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15 March 1972
STUDY OFSOLID ROCKET MOTORS
FOR A SPACE SHUTTLE BOOSTER
FINAL REPORT
UTe 4205-72-7e~-1"3(;,~
(NASA-CR-123623) STUDY OF SOLID ROCKET50TORS FOR A SPACE SHOTTLE BOOSTER. VOLOME1: EXECUTIVE SUMMARY Final Report (UnitedTechnology Center) 15 Mar. 1972 41 P21H G3/28
VOLUME I
EXECUTIVE SUMMARY
N72-22784 ]
15303
https://ntrs.nasa.gov/search.jsp?R=19720015134 2019-07-28T17:47:23+00:00Z
UTC 4205-72-7
STUDY OFSOLID ROCKET MOTORS
FOR A SPACE SHUTTLE BOOSTER
FINAL REPORT
VOLUME IEXECUTIVE SUMMARY
15 March 1972'
Prepared UnderContract NASS":'28431
Data Requirement No. MA-02Data Procurement Document No. 314
for
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION, GEORGE C. MARSHALL SPACE FLIGHT CENTER
MARSHALL SPACE FLIGHT CENTER, ALABAMA
by
United Technology Center
UDIVISION OF UNITED~R~FTCORPORATION
SUNNYVALE, CALIFORNIA
'{'l!l17'crnlNG PAGE BLANK NOT FILMED
ABSTRACT
The study of solid rocket motors for a space shuttle booster by United
Technology Center was initiated on 19 January 1972 under contract to Marshall
Space Flight Center of the National Aeronautics and Space Administration. The
objective was to investigate and define solid rocket motor booster stage designs;
development, production and launch operations programs; and creditable, under
standable costs of selected booster baseline configurations in order to furnish
the National Aeronautics and Space Administration with data necessary to make
decisions regarding the booster for the Space Shuttle Program.
Solid rocket motor booster stage studies were directed toward defining the
basic series burn and parallel burn boosters for both the 120- and l56-in.
diameter solid rocket motors required to accommodate both the 15- by 60-ft and
14- by 45-ft payload bays. Basic vehicle characteristics were amplified by the
orbiter phase B contractor teams to present a broad array of solid rocket motor
stage requirements. The cost and program definition tasks were extended to
reflect the requirements of the basic National Aeronautics and Space Administra
tion mission model of 445 flights plus 3 mission model rates of 40, 20 and 10
flights per year.
During the 8 weeks of the study, documented in this report, UTC elected
to define four typical baseline boosters to reflect orbiter contractor team
requirements, which were the basis for program and cost definition of the total
solid rocket motor booster stage costs. Definition of basic solid rocket motor
cost data was pursued as necessary to fulfill the National Aeronautics and Space
Administration cost estimates requirement. In addition to activities associated
with solid rocket motor booster stage system definition,. the technical consider
ations pertaining to ecology, abort, base heating and the concept of recovery
and reuse of solid rocket motor components were studied. A time-phased program
plan was defined for design, development, test and evaluation, production, and
launch operations to provide a complete basis for cost estimation.
iii
Detailed management, technical, and cost comparisons of the four baseline
solid rocket motor configurations led to the following conclusion:
A. The solid boosters offer maximum confidence for a reliable interim
booster, and all projections were based on creditable data offering
the minimum technical and financial risks.
B. A l56-in.-diameter solid rocket motor booster stage should be
developed for more than one-half the full mission model, and a
l20-in.-diameter solid rocket motor booster stage should be developed
for less than one-half the full mission model.
c. A solid rocket motor booster' stage provides the Space Shuttle Program
with a substantial management reserve in respect to a minimum invest
ment for a design, development, test, and evaluation program and
maximum confidence in satisfying budgetary commitments.
