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,3-2§2-AMOO-75-509
20
FEBRUARY
1976
(NASA-CR-146805)
POSTEIIGHT
ANALYSIS
FOR
N76-21262
DELTA
PROGRAMM
ISSION
NO.
113:
COS-B
MISSION
(McDonnell-Douglas
Astronautics
Co.)
213 pHC $7.75
CSCL
22C fnclas
15212
41G3/18
POST
FLIGHT
ANALYSES
FOR
D
-LTA
PROGRAM
MISSION
NO.
113
-
COS-B3
MISSION
CONTRACT
NAS7-8321,
MMARNN
UL
RECEIVDS.
"AGrh
'2SpT
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A3-262-AMOC-MT5-509
Date:
FEB
2
0
1976
MEMORANDUM
Subject: POSTFLIGHT
ANALYSES FORDELTAPROGRAM
MISSION
NO.
113
-
COS-BMISSION
- CONTRACTNAS7-832*
To:
E.W.Bonnett,A3-900
Copies to:
C.H.
Baumann,
F.M.
Keller,
D.W.
Knebel,
J.R.Reider,
J.C.
Simmons,
D.W.
Tutwiler,
F.B.VanShoubrouek,
A3-200;
F.J.Maguire,A3-G83;
T.
B.
Rehder,
J.
L.
Schmidt,A31-822;
M.D.Steffey,
A41-770;D.
R.
Cummings,
A41-792;D.A.Maclean,
A41-822;
File
From:
C.A.Ordahl,
A3-262
1. Thismemorandum
has been preparedin
accordancewith
COM
15 ofthe
subject contract.
2. On
8August 1975,the
COS-Bspacecraftwaslaunched
successfully
from
the
Western
TestRange
(DeltaProgramMission
No.
113).
The
launchvehiclewas
a
three-stageExtended
LongTankDelta
DSV-3P-IIBvehicle,Serial
No.20018.
3. Postflightanalysesperformed
in
connection
with
DeltaProgram
Mission
No.113 (COS-B
Mission) arepresented
inthe attachments tothismemorandum
(Attachments1through10).
Theseattachments
consistofthe
following:
Attachment
Number Title P_e
1
Section
1.
System
Performance
- COS-BMission 1-1 through
1-26
2
Section 2. Propulsion
Systems
-
COS-BMission
2-1 through2-28
3 Section
3. GuidanceSystem
- COS-B
Mission
3-1
through
3-15
4 Section
4. Flight ControlSystem
- COS-BMission
4-1
through
4-29
5 Section
5. Electronics System
- COS-BMission
5-1
through
5-11
6 Section
6.
Mechanical Systems
- COS-B
Mission
6-1
through
6-10
D. 3360-1114;
EWO
54173; COM
15
rRACTUALDOCUMENT
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E.W.
Bonnett,
A3-900
-2-
A3-262-AMOO-M75-509
Attachment
Number
Title
Pages
7
Section
7. StructuralSystems
-
COS-B Mission
7-1 through
7-6
8 Section
8.
Reliability
8-1
through8-2
9
Definitions
of Performance
Parameters
[Tables
9-1
through
9-2
2-3 and2-4 ofAttachment2(Section
2)]
10
VehiclePerformance
Telemetry
Plots-
COS-B Mission
10-1through
10-73
C.A.Ordahl
Chief
Engineer
Delta
Programs
Engineering
Division
FMW:lsm
Attachments:
As Noted
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Attachment Ito:
A3-262-AMOO-M75-509
ATTACHMENT
1:
SECTION
1.
SYSTEM
PERFORMANCE-
COS-BMISSION
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Attachment1 to:
- A3-262-AMOO-M75-509
MISSION ANALYSIS
-
COS-B
MISSION
INTRODUCTION AND
SUMMARY
Delta
MLission
Number113,COS-B,was launchedfrom
Pad
SLC-2W of
the
Western
Test Range
(WTR)
ataflight
azimuth
of
196degrees from
true north
at 0147:59.595
Greenwhich
Mean
Time
(CMT)
on
August 8, 1975.
This section
provides a
discussion of the
mission analysis
aspects of
the
COS-B
spacecraft launch
and a
description of the
trajectory
flown by the Delta vehicle
fromliftoff
to
third-stage
burnout;
data
pertaining
to
theexperimental
second
stage
restart
are
included.
This section
alsopresents
acomparisonbetween
(1)
the
actual
trajectory
flownby
thevehicle,
(2)
the
guidednominal
trajectory
(Reference
1),
and
(3)he latest
predicted
or
Best Estimate
Trajectory with
launch-day
winds
and
atmosphere
(BET-with-winds).
The
actual trajectory
flown
by the
first and second
stage is
based on
Federal
Electric Corporation (FEC)
radar tracking
data
and NASA-provided
hardpoint
position
and velocity vectors
(Reference
(2). PCM
telemetry data
was
utilized
to
support the determination
of
trajectory
data at the
time
points specified in
the subsections
to
follow. The
following table compares
the
achieved orbit at
spacecraft
injection
(TECO) to
the
nominal
orbit given in the
Orbit
Accuracy Incentive
TWX (Reference 3).
Incentive
flight requirements
within
allowable
tolerances.
may be seen to
have been met;
that
is,
all fall
Parameter
Nominal
(Reference
3)
Achieved
Achieved
Minus
Nominal
Incentive
Tolerance
Apogee
Altitude
(Integrated)
(n.td.)* 53,992
54,433
+44211
+2115
Perigee
Altitude (n.mi.)*
188.i9
187.17
-1.02
-10.0
Inclination
(deg)
90.000
90.155 +0.155
+0.82
Table
1
summarizes the
orbit
parameters of
all
Delta
missions
to dateand,
where
applicable,
includes
thecorrespondingthree-sigma
deviations. Table
2
presents the guided
nominal, BET-with-winds,
and
actual sequence of events for
the
COS-Bmission.
