i s e r
and fittings (fig. A-17). TO reduce the abrasive action between the steel-cable riser
and the spacecraft flowerpot fitting, a lead-
tape wrapping-was added to drogue cable
risers where they contact the flowerpot
fitting.
T
180 i n
diameter
Drogue Mortar-Assembly
Modifications
The drogue mortar function for the
Increased Capability Block I1 system was
identical to that of the Blocks I and 11 con-
figurations. The major difference was that
the new mortar assembly was designed to
deploy a larger and a heavier drogue para-
steel cables chute (fig. A-18).
A- 12
L
-Deployment bag,
pilot parachute
were modifications to strengthen the pilot .33 Ib
parachute. The parachute (fig. A-19) w a s
modified to increase the strength specifica-
tion of the suspension lines from 400 to
600 pounds, and was changed from a
multiple-ply-webbing riser to the integrated
suspension-line and riser configuration
5
similar to that used on the drogue para-
chute. The multiple-ply-webbing riser de-
sign was unsatisfactory because of
susceptibility to improper manufacturing.
In securing the plies of webbing, uneven
gathering of a single ply of webbing often
occurred. This condition was difficult to
detect by visual inspection and resulted i n
1 ".I in: ,I> Steel riser
A-1 3
The parachute used was a convention-
al Apollo 7.2-foot-diameter ringslot para-
chute identical to that used to extract the
main parachutes. To reduce the loads,
Ring, foamed riser- Cover 7
the parachute incorporated fixed o r perma-
nent reefing to reduce its effective drag
area. Because of volume limitation in the
CM forward compartment, the convention- I
al cylindrically- shaped pilot mortar could
not be used. By relocating certain equip-
ment in the forward compartment, it was
possible to install the elliptically shaped
LL
forward-heat- shield mortar assembly LOrifice
illustrated in figure A-21. The Block II Pressure cartridge (two each, not shown)
forward- heat- shield- separation- Breech
augmentation system was successfully dem-
onstrated during the Increased Capability Figure A-21. - Block I1 forward-heat-
Block I1 test program. shield mortar assembly.
A- 14
APPENDIX B
SUMMARY OF TESTS
Each component of the Block I, Block 11, and Increased Capability Block 11 ELS
was subjected to extensive testing during the developmental and qualification phases of
the programs. Considerably more developmental testing was associated with the
Block I effort because this program established the basic designs from which the
Block 11 and the Increased Capability Block 11 system evolved. However, the same
general approach and testing techniques were applied during each of the three programs.
In defining the tests for the ELS, two basic requirements had to be satisfied.
One requirement was to demonstrate that each component and subassembly of the total
system was capable of withstanding the total-mission environment and functioning prop-
erly with adequate margins and within specified tolerances. The second requirement
was to demonstrate that the total system would function properly in all potential flight
modes and that a safe interface existed between the various components of the CM and
the parachute recovery system.
The nature of the ELS made it impractical to combine these two requirements
into a single test program where the parts could be subjected to mission environments
in a laboratory and then to operating conditions in an aerial drop test. Considering
also such factors as the amount of time and handling required between the laboratory
and the test, this approach would not have been valid, nor would the results have been
representative of conditions encountered in an actual mission. Therefore, two separate
and somewhat independent test programs had to be conducted, that is, the aerial drop
tests and the laboratory tests.
The laboratory testing conducted during the program was generally confined to
individual component and lower assembly-level tests. Each component was evaluated
in t e r m s of potential loss of strength or performance (or both) resulting from exposure
to the environments of the Apollo mission and the interface between that component and
the spacecraft. In the laboratory, much effort was expended in testing various compo-
nents to support the selection of the most promising designs, to support the failure
analyses, and to obtain performance data on new designs. A f t e r each component had
demonstrated the required level of performance, it was subjected to qualification testing.
Each qualification test article w a s manufactured, inspected, and accepted as if it were
spacecraft hardware. During the qualification tests, if any item failed to meet the
B- 1,
prescribed levels of performance, the failure was formally reported, and a thorough
analysis was performed to determine the exact cause and the necessary corrective
action.