\
Section
1.0
2.0
3.0
4.0
5.0
6.0
7.0
8.0
CONTENTS
INTRODUCTION
PROGRAM OBJECTIVES
GROUND RULES
SRM BOOSTER STAGE DEFINITION4.1 Baseline Vehicle Descriptions
4.1.1 Payload Bay Orbiter (15 by 60 ft)4.1.2 Payload Bay Orbiter (14 by 45 ft)
4.2 Technical Considerations
SRM BOOSTER STAGE PROGRAM DATA5.1 Design, Development, Test, and Evaluation5.2 Production5.3 Operations
RECOVERY AND REUSE
ESTIMATION OF COSTS
CONCLUSIONS AND RECOMMENDATIONS
v
Page
1
1
1
13399
16161620
22
26
34
Figure
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
Table
4-1
ILLUSTRATIONS
SRM Booster Study Variables
SRM Characteristics
Baseline Launch Vehicle Configurations
Stage Characteristics
Calculated Maximum RC1 in Space ShuttleBooster Ground Cloud
Material Released into Stratosphere
Projected Maximum Sound Pressure LevelsInternal to Launch Vehicle
Design and Development
DDT&E Cost Breakdown
Production Plan
SRM Booster Stage Launch Site Functions
Flight Operation's Manpower Production Flights
Recovery of Solids
Program Savings Based on Recovery and Reuse
SRM Program Schedule
Program Cost Summary, SRM Booster Stage
Annual Funding Requirement
SRM Booster Stage Program Cost Summary
Typical SRM Hardware Cost Percentages
SRM Booster Stage Recurring Costs/Launch
Average Cost Per SRM
TABLE
Stage Characteristics
vi
Page
4
5
6
7
12
14
15
17
18
19
21
23
24
25
27
28
29
30
31
32
33
Page
10
AP
BLOW
CDR
DDT&E
EMI
GLOW
KSC
NASA
NAS/NRC
OLOW
PBAN
PDR
PFRT
SRM
SST
TT
UAC
UTC
ABBREVIATIONS
ammonium perchlorate
booster liftoff weight
critical design review
design, development, test, and evaluation
electromagnetic interference
gross liftoff weight
Kennedy Space Center
National Aeronautics and Space Administration
National Academy of Science/National Research Council
orbiter liftoff weight
polybutadiene-acrylic acid-acrylonitrile
preliminary design review
preliminary flight rating tests
solid rocket motor
supersonic transport
thrust termination
United Aircraft Corporation
United Technology Center
vii
viii
"
1.0 INTRODUCTION
The study of SRM booster stages was conducted to furnish NASA with data
necessary to make decisions regarding the booster for the space shuttle program.
The study was divided into manageable tasks corresponding to the study'
program objectives to define a SRM design; development, production, and launch
support programs; and creditable, understandable costs of chosen booster base
line definitions. As design variables affecting the baseline SRM stage were
identified by Phase B contractors, configuration impact upon nonrecurring and
recurring costs were assessed. Study effort was applied to define recovery
and reuse of SRM components for the purpose of providing comparative data.
2.0 P~OGRAM OBJECTIVES
For the purpose of this study, program objectives were to be met by satis
fying the following:
A. Definition of SRM designs which satisfy the performance and configura-I
tion requirements of the various vehicle/booster concepts. '
B. Definition of the development, production, and launch support programs
required to provide SRM booster stages at rates of 60, 40, 20 and 10
launches per year in a man-rated system.
C. Estimation of costs, including ground rules and assumptions for basis,
for the defined SRM booster stages. These costs identify all hard
ware systems, design, development and test efforts, production efforts,
and launch support operations.
3'.0 GROUND RULES
The study was designed to provide hardware systems and program definition
necessary to establish the estimated cost of providing SRM booster stages stated
1
in const&nt calendar year 1970 dollars. Costs were developed employing
representative elements of costs (i.e., labor, materials, subcontract, 'general
and administrative, and miscellaneous) without fee. Cost estimates were pre
pared for total program funding and program time-phased funding requirements
~n consideration of the four UTC selected SRM stage configurations, using the
NAsA defined mission models and reflecting delta funding requirements of a
basic SRM versus a complete SRM booster stage.
Actual co~ts of previous and current SRM development, production, and
launch operations programs were utilized to the fullest extent possible.
Costs were developed and portrayed in consideration of other element~ of
NAsA direction and SRM booster stage baseline schedule, which depict~ signifi
cant milestones required to satisfy program requirements.
4.0 SRM BOOSTER STAGE SYS~EM DEFINITION
System requirements for SRM space shuttle boosters have been defined by
the various Phase B orbiter contractor teams. UTC began its current SRM study
by requesting basic vehicle data from these contractor teams. Booster siz~s
were then selected to carry out the point design and booster cost analyses.