VEHICLE
DESCRIPTION
The launch vehicle
used
for
the COS-B
mission consists of
a
DSV-3P-lA Extended
Long
Tank
Booster No. 602
(Serial
No.
20020)
powered by a Rocketdyne
RI-27
liquid
propellant engine
and
nine
strap-on Thiokol
TX-354-5
(Castor I)solid
propellant
rockets
with low-drag
nose
cones. The second
stage
is a DsV-3P-4B
(SerialNo.20023)
having
aTRWengine (light
quartznozzlewith
Expansion
Ratio
= 43:1) with
restart
capabilityandaDSV-3P-7A
fairing (SerialNo.
20023).
The
third stage
consists
of
a TE-364-3
engine
(Serial No. 00025).
Based
on
anearth
radius
of3442.62n.mi.
'-
1-I
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METEOROLOGI
CAL
CONDITIONS
Figures 1,
2, and
3 present
the launch
day
temperature,
pressure,
andwind
speedanddirection fromgroundlevelto
100,000
feet
measured
at
WTR
at
the
time
of launch. Figure 1 shows
the atmospheric temperature
was
hotter than
the
reference
temperature untilapproximately
42,000
feet,
colder
between
42,000
feet
and 70,000 feet and
hotter above
70,000 feet.
The
atmospheric
pressure (Figure
2) is generally higher
than the reference
atmosphere.
Figure 3
indicates
that
the
wind speed
was
significantly
lower
than
the 90
percent
IRIG
wind
reference until approximately
65,000
feet and
increasingly
higher above 65,000 feet. The
maximum
wind
speed
was 49
knots at 100,000
feet.
The wind
direction changed
from a
north-westerly
direction
to
an easterly direc
tion
with
increasing
altitude.
PERFORMANCE AALYSIS
First
and second
stage
performance
up
to
second
stage
cutoff
(SECO)
is
based
on
the
radar
tracking data
and
NASA
SECO
hardpoint.
PCM telemetry
data was
utilized
to
determine
the vehicle
velocity at second stage
burnout. Reconstruc
tion
of
the
third stage is b'sed
on the SECO I PC14 data
and NASA hardpoint
data
at third stage
burnout.
FIRST STAGE PERFOH4ANCE
An analysis of
pertinent
data
indicates
that the
first-stage flight was near
nominal
with respect
to
the
vehicle's
instantaneous impact point
(IIP)
and
present position traces, which remained
well within
the three-sigma
boundaries.
Table 3
presents comparison
of the guided nominal,
BET-with-winds, and actual
trajectory at
significant times. The
actual
inertial velocity
at
MECO
may be
seen
to
be
2.1
ft/sec
higher
than
the
guided
nominal and
59
ft/sec
lower
than
the
BET-with-winds.
SECOND
STAGE PERFOMANCE
Table 4 presents trajectory
comparisons
of
second stage
performance
parameters
at significant
event
times
during
the first
burn period
of the second stage.
Actual SECO
I
was
determined from Pal
telemetry data.
Figure 4
compares the
tag
second stage
thrust history
with the actual reconstruc
ted
thrust
basedon DIGS
accelerationdata,
flowmeterweight-flow
data
(Reference
4),
weights (Reference5),
and
actualevent
times
(Reference
6).
Thereconstructed
thrustcurveis higher
thanthe
taganaverage
of
48
poundsoverthefirst
burn.
The
higher than tag
thrust
level
resulted
in
a
1.5 second
shorter burn
while
the
low
(-71
ft/sec) velocity at ignition
results
in a
1.4 second longer
burn,
thus
thefirst burnof
the
second
stagewas
within0.1 seconds
ofnominal.
1-2
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THIRD
STAGE
PERFOR4ANCE
PCM
position
and
velocity
vectorsat
SECO
were'used
inthe
following
manner
to reconstruct
third
stageperformance.
An
MDAC-W
predicted
orbit
was
generated
utilizing
a
PCM
position
andvelocity
vectorat
secondstage
burnout,
vehicle
attitudeas
defined in
theBET-with-winds,
actualthird
stage
ignition
time,
and
the
latest
predicted
third
stageperformance
and
burn
time. Areconstructed
orbit
wasthengenerated
using
the same
initial PC01
position
and
velocity
vectors
and
coasttime as
thoseforthe
predicted
orbitin
order
to
determinethe
third
stage
performance
parameters
required
to
match
the NASA
orbit
parameters
at
a
time
after third
stage
burnout.
Thefollowing
table
presents
asummary
ofthird
stage
predicted
and
reconstructed
performance.
Predicted
Reconstructed
Parameter
Unit
Value
Value
Effective
Specific
Impulse
sec
287.93
+1.10 3a
288.39
Total
Impulse
lb-sec
417678.9
418338.82
Impulsive
Velocity
ft/sec
9197.10
9512.09
Vehicle
Attitude
Error;
Pitch
Component,
deg
0+ 3.90
3a
0.699
Nose-UpPositive
VehicleAttitude
Error;
Yaw
Component,Nose
deg
0+ 3.90 3a
0.649
Right
Positive
Table
5presents
comparisonof
the
MDAC-W
predicted,
BEF-with-vinds,
and
reconstructed
third-stage
trajectory
parameters
at third
stage
ignition
and
burnout.
POSTFLIGHT
STATISTICS
Statistical
information
forpertinent
performance
and
trajectory
parameters
are
presented
in Tables
6
and
7.
Table
6provides
data
as
compared
to
nominal
predictions,
while
Table
7compares
to
tag (BET)
predictions.
1-3
http:///reader/full/418338.82http:///reader/full/418338.82
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REFETENCES
1.
Memorandum A3-200-AAC3-M-75-417,
"Guided
Nominal Trajectory
for
the
COS-B Spacecraft
Mission," dated
18 July 1975.
2.
NASA Memorandum,
"Tracking Data
for COS-B
Mission
(Delta
113)," dated
11 September 1975.
3. TWX
A3-130-Delta/AAC3-750137,
"Orbit
Accuracy
Incentive
Criteria
for
the
COS-B Spacecraft
Mission
-
Contract
NAS7-832,"
dated 11
July 1975.