B-2
~
During the original Block II program, emphasis w a s placed on ground testing to
thoroughly evaluate the changes made from the Block I design. These ground t e s t s
were used extensively to obtain comparative performance data on several main-
parachute-deployment-bag configurations being considered. The ground-test approach
considerably reduced the number of required aerial drop tests and reduced the program
cost.
Following the decision to modify the Block I1 ELS to increase its capability to re-
cover the heavier CM, the establishment of ground-test requirements was necessary.
Because each component of the ELS had been qualified previously for use on manned
spacecraft, the extent to which retest was necessary was governed strictly by the na-
ture and the extent of individual component redesign. F o r example, modifications
were made to the reefing cutters to vary the delay time from 8 to 6 and 1 0 seconds.
This change affected only the pyrotechnic time-delay element in the cutter, not the
structure o r actuating mechanism. Therefore, the redesign reefing cutters were sub-
jected only to the mission-environmental test conditions that could influence the per-
formance of the time-delay element, that is, acceleration, high and low temperature,
and high humidity.
The aerial drop tests were conducted at the A i r Forcernavy Joint Parachute Test
Facility, El Centro, California. This facility was ideally equipped for Apollo-type drop
testing because it provided a fully instrumental test range, an onsite shop, and adminis-
trative office space. Sources for data acquisition included ground- to-air and air-to-air
photographic coverage, ground cine- theodolite tracking stations, and a telemetry
ground station. In addition, the El Centro test facility provided many of the drop air-
craft and the test-vehicle ground-handling equipment. All BP vehicle drop tests were
made from a modified C-133A aircraft provided by NASA and manned by contractor
personnel.
Three basic types of vehicles used in the aerial drop tests included an instru-
mented cylindrical test vehicle (ICTV), a parachute test vehicle (PTV), and the boiler-
plate (BP) test vehicles. Often referred to as a bomb-drop vehicle, the ICTV w a s
simple in concept, rugged in construction, and low in cost (fig. B-1). The ICTV was
used on tests where the CM interface was not a consideration and where it provided
simplicity; minimizing test-preparation time. These vehicles were usually equipped
for telemetry and m5aard photographic dzita acquisition.
B- 3
Figure B-1. - Instrumented cylindrical test vehicle.
Well suited for testing at the total-system level, the PTV was designed to simu-
late the major features of the spacecraft upper deck (fig. B-2). Below the deck level,
the PTV was a simple cone shape of sturdy construction to eliminate impact damage.
Because the total drag area of the PTV was much less than that of a CM, this vehicle
was well suited for conducting system-level tests at dynamic pressure conditions above
the limits for spacecraft.
B-4
Figure B-2. - Parachute test vehicle.
The BP test vehicle was the most elaborate spacecraft- representative test vehicle
used in the parachute drop tests (fig. B-3). The test accurately simulated the CM
weight, the center of gravity, and the geometric profile. In addition to testing the total
parachute recovery system, the test could incorporate the forward-heat-shield jettison.
A l l the various components interfacing with the ELS (location aids, vehicle-uprighting
bags, and other spacecraft components) that could affect, or be affected by, the para-
chutes were installed on the BP. The BP was the first vehicle in which the spacecraft-
configuration ELS sequence controller performed the system-sequencing functions.
Instrumentation in the BP w a s similar to that used on the PTV, consisting primarily of
telemetry and onboard cameras.
B-5
Figure B- 3. - Boilerplate test vehicle.
During the Blocks I and 11programs, the aerial drop tests were classified as being
either developmental tests o r qualification tests. Developmental tests were further
I
grouped into individual test series according to specific test objectives.