Technical integration activities were continued with the Phase B contractors
to obtain more detailed SRM booster requirements, supply technical data to the
contractors, complete the SRM booster-orbiter interface definition, and maintain
liaison with the contractors.
Although contractor r~sponse to this request was varied, sufficient data
were obtained to establish the basic sizing of the SRM boosters for the parallel
burn and series burn modes. Most d&t& received were based upon the 15- by 60-ft
payload bay orbiter. Accordingly, the UTC study was oriented to define cost and
design data for this vehicle. Data were obtained from a single contractor to
define 14- by 45-ft payload bay orbiter booster sizing requirements.
2
The SRM booster stage studies were directed toward defining the basic
series burn and parallel burn boosters for both the 120- and 156-in.-diameter
SRMs. Cost data and design definition of the SRM boosters were directed toward
total stage costs, although the basic SRM data were pursued as necessary to
fulfill NASA cost requirements and to highlight apparent cost discrepancies
between UTC large motor expereence and cost projections by orbiter/booster
contractor teams. The cost and program definition tasks were extended to
reflect the requirements of the four NASA mission models. The study variables
are shown in figure 1.
All basic vehicle characteristics were amplified by the orbiter contractor
teams to present a broad array of SRM stage requirements. UTC elected to define
typical baseline boosters to reflect these contractor requi,rements.
The UTC baseline design uses the TECHROLL@ seal movable nozzle on the basis
of a 10% cost reduction compared to a liquid injection system and a reduction
in actuation torque requirements and movement of the nozzle seal pivot point
co~pa~ed to the flex-seal. Thrust termination devices have been included
on',the basis of prior acceptance and qualification for use in abort systems on
manned flight programs. Forward thrust loading has been selected on the basis. " -
of orbiter contractor preference and a relatively small weight and cost penalty
within the SRM stage.
The UA 1207 seven-segment SRM is a direct derivative of the current Air
Force Titan III five-segment SRM. The components of the 156-in.-diameter SRM
would be scaled up from the 120-in.-diameter SRM currently in flight use. SRM
characteristics involved in these applications and their derivation are presented
in figure 2.
4.1 BASELINE VEHICLE DESCRIPTIONS
4.1.1 Payload Bay Orbiter (15 by 60, ft)
Basic sizing and configuration of the UTC selected baseline SRM shuttle
boosters are shown in figure 3. Series burn and parallel burn configurations
for the 120- and 156-in.-diameter SRMs are shown. Characteristics .of the various
boosters are listed in figure 4.
3
ITEM
SELE
CTED
BAS
EllN
EAL
TERN
ATES
ORB
ITER
PAYL
OAD
BAY
15X
60FT
14X
45FT
LAUN
GH
CO
NFI
GU
RAT
ION
PARA
LLEL
BURN
SERI
ESBU
RN
SRM
DIA
MET
ER12
0IN
.15
.6IN
.
CO
NFI
GU
RAT
ION
BASI
CSR
MST
AGE
HRUS
TVE
CTOR
CONT
ROL
MO
VABL
ENO
ZZLE
NONE
;TE
CHRO
LL8
SEAL
LIQ
UID
INJE
CTIO
N
HRUS
TTE
RM
lNAT
ION
WIT
HW
ITHO
UT
HRUS
TLO
AD.I
NGFO
RWAR
DAF
T
ISSI
ON
MO
DEL
PEAK
RATE
/YEA
R10
,20
,40
,O
R60
T TT M
Fig
ure
1.
SRM
Bo
os'
ter
Stu
dy
Vari
ab
les
UT
C-V
-014
57
ITEM
UA
1205
FOR
'U
A12
071
20
-IN
.SR
M1
56
-IN
.SR
MTI
TAN
IIIC
/DTE
STED
FOR
SHU
TTLE
FOR
SHU
TTLE
PRO
PELL
AN
T0
00
0IN
SUL
AT
ION
0~
~
CA
SESE
GM
EN
TS
00
(1)
o(1
)~
FOR
WA
RD
CLO
SUR
Eo
(2)
~(3
)~
(3)
~
IGN
ITE
RA
ND
S/A
DEV
ICE
00
0~
..