4.
Memorandum A3-226-ADO3-75293,
"Propulsion
Postflight Recon
struction
for COS-B
Delta
Mission
No.
113, Second
Stage,
.DSv-3r-4B (Light
Quartz), Si1
20020," dated
29 September
1975.
5. Memorandum
A3-224-ABE2-75-169,
"COS-B
(Configuration
2913) -
Final
Postflight
Weight Summary,"
dated
2 October 1975.
6.
AVI A3-230-AEFO-AVI-75-234,
"COS-B Sequence
of
Events,"
dated 7
October i975.
1-4
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Table
2
SEQUENCE
OF EVENT3 - COG-B MISSION
Event
Time
Guided
Nominal
Value
From Liftoff (sea)
BET-With-Winds Actual
Value Value
FIFST
STAGE
1. Solid-Motor
Ignition
Arm
2. Solid-Motor
Ignition Command
3.
Telemetry Liftoff
4.
Solid-MotorBurnout
(6)
5. Solid-Motor
Ignition(3)
6. Solid
Motor
Burnout (3)
7.
Solid-Motor
Separation (9)
8.
14CO
Enable
9.
Sensed
0ECO0.5g)
10. Vernier-Engine Cutoff
(VECO)
11. First-Stage/Second-Stage
Separation
Command
-0.90.
-0.20
..-
38.62
39.00
77.81
87.00
230.942
230.322
236.322
238.322
--
38.61
39.08
77.81
87.00
223.653
231.033
237.033
239.033
---.38
o.18
38.47
38.58
77.61
87.34
222.32
227.28
233.80
235.88
SECOND
STAGE - FIRST
BUHUN
13. Second-Stage
Ignition
Command
No. 1
14.
Second-Stage
Engine Start
No. 1
(Steady
State)
15.
Fairing
Separation
(Actual)
16.
Second-Stage
Engine
Cutoff
Command
(SECOM)
No.
1
(DIGS Velocity Cutoff)
17.
SensedSecond-Stage
EngineCutoff
(SECO)
No.
1 (.5g)
243.322
243.662
273.332
530.121
530.832
244.073
244.373
274.033
533.919
534.662
240.81
241.18
270.35
530.56
530.87
1-9.
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Table 2 (Continued)
SEQUENCE
OF EVEITS
- COS-B
MISSION
TimeFromLiftoff(See)
Event
Guided BET-With-Winds
Actual
Nominal
Value Value
Value
STAGE
Fire
SpinRockets, Start Third-
Stage Ignition
Time Delay, and
Start
Third-Stage Timer
3026.621
3027.439
3025.30
Second-Stage/Third-Stage
Separation
Second-Stage
Retro
Initiation
3028.621
3029.439
3027.32
Third-Stage
Ignition
3070.121
3070.939
3071.2
Third-Stage
Burnout
3114.921
3115.739
3115.7
SpacecraftSeparation
3187.00
3187.439
3185.32
1-10
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10RGINAn
PAGE IS
OF POOR QUALM
Table 3
SUWRY
OF
FIRST
STA'E PEBFOh4AI{CE
(COS-a
MISSION)
PARA4ETERS
Item
Unit
Gui
ded
Nominal
Value
BET-With-Winds
Value
Actual
Value
LIFTOFF WEIGHT (LB) 291,785.71 291,288.85 291,290.0
SOLID
MOTORBURNOUT (6)
Time
(Average)
InertialVelocity
sec
ft/sec
38.62
1660.03
38.61
1677.0
38.97.
1668.89
Velocity
(Relative to
LaunchPoint)
InertialFlightPath
Elevation
Angle
Flight
PathElevation
Angle"
Inertial FlightPath
Azimuth
Angle
Flight PathAzimuth
Angle *
Range
Altitude
ft/sec
deg
deg
deg
deg
ft
ft
1,259.8
42.67
63.12
116.37
196.61
6530.9
19,624.4
1,274.5
42.99
63.62
116.17
195.93
6604.3
19,759.2
1,264.1
h3.05
6
4.15
115.51
196.42
6680.2
20,533.4
SOLID MOTOR BUIUOUT
(3)
Time (Average)
see
77.81 77.81
77.61
Inertial
Velocity
Velocity
(Relativeto
LaunchPoint)
Inertial Flight
Path
Elevation Angle
FlightPath
Elevation
Angle*
InertialFlightPath
AzimuthAngle
FlightPath
Azimuth
Angle*
Range
ft/sec
ft/sec
deg
deg
deg
deg
ft
2,657.7
2,640.4
37.34
37.46
161.60
196.48
52,668
2,643.5
2,628.5
37.47
37.55
16l.6o
196.71
52,499
2,679.0
2,664.3
38.27
38.35
161.66
196.67
51,723
Altitude ft
71,030 71,649
72,573
*
Angle
is
with respecttotherelative velocity vector.
1-11
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SUMMARY
Item
;OLID-MOTOR
SEPARATION
(8)
Time(Average)
Inertial
Velocity
Velocity
(Relative
to
Launch
Point)
Inertial
Flight
Path
Elevation
Angle
Flight
Path Elevation
Angle*
Inertial
Flight
Path
Azimuth
Angle
Flight
PathAzimuth
Angle*
Range
Altitude
GUIDANCE
INITIATION
Time
InertialVelocity
Velocity
(Relativeto
LaunchPoint)
Inertial Flight Path
Elevation
Angle
Flight
Path Elevation
Angle*
InertialFlight
Path
Azimuth
Angle
FlightPathAzimuth
Angle*
Range
Altitude
OF
Table
3
(Continued)
FIRST STAGE
PERFOMANCE
(C-B MISSION)
PARAMETERS
Unit
Guided
Nominal
Value
BET With Winds
Value
Actual
Value
see
ft/sec
87.00
2853.1
87.00
28,33.1
87.34
2891.0
ft/sec 2878.3
2864.8
2926.4
deg
33.11
33.19
33.85
deg
32.61
32.60
33.22
deg 166.10 166.33
166.75
deg
n.mi.
n.mi.