The aeria1,drop tests made during the Block I developmental program a r e illus-
trated in figure B-4. During the first year of the Block I developmental effort, many
single- main-parachute ICTV drop tests were made to evaluate various main-parachute
design concepts. In 1963, much of the developmental testing consisted of multiple
1%2 1963 1964 1%5
eries Series objectives
J A S O N D J F M A M J J A S O N D F M A M J J A S O N D J F M A M
1 Preliminary developmental tests, one T
main parachute
3 Developmental tests, one main parachute m T lCTV
6 Developmental tests, one main parachute m nmm T T ? W v PN
A BP test
7 Developmental tests, three main cluster V W vehicle
13 Developmental tests, three main cluster vv
17 Drogue tests, one and two parachutes v v V
parachute (cluster) t e s t s and the higher-level system tests. During these tests, the
main-parachute-cluster interference problems became a concern. Several two-main-
parachute ICTV drop tests were conducted during the first 6 months of 1964 to evaluate
the main-parachute modifications incorporated to improve cluster inflation. Following
a limited s e r i e s of total-system verification tests, the Block I developmental drop test
was concluded early in 1965 with drop tests to demonstrate the ultimate strength capa-
bility of the drogue- and main-parachute designs.
A s the time approachdd f o r the qualification aerial drop tests to begin to support
the spacecraft flights, certain qualifiable configuration components apparently were
not going to be available f o r the initial tests. Problems encountered during the latter
part of the developmental program required that some late changes be made to certain
components; time was insufficient to incorporate all these changes in the initial system-
qualification drop tests. Although a basic rule for qualification testing states that items
subject to qualification be of the spacecraft configuration, compelling schedules neces-
sitated initiation of total-system testing with certain items that were not yet in a quali-
fiable configuration. These items included the drogue and pilot mortar cartridges,
the main-parachute disconnect and cartridge, and the reefing-line cutters. The use of
these interim components was permitted on the basis of similarity to final design and
on the fact that the flight items were to be subjected to laboratory qualification tests.
The plan w a s to phase the final-design components into the qualification drop-test
program at the earliest date.
B- 7
BLOCK I SYSTEM-QUAL1 FICATION DROP TEST
From May 6, 1965, to February 24, 1966, 12 aerial drop tests were made to
complete qualification of the Block I ELS and to rate the system suitable for manned
flights. A summary of this program is presented in table B-11.
60-1
62-1 1 /“E
1 5-6-65
6-3-65
1 Normal entr)
a1ti tude
b62-3
62-2 I 8-5-65
8-19-65
Pad a b o r t
IMediuni-
altitude
abort
altitude
abort
altitude
aDynamic p r e s s u r e . b T e s t 62-3 p r o g r a m e r parachute failed to disconnect. ‘Test 62-4 failed to m e e t requifed t e s t conditions
B-8
In these tests, the service ceiling of the C-133 drop-test aircraft (30 000-feet
maximum altitude) and other limitations would not allow complete duplication of the
operational flight modes. F o r example, i n an operational normal entry o r high-altitude
abort, the drogue parachutes would normally be deployed by closure of the high-altitude
baroswitches at approximately 24 000 feet. In a drop-test simulation of these opera-
tional modes, the BP test vehicle w a s released at 30 000 feet. To allow for a minimum
auxiliary-brake parachute- stabilization interval and some-finite free-fall interval to
achieve the required dynamic pressure condition, drogue-parachute deployment (initi-
ated by an auxiliary events controller) had to be delayed to altitudes moderately below
the normal 24 000-foot level. A s a result, two compromises were made in the total
performance of the ELS, that is, recovery was not initiated by the sequence-controller
barometric switches, and the normal drogue-parachute operating interval was reduced
from a normal 50 to 55 seconds to approximately 25 seconds. The first compromise
w a s reconciled by monitoring the sequence controllers for proper closure of the baro-
metric switches, thus acquiring evidence of satisfactory performance. Analysis
showed that the lack of the full drogue interval would not have a significant effect on
the test conditions at the time of main-parachute deployment because there would be
no substantial improvement in vehicle dynamic stability after the 25-second test inter-
val. Secondly, the shorter drogue interval represented a more demanding condition
pertaining to total-system operation than that which would be experienced in a compara-
ble spacecraft operation.