~TH
RUST
TE
RM
INA
TIO
NSY
STEM
0(2
)0
0A
FTC
LOSU
RE
00
0~
NO
ZZ
LE
0~
~~
THR
UST
VEC
TOR
CO
NT
RO
L0
0•
•O
NB
OA
RD
POW
ERSY
STEM
0~
(4)
~(4
)~
FLIG
HT
INST
RU
ME
NT
AT
ION
00
00
SUPP
OR
TL
ON
GE
RO
NS
0~
~~
AFT
SUPP
OR
T0
0~
~
FOR
WA
RD
SUPP
OR
T0
00
0
(l)T
WO
AD
DIT
ION
AL
SEG
ME
NT
S
(2)
THR
UST
TE
RM
INA
TIO
ND
ELET
EDFR
OM
PRO
DU
CT
ION
MO
DEL
S
(3)
EXTE
ND
ED
(4)
RE
DU
ND
AN
CY
AD
DED
oSA
ME
AS
UA
1205
~SA
ME
DES
IGN
AS
UA
1205
Fig
ure
2.
SRM
Ch
ara
cte
rist
±cs
•D
IFFE
REN
TFR
OM
UA
1205
UT
C-V
-015
41
ORB
ITER
PAYL
OAD
BAY
__
_15
X60
FT
MA
XIM
UM
ACCE
LERA
TIO
N
------<
3G
MA
XIM
UM
DYN
AMIC
PRES
SURE
__
__
---:....-
_<
650
PSF
SER
IES
(S6-
12O
)
SER
IES
(S3-
156)
PARA
LLEL
(P4-
-120
)
PARA
LLEL
(P2-
156)
Fig
ure
3.
Bas
elin
eL
aunc
hV
ehic
leC
on
fig
ura
tio
ns
UT
C-V
-015
40
S3-1
56S6
-12O
P2-1
56P4
-120
LAUN
CHC
ON
FIG
UR
ATIO
NSE
RIES
SERI
ESPA
RALL
EL_
_PA
RAL
LEL
NUM
BER
OFSR
Ms
36
24
PRO
PELL
ANT
WEI
GH
T,LB
X10
-6
SRM
1.07
50.
592
1.24
50.
592
BOO
STER
3.22
53.
552
2.48
92.
368
THRU
ST,
LBX
10-6
SRM
(SEA
LEVE
L)2.
178
114
02.
549
1.40
0BO
OST
ER6.
534
6.84
05.
098
5.60
0
ACTI
ONTI
ME,
SEC
141
135
137
140
STAG
EM
ASS
FRAC
TIO
N0.
869
0.86
00.
880
0.86
4
CONT
ROL
REQU
IREM
ENTS
DEFL
ECT
ION
,DE
GREE
±12
±12
±6±6
RATE
,DE
GREE
/SEC
55
55
PRO
PELL
ANT
PBAN
~
ACC
ELER
ATIO
N,
GLI
FTOF
F1.
25TO
1.3
..FL
IGHT
LESS
THAN
3..
MA
XIM
UM
DYN
AMIC
PRES
SURE
,P
SF
_LE
SSTH
AN65
0..
Fig
ure
4.
Sta
ge
Ch
ara
cte
rist
ics
UT
C-V
-014
59
The series burn configuration of three 156-in.-diameter SRMs (S3-156) is
shown in more deta~l in figure 2-3, volume II. Three 156-in.-diameter'SRMs
(each SRM contains 1,080,000 1b of propellant) are grouped triangularly in a
tank-end load configuration. Three identical SRM assemblies plus an interstage
assembly, or HO tank adapter, make up the booster. Vehicle prelaunch support
is provided by four ground support fittings at the base of each SRM. Two
diametrically opposed TT ports are provided at the forward end of each SRM. The
physical arrangement of the SRMs and TT ports has been selected to provide the
maximum miss distance of TT debris to the orbiter.
The series burn configuration of six 120-in.-diameter SRMs (S6-120) is
shown in figure 2-4, volume II. The six UA 1207 motors (each containing 592,000
1b of propellant) are arranged in a rectangular 2- by 3-ft tank-end load con
figuration. The booster is made up of six identical SRM assemblies plus an
interstage assembly as in the 156-in.-diameter case. TT is provided at the
forward end of each SRM with two ports arranged 900 apart on each motor.