196.50
12.06
14.08
196.93
12.00
14.17
196.93
12.12
14.56
sec
ft/sec
125.0
4486.2
125.0-
4465.7
125.0
4528.9
ft/sec
4609.7
4594.5 4674.9
deg
20.52 20.43
21.24
deg 19.77-
19.66
20.38
deg
177.35
177.58
178.53
deg
n.mi.
n.mi.
194.38
23.66
23.66
194.65
23.68
23.68
195.4o
24.43
24.43
1-12
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Table 3 (Concluded)
SU44MAY OF FIRST STAGE PERFORMANCE
PARAMETERS
Item
MAIN
ENGINE
CUTOFF
SIGNAL
Time
InertialVelocity
Velocity
(Relative
to
Launch Point)
Inertial
Flight
Path
ElevationAngle
FlightPath
Elevation
Angle
*
Inertial
FlightPath
Azimuth
Angle
FlightPathAzimuth
Angle *
Longitude
Geodetic Latitude
Range
Altitude
IIP
Time
IIP Range
Weight
Liquid
PropellantUtilization
*
Angle
is
with
respect
to
the
(Cos-B MISSION)
Guided
Unit
Nominal BET-With-Winds Actual
Value
Value
Value
sec
229.942 230.653 226.835
ft/sec
-16443.9
16505.0
16446.0
ft/sec 16486.4i
16547.5
16482.6
deg
11.31 11.30
11.18
deg
11.11 11.10 10.99
deg
179.53 179.54
179.28
deg
184.22 184.21
183.96
deg
121.19
121.19
121.18
deg
31.71
31.69 31.78
n.mi. 184.64
185.67
180.4O
n.mi. 58.58
58.78
57.81
sec
659.52 663.15 652.30
n.m. 1297.21
1309.7 1283.0
lb.
20924.4
26833.4 26833.4
%
99.81
99.81
99.84
relative velocityvector.
1-13
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Table
4
SUAARY OF SECOND STAGE
PERFORMA CE
PARAMETERS
(COS-B MISSION)
Gui
ded
Item Unit Nominal BET
With
Winds Actual
Value
Value
Value
SECOND STAGE
START
Time (Steady
State) see
243.662
244.373
240.811
Inertial Velocity
ft/sec 16426.6 16487.9 16416.5
Velocity
(Relative
to
Launch
Point) ft/sec
16470.0
16531.4
16460.1
Inertial
Flight-Path
Elevation
Angle
deg
o10.47
10.47 10.60
Flight-PathElevation
Angle' deg 10.27 10.27
I0.40
Inertial Flight-Path
Azimuth
Angle deg 179.54
179.55 179.55
Flight-Path
Azimuth
Angle*
deg
184.26 184.25
-
184.26
Range
n.mi..
220.5 221.66 216.9
Altitude
n.mi. 65.5 65.72
64.9
Weight lb
15744.6
15706.0
15708.7
NOSE FAIRINGJETTISON
Time
sec 273.32 274•033 270.35
Weight of Fairing
lb
1320.0 1305.0 1305.0
*
Angle is with
respect
to
the
relative
velocity
vector.
1-14
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---
Table 4 (Continued)
SUMMARY OF SECOND
STAGE PERFORMUCE
PARAMETERS
Item
SECOND
STAGE
FIRST
BURNOUT
SECO)
1
Time
Inertial Velocity
Velocity (Relativeto
Launch Point)
Inertial Flight
Path
Elevation
Angle
Flight Path Elevation
Angle*
Inertial Flicht
Path
Azimuth
Angle
Flight
PathAzimuth
Angle*
Range
Altitude
Weight
Longitude
GeodeticLatitude
Radius
of
Apogee
Radius ofPerigee
Inclination
Eccentricity
Unit
see
ft/sec
ft/sec
deg
deg
deg
deg
n.mi.
n.mi.
lb
deg
deg
n.mi.
n.mi.
deg
Guided
Nominal
Value
530.832
25636.8
25678.2
-0.98
-1.07
179.86
183.26
ll4o.6
121.4
5170.0
122.36
15.75
3678.4
3531.4
89.869
0.020?
BETWith
Winds Actual
Value
Value
534.662 530.873
25639.3
25636.2
25680.8
25677.7
-0.97 -0.97
-1.07
-1.07
179.86 179.86
183.26
183.26
1154.5 1152.64
121.2 121.54
5194.4
5143.62
122.37
122.36
15.52 15.55
3679.0 3677.5
3531.6
3531.2
89.869
89.869
0.02044 0.02029
Angle is with
respectto
the-
relative
velocity
vector.
1-15
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Table 4 (Concluded)
SUMMARY
OF SECOND STAGE PERFURMANCE
PARAMETEIS
(cos-B
MISSION)
Guided
Item Unit
Nominal
BET-With-Winds
Actual
Value
Value Value
PERFORMANCE
MNTE1S
(FIRST
BURN)
Burn Period (Steady
State
286.459
289.546
289.42
to
SECCM
1)
Thrust
(Average) lb 9,707.09
9,548.36
9,596.9
Specific Impulse
(Average)
see
301.77
300.63
300.66
Total
Second
Stage
Impulse lb-sec
2,780,681.9 2,764,689.6
-2,777,539.2
Total
Propellant
Consumed
to SECOM
(Steady
)
State
lb
9,236.4 9,196.45 9,238.0
Propellant
Consumption
(Steady
State
to
SECO
I
)
%
92.01 9i.