Although desired dynamic pressure and vehicle attitude could be programed into
the tests, a representative flight-path angle could not be achieved. Also, the marginal
stability of the BP vehicles made the acquisition of desired initial attitudes and CM
dynamics at the end of long free-fall intervals very difficult. At the end of the free-
fall interval, these test limitations generally resulted in vehicle dynamics more severe
than those predicted for an actual mission. Because the test conditions were more
severe than those the spacecraft would experience, the tests were judged as a satisfac-
tory demonstration of the systems capabilities.
The major changes f r o m the Block I to the Block 11 ELS concerned the redesigned
main-parachute deployment bag and retention system, the steel-cable main-parachute
riser, the flowerpot parachute-attachment fitting, and the modified pilot-parachute
mortar. The Block 11 developmental drop-test program was oriented to demonstrate
that these modifications did not degrade the strength o r the performance of the system
in any way.
B-9
BLOCK I I QUAL1 FICATION DROP TESTS
From October 19, 1966, to January 17, 1967, a series of four total-system
aerial drop tests was conducted using the BP-6 vehicle, modified to the Block 11 con-
figuration. This s e r i e s of tests (table B-111) completed qualification on what was then
believed to be the final ELS. Before entering this test series, a basic ground rule was
established stating that only qualified o r qualifiable parachute system components and
installation specifications would be used in the test series. This policy was maintained
throughout the test series; in contrast to the Block I test series, no configuration
changes were made to the hardware during the Block 11 qualification drop-test program.
~~
13-4 10-19-66 High-altitude 1 1 000 No 26.8 37.6 24.7 22. 5 114.8 36.4 10.8 50.1 75. 6
abort
73-1 12-1-66 Pad abort with 11 785 Yes 8.9 46.4 73.08 8. 5 58.8 74.8 7.9 68.4 76.8
short drogue
interval
__
73-3 12-20-66 Medium-altitude 1 1 785 Yes 19.2 46.7 49.7 18.7 53.251.5 4.4 49.7 101.8
abort with
extended
drogue
interval
73-5 1-17-67 Normal entry 1 1 785 Yes 25.0 70.1 29. 5 24.4 76.7 31.3 10.6 51. 6 80.4
'Dynamic pressure.
A second basic difference between the Blocks I and I1 test programs was an
attempt in Block 11 to eliminate the off-limit and overtest conditions prevalent in the
Block I series. Two of the Block I tests had to be repeated, and other t e s t s were
difficult to rationalize as fully valid because of their severity. In the Block I effort,
long free-fall periods, used to obtain high dynamic pressures, often resulted in a very
unstable vehicle and higher- than-desired dynamic pressures. In the Block 11 effort,
smaller programer parachutes were used; they remained attached to the vehicle and
remained operative until attaining the desired test conditions and attaining the initiation
of the recovery sequence. This technique permitted control over the attitude of the
vehicle and over body r a t e s to the point where the ELS functions began. This technique
resulted in near-nominal and sometimes below-nominal conditions for certain flight
modes; however, the t e s t s were far more representative of spacecraft conditions than
were many of the Block I tests.
B- 10
The Block II qualification drop tests were also characterized by two added opera-
tional simulations not demonstrated i n Block I. The first simulation involved an early
manual main-parachute deployment whereby the crew elects to override the automatic
sequence. This operation results i n a short drogue-parachute interval. The second
simulation concerned the main-parachute inhibit mode whereby the crew elects to delay
automatic main-parachute deployment and to extend the drogue interval. This tech-
nique could be used to avoid drifting back to a land landing in the event of a near-pad
abort.
One aspect of the Block II qualification drop tests, which might be considered as
being off -limit, concerned test-vehicle recovery weight. Although the specifications
still reflected a maximum vehicle weight of 11 000 pounds at the time of the tests, the
projected weight for Block II spacecraft had risen to 11 785 pounds. The three final
tests were conducted with the vehicle at this increased weight.