The parallel burn configuration of two 156-in.-diameter SRMs (P2-156) is
shown in figure 2-5, volume II. Each of the two SRMs contains 1,250,000 1p of
propellant. Two identical SRM assemblies are strapped onto opposing sides of
the orbiter HO tank. Precise location of the SRMs is not critical to SRM design,
and the preferred 10cat~on may be selected on the basis of vehicle dynamics.
SRM thrust transmission to the HO tank and orbiter-HO tank ground support is
provided by a structural skirt at the forward end of the SRM. ,Total vehicle
ground support is provided by structural skirts at the aft end of the SRMs.
Two diametrically opposed TT port~ are provided at the forward end of each SRM.
The ports are located to provide a maximum miss distance to the HO tank and
orbiter while providing a laterally balanced thrust to allow abort orbit SRM
ejection.
The 120-in.-diameter SRM parallel burn booster is shown in figure 2-6,
volume II. The configuration of the four identical UA 1207 SRMs is similar to
that of the 156-in.-diameter motors. Use of the four SRMs establishes a broader
ground support base to react prelaunch wind and engine start transient overturn
ing loads.
8
4.1.2 Payload Bay Orbiter (14 by 45 ft)
Vehicle sizing data were obtained to allow basic definition of candidate
boosters for the 14- by 45-ft payload bay orbiter. Table I provides the basic
vehicle data and indicates how the 15- by 60-ft payload booster configurations
can be used to define a 14- by 45-ft payload booster. Design definition and
cost data were not prepared specifically for these configurations. However,
the basic 15- by 60-ft payload booster data can be applied to these boosters.
For design and cost purposes, the S2-156 series burn booster with two
156-in-diameter SRMs is identical to the P2-156 15- by 60-ft payload booster.
The SRMs are attached to the HO tank in a parallel fashion, but the engines
are designed to operate in a series burn mode. The ballistic design will be
varied from the P2-156 in precise throat size and propellant burning rate to
meet the S2-156 thrust requirements.
The S4-120 series burn configuration utilizes four identical VA 1208 SRMs
with an intertank structure similar to that of the S6-120 booster. This UA 1208
is similar to the S6-120 UA 1207 with differences in an additional segment, .~
shorter forward closure and reduced throat size and propellant burning rate to
produce the required S4-120 thrust.
The P2-156 14- by 45-ft payload booster is identical in concept· to the
P2-156 15- by 60-ft payload booster. The attach structure provisions are
identical, and the S3-156 model SRM is utilized. Nozzle throat size and pro
pellant burning are adjusted to provide the proper thrust characteristics.
The P3-120 SRM design is identical to that of the P4-120 with the exception
of adjustments to the nozzle throat size and propellant burning to provide
proper thrust tailoring.
4.2 TECHNICAL CONSIDERATIONS
Studies were conducted in such areas as ecology in respect to pollution
and acoustics, abort requirements and techniques, thrust neutralization, base
heating, and SRM booster stage integration. Based on currently defined vehicle
9
TABLE I
STAGE CHARACTERISTICS
S2-156 S4-120 P2-156 P3-120
1,198,000 1,198,000 1,650,000 1,650,000
2,820,556 2,847,876 2,446,248 2,045,113
4,018,556 3,934,876 3,096,248 3,695,113
Series burn Series Parallel Parallelparallel attach
2 4 2 3
......-------- PBAN-~-------..~
......------- 1. 25 to 1.3 --------
---------Less than 3.0-------l-
......----'-----Less than 650-------..-
0.620 1.080
2.480 2.160
1.267 1.800
5.068 3.600
126 130
0.870 0.883
1.220
3.660
130
0 0 867
0.592
1. 776
+6
.5
+6
5
+12
5
1.250
2.50
2.600
5.200
128
0.887
+6
5
Booster
Action Time
Stage Mass Fraction
Control Requirements
Deflection, degrees
Rate, degrees/sec
Propellant
Acceleration, g
Liftoff
Flight
Maximum Dynamic Pressure, 1b/ft2
Number of SRMs
Propellant Weight, 1b x 10-6
SRM
Launch Configuration
OLOW
BLOW
GLOW
Booster
Thrust, 1b x 10-6
SRM (sea level)
10
requirements and SRM booster stage configurations, no technical condition was
identified concerning the application of SRMs for a shuttle booster. More
precise definition of requirements and attendant interfaces will dictate further
analyses and engineering design. The subject of ecology is included in this
summary report because of the varied interests in the effects associated with
SRM use rates required in support of the NASA basic mission model.