8l
•92.3T
1-16
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Table
5
SUMMARY
OF
THIRD STAGE
PERFOICE
PARAMIETERS
(COS-B
MISSION)
Pre-
Recon-
Item
Unit
dicted BET-With-Winds
structed
Value
Value
Value
THIRD
STAGE
IGNITION
Tine
sec*
3,068.82
3,070.94
3,068.82***
Inertial
Velocity
fps
25,159.8
25,152.8
25,159.8
Velocity
(Relative
to
LaunchPoint) fps 25,199.8 25,192.7 25,199.8
Inertial
Flight
Path
Elevation
Angle*
ddg
1.09
1.10
1.09
Flight
PathElevation
Angle*
deg
0.97
0.97
0.96
Inertial
Flight
Path.
Azinuth
Angle
deg
o.i|
0.14
o.14
Flight
Path
Azimuth
Angle*
deg
356.77
356.77
356.77
Longitude
deg
-47.01
-47.01
-47.00
Geodetic
Latitude
deg
-23.19
-23.19
-22.91
EulerAttitude
Angles'
Pitch
(epB)
deg
169.56
169.56
170.37
Y
a(*PB)
deg
15.57
15.57
15.01
Roll
(P
)
deg
-79.09
-78.79
-79.00
Range
n.mi.
9,849.9
9,862.7
9,849.9
Altitude
n.mi.
189.2
190.7
189.2
Weight
lb
2,262
2,262
2,262
*Angle
is
with
respectto
the relative
velocity
vector.
**Euler
angles
epB, *PB9
and *PB
arethe
angles
specifying
theorientation
of
thevehicle
axes
(xB YPB
3
and
ZpB)with
respect
to
an
inertial
reference
platform.
The
order
ofrotation-
is:
Pitch,
OPB
about
YPB
(positive
turning
ZPBintoXPB)
;
yaw,
PPB
about
ZB
(positive
turningXPB
into
YpB);
and
roll,
*PB
about
XPB(positive
turning
YPBintoZpB)
,
in degrees.
**'41.5 seconds
afterstage
II/Ili separation.
.
,-
oo
http:///reader/full/3,068.82http:///reader/full/3,070.94http:///reader/full/3,068.82http:///reader/full/3,068.82http:///reader/full/3,070.94http:///reader/full/3,068.82
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SUMMARY OF
Item
THIRD
STAGE BURNOUT
Time-
Inertial
Velocity
Velocity (Relative
to
LaunchPoint
InertialFlightPath
ElevationAngle
FlightPathElevation
Angle*
InertialFlightPath
AzimuthAngle
Flight
PathAzimuth
Angle*
Longitude
Geodetic
Latitude
Euler
Attitude
Angles**
Pitch
(ePB)
Yaw(
pB)
Roll
(OPB)
Table 5 (Continued)
THIRD STAGE
PERFO1ANCE
,(COS-B
ISSION)
Pre-
Unit
dicted
Value
sec
3,113.62
ft/sec
34,615.3
ft/sec
34,648.7
deg
2.56
deg
2.44
deg
360.00
deg 357•48
deg -46.82
deg
-19.52
deg
169.56
deg 15.57
deg
-78.79
PIRAMETERS
BET-With-Winds
Value
Recon
structed
Value
3,115.739
34,609.8
3,113.62
34,622.8
34,643.1
34,660.5
2.49
2.76
2.37
2.64
360.00
359.84
357.119
-46.83
-19.79
357.32
-46.82
-19.52
169.56
15.5T
-78.79
170.37
15.01
-79.00
*Angleit
withrespectto
therelative
velocityvector.
**Euler
angles
"pB'*PB,and
OPBare
the
angles
specifyingthe
orientation
ofthe
vehicle
axes
( B, YPB'
andZP)
with
respectto aninertial
referenceplatform. The
order
ofrotationis
pitch,0p aboutY.,
(positivetuning
ZPB
into XPQ; yaw,
pp.
about
Z
3
(positive
turning
XPB
into
Yp);
and
roll, .PBand
XPB
(positive
turning
YPB-into
t B
)
,
indegrees.
1-18
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Table
5 (Concluded)
SUIRIARY
OF
THIRD STAGE
PERFOPBAICE
PARAMETER&
•(COS-B
MISSION)
Pre-
Recon-
Item
Unit
dicted lBET-With-Winds
structed
Value
Value
Value
Inertial
Attitude Angles*
Elevation
Angle(e'L)
deg 5.30
5.03
6.04
Azimuth
Angle ( 'L)
deg 359.62
359.63
359.02
Roll
Angle
(L
)
deg
-82.11
-82.10
-82.16
Range
n.mi.
9,681.5
9,695.4
9,681.4
Altitude
n.il.
194.79
196.2
195.2
Weight
lb
811.5
811.6
811.5
TotalThird
StageImpulse
lb-sec
417,678.9
417,678.9
418338.8,
Spacecraft
Weight
lb
612.0
6!1.5
614.15
Radius of
Apogee
n.mi.
57,670.8
57,732.8
58,204.6
Radius
of
Perigee
n.mi.
3,629.7
3,631.4
3,628.8
Inclination
Angle
deg
90.000
90.000
90.155
Eccentricity
-
.8816
.8816
.8826
ArgunentofPerigee-
deg
335.13 335.00g
334.7
The
vehicle
centerline
elevation
angle
0'L
istheangle
between
the
vehicle
centerline
and
the
plane
pernendicular
to
the radius
vector
from
the
centerof
the
earthto the
vehicle
(positive
for
the
vehicle
nosepointing
away
from
the earth),
in degrees.
Vehicle
centerline
azinuthangle
is
the
angle
between
thelocal
meridian
andthepro
jectionofthevehicle centerlineonto
a
planeperpendicularto
the
radius
vectorfrom
the
center
oftheearth
to the
vehicle
(positive
clockwise
from
true
north),
in
degrees.
-
Thevehicle
instantaneous
geocentric
roll angle
*'L
is
in
degrees.