The results obtained from the Block I1 qualification drop tests demonstrated the
capability of the Block I1 ELS to land the CM safely under the conditions stipulated in
the specifications.
v
__
1967 1968
Series
~
B-11
The 80 and 81 s e r i e s were quite similar in that both were aimed at developing
I
the two-stage main-parachute reefing system. The 80 s e r i e s was single main-
parachute tests; on the other hand, the 81 s e r i e s used clusters of two main parachutes.
In the 80 series, Seven single parachute tests were conducted with the primary objec-
tives being to confirm the effective canopy drag areas and load estimates, to determine
f i l l rates, to establish the time interval for the reefed stages, and to determine the
second-stage reefing-line load. Four 81 series cluster tests were conducted to evalu- I
ate staged reefing with a cluster of two parachutes, representing the design condition,
to establish reefed load sharing, to determine quantitatively the effect of nonsynchro-
nous deployment and disreef, and to obtain further verification of selected reefing I
parameters.
Using single main parachutes and the ICTV, the 82 s e r i e s were verification tests
of canopy strength. This s e r i e s consisted of four tests conducted to demonstrate the
ultimate load-carrying capability of the main parachute in the first-stage reefed con-
dition, the second-stage reefed condition, and the full-open condition. I
The 83 and 84 s e r i e s were similar because both were combined drogue and clus-
tered main-parachute tests using the PTV. The 83 s e r i e s was planned primarily as
developmental type tests to establish drogue reefing parameters and t o obtain additional
performance data on main-parachute clusters. The planned 84 series w a s to have been
conducted with final spacecraft-configuration hardware and w a s to have included drogue-
parachute ultimate-strength verification tests. Because of delays in the test schedule,
the unavailability of spacecraft-configuration hardware, and'the close similarity
between the objectives of the 83 and 84 s e r i e s , only one of the 83-series tests was
actually conducted to demonstrate the effect of a missed second-stage reefing in one of
a cluster of three main parachutes. The test results supported an analysis showing
that the total axial load generated by a single main parachute (in a three-parachute
cluster) would not exceed the structural capability of the parachute if i t prematurely
disreefed from, o r bypassed, the second-stage reefing.
During the qualification tests, each tested component was identical to the space-
craft production design with one minor exception. During the three final tests, strain-
gage link assemblies were incorporated in the main-parachute r i s e r to obtain parachute
load data under simulated operational conditions of the spacecraft.
B- 12
The Increased Capability Block 11 qualification program consisted of the following
tests (table B-IV).
1. Two pad-abort simulations were conducted; one test demonstrated the by-
passed drogue condition. One test included jettison of the forward heat shield.
3. The third test conditions concerned three normal- entry simulations, all
incorporating jettison of the forward heat shield. Two of these tests demonstrated
single-drogue parachute conditions; one of these tests demonstrated the system with
a 1 3 500-pound vehicle recovery weight. The apex-forward attitude of the vehicle,
required to achieve the desired test conditions, prevented the use of the forward-heat-
shield system during three of the tests.
The drop tests made during the qualification series were dispersed as widely as
possible over the potential operational envelopes for the drogue and main parachutes
(figs. B-6(a) and B-6(b)). Included in figure B-6 a r e the test conditions at which the
drogue and main parachutes have been demonstrated during the total system-level (BP
vehicle) tests. Limitations imposed by the drop aircraft prevented drogue-parachute
deployments at the higher altitudes; however, the high-dynamic-pressure conditions
were well demonstrated. For each of the qualification tests, the conditions obtained
were very close to desired values, and no discrepancies of sufficient magnitude were
encountered to invalidate the test o r to prevent fulfilling the test objectives.
On the basis of the performance of the parachute system during each of the
qualification drop tests, and on successful completion of laboratory qualification at
the component and lower-assembly level, the Increased Capability Block II ELS was
verified for use on manned Apollo spacecraft.
1
drogue)
1 85-4 6-li-68 Norniaientry
(one d r o m e ) I
25,fi I
I
19.6 29.1
I
24.9
I
92.5
I
30.9
I
4.3 65.1 96. 5
85-7 7-3-68 Highaltitude 1 3 OOO No 28.9 42. 0 20.1 22. 5 149. 5 35.4 4.4 63.3 91.6
abort (one
drogue and
two mains)
aDwarnie p r e s s u r e .