From a study of the composition of the exhaust gas of a solid propellant
space booster stage in the atmosphere, it appears that most constituents are
present in quantities too small by several orders of magnitude to present any
problem. Because the solid propellant produces HCl, which is not present in
the exhaust of most liquid engines, only the behavior of the part of the exhaust
gas which can come back to earth is shown (figure 5). Following ignition, the
bases initially swirl about the launch platform, mixing with the surrounding
air and forming a ground cloud. After about 22 sec, the vehicle has risen so
high that the exhaust gases no longer reach back to the ground cloud. Because
the cloud is hot it will rise to a height of 1,330 ft (figures applicable to
the Titan III-C and documented by photographic observations are shown in paren
theses). The cloud then drifts downwind and grows by turbulent ·diffusion until
at a distance of 67 miles, under adverse meteorological conditions, its edge may
reach the ground again.
Based on analytical predictions for a series burn booster using six UA 1207
SRMs, the total weight of HCl emitted into the ground cloud is.139,000 lb, and
the HCl concentration at first is 965 ppm. By the time the cloud has risen to
its terminal altitude, the HCl concentration has dropped to 450 ppm; when the
cloud touches the ground it has decreased to 3.7 ppm. The Committee on Toxi
cology of the National Academy of Science/National Research Council has recently
set a Short Term Public Limit of 4 ppm for 10 min and a Public Emergency Limit
of 7 ppm for 10 min. Neither of these limits is reached under the adverse
conditions on which the calculations were based. It is of interest to note
that these limits represent values at which the impact may be no more than a
strong odor or, at the most, slight irritation of the mucous membranes, with
no adverse health effects.
11
67M
I(4
2M
I)
139,
000
(22,
400)
4PP
MFO
R10
MIN
7PP
MFO
R10
MI.N
----- --
----
----
---
I·1
420
FT
CA
LC
UL
AT
ION
SBA
SED
ON
WO
RST
-CA
SEA
NA
LY
SIS
SIX
UA
1207
SRM
s(S
ERIE
SB
UR
NC
ON
FIG
UR
AT
ION
)
PLUM
E~I\
1\ 'I(
~ "W
I.ND
.•,)
'-A
(900
FT)
HCI
CONC
ENTR
ATIO
N,PP
M96
5(6
11)
450
(280
)TO
TAL
WEI
GHT
OFHC
IIN
GRO
UND
CLO
UD,
LBN
AS/N
RC
SHO
RTTE
RMPU
BLIC
LIM
IT,
NAS
/NR
CPU
BLIC
EMER
GEN
CYLI
MIT
,
*NU
MBE
RSIN
PARE
NTHE
SES
ARE
FOR
TITA
NII
I
Fig
ure
5.
Cal
cula
ted
Max
imum
RC
Iin
Sp
ace
Sh
utt
leB
oo
ster
Gro
und
Clo
ud.
lIT
C-V
-01
47
1
Because the atmosphere becomes relatively thin in the stratosphere above
40,000 ft, it is of some concern whether the exhaust gas ejected by the space
shuttle boosters from this altitude until burnout at 135,000 ft will be large
enough to constitute a pollution problem. Because recently there have been
major discussions of this problem in connection with the commerica1 operation
of supersonic airplanes flying in the strato~phere~ the quantity of material. ,
ejected at the maximum launch rate of 60 per yearby a space shuttle booster
using six UA 1207 SRMs was compared to that emitted by an SST fleet of 400 air
planes. It is obvious that the quantities of water, CO2, CO, and NO ejected
by the space shuttle boosters are three or more orders of magnit.~~e smaller.