ORIGINAL
PAGE 18
OF POOR QUAL
1-19
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Table 6
SUM4ABY
OF
POSTFLIGHT
STATISTICS
AO'NAL
PRDICTCONS
(COS-3
MISSION)
BOOSTER:
DSV-3P-1A
Extended
Long
Tank
SECOND
STAGE: DSV-3P-B
(Quartz)
COS-B No.
of
Mean
Sigma
About
Parameter
Deviation
from
Guided
Nominal
Samples
Deviation
Mean
SOLID
NOTo0s
(6)
Drop
Time
+0.34 .
-0.235
0.6 6
GUIDANCE
INITIATION
Altitude
(n.mi.)
+0-77
10
0.623
0.527
Inertial
Velocity(ft/see)
+92,7
10
60.7 97.7
Inertial
Flight-Path
Elevation
Angle (deg)
+0,72
10
0.94B
'0.360
Inertial
Flight-Path
Azimuth
Angle. (deg) +1.18
10 -o0;01
0.621
NECO
Tine
(see)
-3.107
10
-1.555
2.825
Altitude
(n.mi.)
-0.77
10
0.037
0.523
Inertial
Velocity
(ft/sec)*
154
10
179.2
110,7
Inertial Flight-Path
Elevation
-0.12
10
0.040
0.130
Angle
(deg)
Inertial
Flight-Path
Azimuth
-0.15
10
0.
016
0.079
Angle
(deg)
BasedAn
DIGS telenetmr
data
and predicted
nominal
MECOsalocity.
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TABLE
6
(Concluded)
SUMMARY
OF
POSTFLIGHT STATISTICS
NO4INAL
PREDICTIONS
COS-B
MISSION
BOOSTER:
DSV-3P-1A
Extended Extended
Long Tank
SECOND STAGE:
DSV-3P-4B
(Quartz)
Parameter Deviation
COS-B
from
Guided
Nominal
No. of
Samples
Mean
Deviation
Sigma
About
Mean
SECOND
STAGE
First Burn
Time (sec)
Propellant
Consumption
Through End
of Primary
Mission
(%PU/oPU)
2.892
0.66
l0
10
0.488
-0.771
2.902
1.320
SECO
I
Altitude
(n.mi.)
Inertial
Velocity
(ft/sec)
Inertial
Flight-Path Elevation
Angle (deg)
Inertial
Flight-Path Azimuth
Angle
(deg)
+0.10
0.6
0.01
0.00
1l
11
ii
11
o.o14
-0.493
0.0004
0.063
0.242
3.644
-0,010
0.090
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Table 7
SU4MARY
OF POSTFLICUT
STATISTICS
TAG
PREDICTIONS
CS-B
MISSION
BOOSTER:
DSV-3P-lA Extended
Long
Tank
SECOND STAGE:
DV-3P-4B
(Quartz),
Parameter
COS-B
Deviation
from
Tag
No. of
Samples
Mean
Deviation
Sigma About
Mean
BOOSTER
Burn
MECO
MECO
Time (see)
Inertial
Velocity
(ft/sec)*
Altitude
(n.mi.)
-3.497
+92
-0.98
10
10
10
-i.736
176.9
0.021
1.820
121.2
0.533
SECOND STAGE
FirstBurn
Tire
(sec)
Propellant
Consumption
Through
End
of Primary
Mission
(%PU/CPU)
Tailoff
Impulse (lb-sec)
-0.227
0.035
147
10
10
10
-4.988
0.035
44.778
2.867
i.o66
102.404
* Based
on
DIGS telemetry
data and predicted
nominal
MECO velocity.
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____
- -- - -- --------
------- -
-
--------------------
EXTENDED LONGTANKDELTACOS-B
MISSION FIGURE
I
UPPERAIR
TEMPERATUREDATA
WESTERN
TEST
RANGE VANDENBERG AFB
8 AUGUST
1975
0147
GMT (
1847
PDT)
10
0-:'!!'-P==
:
.
.. ..
,..
..
.
.....
...
..
9
0'----r
~1962:
INTERNATIO10
1
:z
...
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--
ORGAN*1ZAT[ON
:r .....
..
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,VIO
,
..
......
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T W!, .........
E: ...
I ' -ATMOS6 .
8/9/2019 Delta 113 Postflight Report
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-
-
-
-
--
-
-
-
-
-
-------
EXTENDED
LONG
TANK
DELTA
COS-B
MISSION
FIGURE
2
PRESSURE
DATA
WESTERN
TEST
RANGE,
VANDENBERG
AFB
8
AUGUST
1975
0147
GMT L847
PDT)
.+ ...............----
.-.., ...
. --
..... , . . ..-.. ..
. ..
....
-
.,-
.....
.. ....
NIA..
...
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NR
9
O
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T-A--.-------
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.. .......
90
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t-
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'z~~
1
1
90- -----------
-
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.-.
CT--A
-
ERE:
TMOSP-
--
---
__ ......
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._ _:E
_
+'
I
4 ~:-L---'-.-I9SJNTERNA1OACIL
-
---
--
AVIA-TION
ORGANIZATIO
I.-- --
:- - -:-+-4
... .. { ..... -----..
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'_-- - - -
- ..
.:.
---------------
ATM0SPHE.RE
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...... _
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----
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- +..
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zz.tz- --. -t -"+.'%.
"-
.- -
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,- -
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--
. -
- - - - - - - - - - - - -~~- --
. . .----- .
0
200
400
600 800
200
PRESSURE.,
P
(MILLIBARS)
1-24
http:///reader/full/ATM0SPHE.REhttp:///reader/full/ATM0SPHE.REhttp:///reader/full/ATM0SPHE.REhttp:///reader/full/ATM0SPHE.RE
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-
-
-
-
-
-
-
-
-
-
-
-
-
-
-
-
i
-
-
-
1
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.
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t
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1
2
-
,
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EXTENDED
LONG
TANK
DELTA
COS-B
MISSION
FIGURE
4
VEHICLE
TOTAL
THRUST
SECOND
STAGE
SIN
20020
.
.
..
.A
G
:
T
k
R
U
S
.
.
.
-:.
.
.