B-13
m
40000
'0 o BP test points
/ Normal entry
r 30000- ,' " and hiah-
u
.-c
20000-
High-altitude abort
10000 -
0
4 .1 .2 .3 .4
,1 .2 .3 .4 .5 .6 .7
Mach number Mach number
B-14
MSC-05805
APPENDIX c
APOLLO 15 MISSION
MAIN PARACHUTE F A I L m
PREPARED BY
APPROVED BY
James A. McDivitt
Colonel, USAF
Apollo Spacecraft Program
c-1
M A I N PARACHUTE FAILURE
c-3
h
Figtire 1.- Spacecraft descetiding with one main parachute failed
c-4
Main parachute
cznopy
link
ser
Fabric riser
7
Steel cable riser
Fabric riser
c-5
l i n e s which are a t t a c h e d by s i x s t e e l connector l i n k s t o s i x i n d i v i d u a l
l e g s of a f a b r i c r i s e r . The s i x l e g s of t h e f a b r i c r i s e r coverge i n t o
a s i n g l e l e g which connects t o t h e end of a s t e e l c a b l e r i s e r . The t h r e e
s t e e l cable r i s e r s of t h e p a r a c h u t e system coverge and a t t a c h t o t h e com-
mand module through t h e p a r a c h u t e attachment and disconnect assembly.
DISCUSS1ON
A d i s c u s s i o n of t h e a n a l y s i s , t e s t s , c o n c l u s i o n s , and c o r r e c t i v e
a c t i o n s a r e contained i n t h i s r e p o r t . A l l times shown i n t h i s r e p o r t
a r e elapsed time from range z e r o . Range zero i s t h e n e a r e s t i n t e g r a l
second p r i o r t o l i f t - o f f .
FLIGHT DATA
This r e s u l t a n t v e c t o r l o c a t e s t h e f a i l e d p a r a c h u t e as shown i n f i g -
ure 4 . The computed f o r c e v e c t o r w a s s u b s t a n t i a t e d by body-mounted r a t e
gyro data.
PHOTOGRAPHIC DATA
C-6
clouds ZY5:0Y:12 purge start 295:09:21.8
Last frame on onboard
camera film 295:08:43
First visual Reaction control system
First sighting of Second colored cloud
sighting depletion firing start Spacecraft f ~ b s c ~ by
~ r d Falling object signted below
295:08:05 295:08:22 damage 295:01):15 295:w:za
clouds 2Y5:08:47 the spacecraft 29545%
7 7 I
t5.0
Pitch rate,
deglsec
-2.5
t2.5
t5.0
Roll rat.?,
dzylsec ’”:
-2.5
350
Z-axis
acceler ation,
1 -5.0
Reaction :%
control 9
Reaction 305 -1
system
helium
manifold 2 260
pressuree 245
control
system
helium 275
40
E
-2
Y-axis
t2 r
manifold 1 260 acceleration, 0
psia Reaction
230 pressure, 245 9
psia bus A voltage, control -1
215
200 system 4000
215 t tank 2 Reaction
200 helium control
’i pressure,
psia lo00
system
tank 1
helium
pressure, lo00
psia -2
c-7
i
c-9
A
a. Apparently t h r e e of t h e s i x legs of t h e f a b r i c riser were t a k i n g
t h e load.
CREW OBSERVATIONS
c-11
rc
Y
4
Lc
0 0
m se
r o
L
m aN
.-e
24 0
0 d
Y
4
. . -
0 0
.E 0
m e Qa
c-12
RECOVERED PARACHUTE INSPECTION
c-13
Failure
point
C-14
C-15
c . The forward h e a t s h i e l d mortar had f i r e d and t h e ramp had i t s
normal s c r a t c h e s . One pyrotechnic connector w a s b e n t , probably as a r e -
sult of ground handling.