Because the SSTs do ,not emit HC1 this gas cannot be compared in the same way,
but its concentration, when spread throughout this band of the stratosphere,
was compared with the actual concentration of ozone in the same band. The
concentration of HC1 resulting from the space shuttle boosters would be one
million times less than the concentration of ozone. Therefore, HC1 would have
no effect on the spectral absorption characteristics of the atmosphere. Particu
late matter emitted by the space shuttle boosters would be less than that emitted
by the SST fleet, but only by a factor of 13. Ho~ever, the quantity of inter
stellar dust particles which pass through the atmosphere on their way to the
earth's surface is one thousand times greater than that deposited by the space
shuttle booster. Therefore, it is expected that pollution of the stratosphere
by the space shuttle booster'is not a serious problem (figure 6).
The Titan III-C launch vehicle uses two UA 1205SRMs. Detailed acoustic
measurements made on these vehicles allow a projection of noise levels to be
expected within the HO tank and orbiter vehicle. The maximum noise should occur
shortly after launch and diminish to a minimum about 15 sec after liftoff.
Aerodynamic noise then becomes dominant, with the peak level' of lS0 to 158 db
occurring at the time of maximum dynamic pressure (figure 7).
13
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5.0 SRM BOOSTER STAGE PROGRAM DATA
An important element of the study was definition of programs required for
design, development, test, and evaluation; production capabilities and processes;
and launch operations and logistic concepts. These programs were defined
employing the NASA-established ground rules, assumptions, and mission models,
which were the basis for the estimated booster costs furnished to NASA.
5.1 DESIGN, DEVELOPMENT, TEST, AND EVALUATION
DDT&E plans were derived from previous development experience with the
120- and l56-in.-diameter SRMs. Figure 8 portrays a representative design and
development plan and a schedule time-phased to support key shuttle system
milestones. The plan and schedule are applicable to both motors, except that
differences will occur in long lead-time procurement items, scope of full-scale
development testing, and variations in requirements for qualification testing.
Figure 9 summarizes the DDT&E funding requirements for the four baseline
configurations which encompass those elements of cost included in this nonrecur
ring cost category, such as inert motors for dynamic testing and development
flight test motors. For general comparative purposes, the net DDT&E funding
requirements are shown.
5.2 PRODUCTION
Production planning encompassed the support required for the recurring
operational phase and was defined in terms of all SRM booster ~tage configura
tions and mission models. Other variations also were introduced per the NASA
requirements and are reflected in the various cost presentations (figure 10).
Figure 10 summarizes the results of reviewing required production capacities
and establishing a plan to provide a production capability necessary to support
the Space Shuttle Program. This can be accomplished by expanding the existing
capabilities at the UTC Development Center, California, and by establishing an
additional production capability at the UA facility in Florida. These require
ments will be provided by use of corporate financing.
Stage structures and other inert hardware components also can be provided
by expansion of existing capabilities.
16
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Maximum use of the subcontractors and vendors currently qualified for the
Titan III Program has been considered. Production can be expanded as necessary
to support selected mission models.
With the exception of AP, production capacities exist to support the
annual peak requirements for propellant raw materials presently defined.
Current capacities for AP are in the range of 35 to 50 million pounds per
year. Because of the limited commercial use for this product and an estimated
production facility expansion cost of $15 to $25 million, a Government
industrial facility will be required to support this significant production
requirement.
Element
Aluminum powder
AP
Epoxy resin (DER 332)
PBAN Polymer
Other materials
5.3 OPERATIONS
Annual PeakRequirement, lb
36 million
150 million
7 million
25 million
As required
Operations planning incorporated previously discussed ground rules, assump
tions, and mission models in addition to assuming all flights were from KSC per
the NASA instructions. The concepts defined the transportation, logistics,
material handling and launch base activities necessary to successfully imple
ment a factory-through-launch plan for a man-rated SRM booster stage. This
concept and the ground systems required were derived from significant Titan III
experience gained from launch operations from both the East and West Coasts.
This ground system has been demonstrated completely and is capable of accommodat
ing the space shuttle booster system. Figure 11 identifies launch site functions
in the three major categories of SRM operations required for flight. Within
each category, discrete operations are identified which form the basis for
establishing facilities, ground support equipment, manpower, and associated
requirements.