MA
..
HIGH-A
AD.:LOW
EXTREMAa
.--
RECONSTRUCTE
D-TH
RUST_4>-Tm
>1-_
----o----.CTU.-.--LE
_
gm
.
10.5-§
4
~
'
t
m
....
O,~~~~
~~~O- .....
..- .........
_.=.- -= ---
. ... . .
... .
=:.
=: ---------..
=.--:
T-
SM
MAA...
F.
........
=
=
. f
ypf.lf-fi-
Y....
..
.
t.z--,.
10.0-
H; _-t .
...
.
4
L
...
~ .i
.
............
......
...
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9.0-
r
4
t--:t
S....... ..
-HH7
<
8.5
4.
__
4-_ --
-_-
w_.r
00
700...,-
.060
-
.....
0
=C-
.--=
--
-
=-
......
...........
.
.
.
'
a- 0
200 300 400 500
600 700.
TIME.FROM
LIFTOFF,
t
(SEC):
1-26
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Attachment
2to:
A3-262-AMOO-M75-509
ATTACHMENT'2:
SECTION
2. PROPULSION
SYSTEMS -
COS-H MISSION
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Attachment
2 to:
A3-262-AMO0-M75-509
SECTION 2
PROPULSION
SYSTEMS - COS-B MISSION
2.1 INTRODUCTION
Overall performance
of
the
COS-B
launch
vehiclepropulsion
systemswas
satisfactory
throughout
first,second, and
thirdstage
flight. All data
returnedby
telemetry
channels usedto
monitor
propulsion
systemsperformance
weresatisfactory
with two exceptions. The
LOX pumpinlet
pressuretransducer
failedatapproximately
74
seconds
from
liftoff,andthe
second
stage
chamber
pressure
exhibited
the
same
anomalouscharacteristics observed
onprevious
SSPU
flights.
Asummary
ofvehicle
model
and serial
numbers
isprovidedin
Table 2-1.
There
were
sevenfirst
flightitems
on
this
flight. Thesefirst
flightitems
are
listed
in
Table 2-2. All flighttimesin
the text
are
givenin
secondsafter
DIGS
indicated-liftoff
unless
otherwise
noted. The
"DIGS
Liftoff"
time
for
this
flight
was
defined as
the
timeatwhichthevehicleachieved
approximately
37.5 ft/sec
2
acceleration
(about5.3
ft/sec
2
offthe
pad)
forfourconsecutive
20-millisecond
time
intervals. The
propulsion
system
sequence
of
events
is
summarizedin
Table2-3.
Reconstructedboosterand
second
stane
performance
parameters
arecompared in
Table2-4withcorrespondingvalues
from
the latest
boosternominal simulation
(Reference
2-1) and the Detailed
TestObjectives
Report
(DTO.,
Reference 2-2).
Values from
both
reports
are referred
to as nominal
values sinceresultsofthe
nominal
simulation
are
used
to
generate
the
DTO.
Firstandsecondstage
reconstructedvalues
are
compared
also
with
valuesfrom
the
BestEstimateTrajectory
without
winds
(BET,
Reference
2-3) andthe
Propulsion
preflight
tagpredictions
(References2-4 and 2-5) in
Table
2-4.
Cumulativestatistics
forall comparisons
in
Table2-4
arepresented
in
Table2-5.
2-1
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In
this section, the
word "predicted"
refers to preflight
predictionsof
performance for
the
Propulsion systems
utilized
on
this
vehicle
(References
2-4
and
2-5)
anddoes not
denote nominal
or
DTO values.
The
word "reconstructed"
refers
to
postflight reconstructions ofperformance
oenerated
usingtelemetered
system
pressures, temperatures,
event
times,
acceleration, and
(forthesecond
-stage)
flowrates.
The
term "internal" refers
to
reconstructions
based
on
pressures,
temperatures,
andevents.
The
term"external"
refers to
reconstructions
using
acceleration
and.the
internallyreconstructed
vehicle
mass history.
2.2 FIRST
STAGE PERFORMANCE
Performance
of the first
stage
propulsion systems
is described
in
thefollowing
paragraphs.
2.2'1
Main
Engine
All valid telemetry
data
indicated
that
the
main engine
flight
performancewas
satisfactory,
as summarized
in
Table
2-4.
Figure
2-1
shows
the
nominal and
actual
start
sequence.times.
Theagreement
observed between
the nominal
and actual
sequence
times indicates
a
normal
start
sequence based
on available
engine
statistics. The
mainengine start
sequence was initiated
2.338
seconds
prior to liftoff
(DIGS).
Main engine
cutoff (MECO)
occurred
226.835
seconds
after
liftoff
due
to
actuation
ofthe fuel
injector
pressure
switches
(FIPS). Thepropellant residual
atMECO
was 281
pounds
of
fuel
or
19
pounds
less than
the
loaded
bias of 300 pounds.
The
residual
corresponds
toapropellant
consumption
andapropellant
utilization
(PU)
of
99.84percentand
mixture
ratio variations of
-0.0007
mixture ratio
units
(mru)
from
the
preflight prediction
and
-0.0187
mru
from
the
ground test
tag
prediction.
Figures
2-2and
2-3 present
the
internal
reconstructed
liquid
engine
thrust
and
flowrate histories,
respectively. The
overall
performancewas
good
and
generally
verified
the
performance
model.
2-2
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TheDIGS
thrust
acceleration
measurements,together
withpreflightpredicted
drag
andpostflight internal reconstructed
vehiclemass,
were
used
to compute
total
external boosterthrust
and specificimpulse.
Figure
2-4
depictsthe
internaland
external reconstructedthrust
historieswhich agreeveryclosely
from
liftoff
to
MECO.
Table
2-6
presents
a
comparison of averagesfortotal vehiclealtitude
thrust
andspecific
impulsebetween120
seconds
and
MECO
and
the
averagemixture ratio
over
theentireflight. Theinternal
reconstruction
indicates
first
stage
Isp
was
about
0.52seconds
higherthan predicted; the
external
reconstructionshows
a
decrease of
about
1.45
seconds
from
the
predicted
value.