FAILURl3 ASSESSMENT
The i n v e s t i g a t i o n w a s e s s e n t i a l l y d i v i d e d i n t o t h r e e a r e a s which
were l i k e l y s u s p e c t s as t o t h e cause of t h e parachute f a i l u r e .
C-16
Forward Heat Shield
T r a j e c t o r y a n a l y s i s . - A t r a j e c t o r y a n a l y s i s was performed u s i n g s i m -
u l a t i o n s t o determine i f t h e forward h e a t s h i e l d could have c o n t a c t e d t h e
main p a r a c h u t e s . The simulations were based on t h e point-mass equations
of motion, which used t h e known mass and aerodynamic c h a r a c t e r i s t i c s of
t h e forward h e a t s h i e l d and s p a c e c r a f t parachute systems and t h e measured
downrange and crossrange winds.
Photographic a n a l y s i s . - A c l o s e examination of t h e t e l e v i s i o n r e c o r d
of s p a c e c r a f t descent on t h e main parachutes e s t a b l i s h e s t h a t t h e forward
h e a t s h i e l d w a s below t h e s p a c e c r a f t at t h e time of t h e f a i l u r e . S p e c i f -
i c a l l y , t h e forward h e a t s h i e l d i s seen below t h e s p a c e c r a f t i n frame 588
( f i g . 9 ) a t 295:09:11:3, approximately 2 seconds b e f o r e t h e anomaly occur-
red. By c o r r e l a t i o n w i t h frame 775, which shows t h e parachute and forward
h e a t s h i e l d i n t h e same frame at 295:09:17.5, and by d i r e c t measurement
of t h e s e p a r a t i o n d i s t a n c e between t h e two o b j e c t s and measurement of t h e
known p a r a c h u t e dimensions, t h e v e r t i c a l s e p a r a t i o n d i s t a n c e s between t h e
forward h e a t s h i e l d and t h e s p a c e c r a f t were 580 f e e t f o r frame 588 and
1020 feet f o r frame 775.
C-17
Television frame 588
295:09:11.3 elapsed time
Direct measurement
AH (1020 ft.) = 580 f t 4.04 inches
A H = 580 f t
0 1
C-18
0 0 0 0 0 0 0 0
0 0 0 0 0 0 0 0
U 0 9 N 03 d 0 9
9 9 In m d U d m
ij 'apnwiv
c-19
Assessment o f p r o b a b i l i t y of f o r v a r d h e a t s h i e l d c o n t a c t i n g s p a s -
c r a f t . - An assessment of t h e p r o b a b i l i t y of t h e forward h e a t s h i e l d con-
t a c t i n g t h e s p a c e c r a f t w a s made t o determine t h e hazard a s s o c i a t e d w i t h
c o n t a c t . Actual wind d a t a i n t h e form of frequency of occurrence of
winds as a f u n c t i o n of a l t i t u d e , wind v e l o c i t y , and d i r e c t i o n were used
as a b a s i s f o r t h e s t u d y . Wind d a t h were a p p l i e d t o nominal t r a j e c t o r i e s
of t h e s p a c e c r a f t and forward h e a t s h i e l d i n a p l a n a r ( 2 dimensional)
a n a l y s i s which y i e l d e d t h e frequency of occurrence of s p e c i f i c v a l u e s Of
range s e p a r a t i o n between t h e two bodies a t i n t e r c e p t a l t i t u d e . Range
s e p a r a t i o n values of l e s s t h a n 100 f e e t between t h e two v e h i c l e s were
considered c o n t a c t . The cumulative p r o b a b i l i t y of c o n t a c t i s 0.093 per-
c e n t . This a n a l y s i s considered no t r a j e c t o r y d i s p e r s i o n s . Subsequent
refinement of t h e p l a n a r a n a l y s i s t o i n c l u d e e f f e c t s of l a t e r a l d i s p e r -
s i o n (due t o t h e moderate l i f t o f t h e forward h e a t s h i e l d system and t h e
s p a c e c r a f t on t h e drogue p a r a c h u t e ) provided a method which i s much l e s s
sensitive t o variation i n i n i t i a l conditions, principally i n f l i g h t path
a n g l e . The r e f i n e d a n a l y s i s a l s o y i e l d s a c o n t a c t p r o b a b i l i t y of about
0.1 percent.