20
RECE
IVE
AND
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ECT
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ITE
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Fig
ure
11
.SR
MB
oo
ster
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urC
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1478
The three functional categories relate to the following SRM activities:
(1) receipt, inspection, and subassembly required to initiate SRM boos~er stage
assembly on the launch platform, and recycle and turnaround operations assoc
iated with transportation and transportation ground support equipment; (2) assem
bly, checkout, and acceptance of the SRM booster stage,'integration with the
orbiter propellant tank and the orbiter, and integrated system testing required
prior to transfer to the launch area; and (3) launch pad final preparations,
prelaunch functional checkout operations, and launch support.
Launch site manpower requirements are a function of the booster configura
tion, mission model, and definition of the ground support system, as presented
in figure 12.
6.0 RECOVERY AND REUSE
Recovery and reuse of SRMs received considerable emphasis during this study.
Although SRM recovery was studied in some detail by UTC in the early 1960s for
application with the space shuttle system, a complete update and review was done.
The results of the current UTC recovery and reuse study are summarized in
figure 13.
The launch cost savings which accrue to the program by recycling 7, 14,
21, and 100 times represents a very significant savings. In each of the NASA
mission models, a savings of approximately 25% to 35% of the total SRM fabrica
tion cost can be realized by recycling. These savings (figure 14) represent
worst-case conditions for recovery and refurbishment and include attrition rates
of 16% to 25% for recovered hardware for 7 reuses and up to 42% for 100 reuses.
The cost analysis was based on recovering only major components of the SRM with
no redesign required, except for incorporation of the parachute recovery system
into the nose section of the SRM.
It is of importance to recognize that if the results of subsequent studies
on recovery of SRMs reveal that recovery and reuse of SRM components is not
feasible, the maximum financial risk will remain limited to the expendable SRM
costs as indicated.
22
Fig
ure
12
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61
7.0 ESTIMATION OF COSTS
Estimation was a significant objective of this study. Information from
previous and current development and production programs were used with the
SRM system definition and program planning from this study, to accumulate
creditable and understandable cost data. The cost data then were portrayed in
the formats and cost elements established by the NASA direction. In addition,
various cost data presentations were established to assist NASA in their deter
minations. Figure 15 presents the SRM program schedule used for cost extimation.
The SRM booster stage cost summary presented in figure 16 is the result of
UTC's cost estimations based on the defined baseline configurations and programs
and the variations due to mission models with and without recovery. A time
phased annual funding requirement for a mission model of 10 launches per year
is shown in figure 17.
Although this study involved SRM booster stage requirements for the
14- by 45-ft payload bay orbiter as a secondary effort, a program cost summary
for three candidate SRM booster configurations was prepared and is presented
in figure 18.
To provide further cost parameters for use by NASA, figures 19 and 20 are
presented. Figure 19 provides information useful in analyzing recovery savings,
production funding, and vendor recurring funding requirements. Figure 20
presents information reflecting recurring costs of the varous mission models
of the four baseline configurations and showing the delta cost impact of SRM
recovery and reuse.
Figure 21 presents information significant to the impact of- funding require
ments as a function of launch rate based on the four mission models examined.
These costs are in millions of dollars and assume no recovery or reuse.
26
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AL
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RATE
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CONF
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68
8.0 CONCLUSIONS AND RECOMMENDATIONS
the UTe study of SRMS for a space shuttle booster investigated and defined
four SRM booster stage designs; established concepts and plans for development,
production, and launch operations programs; and determined creditable, undet
standable costs for the selected baseline booster configurations.
The detailed management, technical, and cost comparison of these study
results leads to the conclusion that:
A. The solid boosters offer maximum confidence for a reliable interim
booster; all projections were based on creditable data offering
maximum technical and financial risk.
B. A l56-in.-diameter SRM booster stage should be developed for flights
in excess of one-half the full mission model, and a l20-in.-diameter
SRM booster stage should be developed for less than one-half of the
full mission model.
C. An SRM booster stage for the Space Shuttle Program requires a minimum
investment in a DnT&E program and yields maximum confidence in satis
fying known budgetary restraints.
It is recommended that the results of this study be accep~ed as the basis
for a decision to select the SRM booster stage for the space shuttle system.
It is further recommended that near-term design, development, test and
evaluation projects be implemented for SRM recovery and reuse technology, thrust
vector control development and test, and for advancement of the manufacturing
industries' state of the art.
34