c-20
TABLE I.- COMMAND MODULE/FORWARD HEAT SHIELD TRAJECTORY PARAMETERS
I n i t i a l Conditions
Forward h e a t s h i e l d j e t t i s o n
A l t i t u d e , ft .......... 23 300
F l i g h t path angle, deg ...... -73 1
Dynamic p r e s s u r e , l b / f t
2
..... 124
Spacecraft weight, lb ....... 12 810
Forward Heat S h i e l d
............
Weight, l b
Drag a r e a , .
CDS ft
2
........
L i f t c o e f f i c i e n t , CL .......
310
27.75
0
Spacecraft
A l t i t u d e , ft , , ., ,
, , 6 415
.
Time from forward h e a t s h i e l d
j e t t i s o n sec . . . . . . . . . .
No-wind range s e p a r a t i o n , f t . , . .
135.2
a
-755
a
Spacecraft downrange of forward h e a t s h i e l d .
c- 21
4 Forward heat shield leading edge
70 ft/sec
c-22
I lcan izing
C-23
These t e s t s showed t h a t t h e forward h e a t s h i e l d c o n t a c t i n g <ne para-
chute could damage some of t h e suspension l i n e s , but would probably not
cause a l o s s of r i s e r l e g s .
C-24
acceptable damage should t h e heat s h i e l d contact the s p a c e c r a f t and i t s
1 parachute system, no c o r r e c t i v e a c t i o n i s r e q u i r e d .
C-25
and none were observed. This unnotched s t u d w a s t h e n remounted i n a l i n k
assembly, t o r q u e d t o 120 i n - l b , which i s t w i c e s p e c i f i c a t i o n l e v e l , and
p l a c e d i n s e a water f o r 24 h o u r s . The l i n k s and s t u d s were t h e n a i r d r i e d ,
disassembled, and examined f o r c r a c k s . No cracks were found.
C-26
i 3. For t h e f a i l e d s t u d s , t h e flaws probably o c c u r r e d a f t e r t h e
plating operation.
C-27
i n t o the wind. When t h e canopy w a s f u l l y i n f l a t e d and s t a b l e , s e l e c t e d
r i s e r s were p y r o t e c h n i c a l l y severed. I n d i v i d u a l riser l e g l o a d s , t o t a l
r i s e r l o a d , and photographic documentation were o b t a i n e d .
C-28
4x lo3 .
0 -
-0.4 0 0.4 0.8 1.2 1.6 2.0
Time, sec
C-29
c. Burning f u e l can cause damage t o t h e r i s e r s , suspension l i n e s ,
o r parachute canopy.
C-30
P o s t f l i g h t t e s t i n g of t h e command module r e a c t i o n c o n t r o l system
showed it t o b e i n normal working o r d e r . T e s t i n g i n c l u d e d leak checks
of t h e p r o p e l l a n t t a n k b l a d d e r s , engine valve leak t e s t s , engine valve
s i g n a t u r e t r a c e s t o v e r i f y p r o p e r opening c h a r a c t e r i s t i c s , and e l e c t r o n i c
t e s t s t o v e r i f y t h e e l e c t r i c a l w i r i n g and t e r m i n a l b o a r d connections.
CON CLUSIONS
The a n a l y s i s of t h e d a t a and r e s u l t s of t h e s p e c i a l t e s t s l e a d t o t h e
following conclusions:
C-31
a . The most probable cause of t h e anomaly was -the b u r n i n g o f r a w
f u e l (monomethyl hydrazine) b e i n g e x p e l l e d d u r i n g t h e l a t t e r p o r t i o n of
t h e d e p l e t i o n f i r i n g and t h i s r e s u l t e d i n exceeding t h e p a r a c h u t e - r i s e r
and suspension-line temperature l i m i t s .
CORRECTIVE ACTION