Thepreflight pre
dictedand internal andexternal reconstructed
valuesare
comparedwith
the
ground
testtag
prediction
in
Table
2-6.
Cumulativestatistics for
these
parameters
are
also
tabulatedin Table2-6.
2.2.2
POGO
Suppression
System (PSS)
The
PSSwaspressurized
froma464psia
reoulated
AGE
sourceuntil
liftoff.
No
inflightpressurization
wasprovided,
thus
the inflightLOX
volumewas a
functionofthe
ullagegas mass and
temperature
and
of the LOXpump
inlet
pressure. -The',iotemperatureprobesinthe
PSS showedthat
the
PSS performed
satisfactorilyandthe LOX
and
ullage
gas
volumeconstraints
weresatisfied
throughout theflight.
The upprprobe
may
havebeen
coveredmomentarilyby
splashingatliftoff.
2.2.3 Vernier
Engines
Vernierengine
performanceappearedsatisfactory
based on telemetered
chamber
pressuredata
from
vernier
engineNo.2. Reconstruction
indicatesthattheengine
was
operating
at
a
thrust
level of
977pounds durina
vernierengine
solo,
which
is
less than
the
nominal thrust
of
1002pounds.
2-3
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c 3cq
not~ors
ic
Based
on telemetered
data and
reconstructed performance
values,
the performance
of
the
solid
motors
was satisfactory. Table 2-4sumnarizes
solid
motor
per
formanceandTable
2-5
presentscumulativestatistics
for
Castor
II
motors.
The reconstructed solid
motorperformance-is
based
on
event
times
andthe
chamberpressure
histories.
Total burn
times for
both
the
qround-ignited
and
altitude-ignitedmotor
sets- weregenerallyslightly
less thanoredicted. Web
times
were
slightly
greaterthan predicted. All
burn andwebtimes werewell
withinthe allowable
dispersion
band of
the Castor
II
motors.
All
of
thesolidmotorstart
andthrustbuilduptransientswerenormal. Total
thrustand
flowrate
historiesfor
the
solid
motors
plus
themafnengineare
shownin
Figures
2-4and2-5,respectively.
2.3 SECONDSTAGEPERFORMANCE
The
second
stageengine
operated
normally
forthe
first
burn-of
289.75seconds,
whichwas
0.75
secondsshorter
than
theBET
prediction. Theexperimental
restarthadadurationof
25.06
secondsfora
total
burn
time about
6.1
seconds
less than
the
predicteddepletion
burn time
of
320.6
seconds.
A
fuel
depletion
wasobserved.
Asummaryofsecond
stage
firstburn engineperformance
is'
presentedin Table
2-4
whilespecific
impulse,thrust,andflowrate histories
are
depictedin
Figures
2-6,
2-7,
and
2--B, respectively. Table 2-5presents
statistical
dataforvaluesin Table2-4.
2.3.1
FirstBurn
Duringfirst
burn
operation,the
secondstacetemperaturesandpressures
were
nominal.
Thepropellant
tanks pre-pressurization
sional occurredat
229.83
seconds,
3.00
seconds
afterMECO(FIPS)command.
Duringthe
pre-pressurization,
the
helium-bottle,helium
regulator,and
propellant
tank
pressures
wereas expected.
2-4
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2.3.1.1 FirstBurn rransientPerforana±
The total
start
transient
impulsecalculated using
chamberpressuredatawas
346pound-seconds
compared
to
theprevious
average-flight
value
of383pound
seconds.
The
total propellant
consumed
duringthe
starttransientwas
2.31
pounds compared to
the
averagevalue
experienced of 2.43
pounds. The
shutdown
propellantflowto
propellant
valvesclosure
was6.36pounds
compared
to
the
average
value
experienced of6.23 pounds.
DIGS
accelerometer
dataindicate
ashutdownimpulse
of
3187pound-seconds,
compared
to
the
3040
pound-seconds prediction
derived
fromanalysis of
data from
previous flights. The
shutdowntransient
performance
is summarized in
Table2-4.
2.3.1.2
Steady-State
Performance
Second
StageignitionCommand-No.
1(SSIC
No.
1) occurred
240.81
seconds
after
liftoff
and Engine
StartNo. 1occurred
0.37secondlater
at241.18
seconds.
SECOMNo.
I
occurred
530.56seconds
after liftoff
as theresultof
a
planned
DIGS-initiated
cutoffcommand.
Therefore,the
propulsionsystem
firstburn
steady-state
poweredflightduration
(fromEngine
Start
No.
1to
SECOMNo. 1)
was289.42seconds.
This
timewas
1.08
secondsshorterthanthe
BETpredicted
duration.The
reconstructed
average
thrustwas-
higherthanpredicted,as
was
theaverage
flowrateyieldingan
averagespecific
impulse
thatwas 1.75seconds
lower
thanthe
BETprediction.
Althoughit
did
not
adversely
affect
theprimarv
mission,theCOS-Bvehicle
experienced
ananomalousvibration
which
occurred
from165
to 212seconds into
second
stageburn. This anomaly
isdiscussedin
detail in
Anomaly
Report
No.
T00166.
The
vibration-had-an
acceleration
level of
approximately
2g'szero
to peakinthethrust
axisat
afrequency
ofapproximately
130Hz,as
measured
at
theguidancesection.
The
fuel manifold
used
on
COS-Bwasof
anew
buy,built
especiallyfor the
Delta
-Program(previous
fuel manifolds
weredesigned
for
the
LunarModule
Descent
Engineor
LMDE).
The
Deltafuel manifold-Incorporated
minorproduction
changes, including
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a
weldbeadat
theinlet. Correctivemeasures
being
considered at this
time
includethe
removal oftheweld
beadandstiffeningof
the
thrust
mount.
Analysis
is
continuing
as
additional
test
data become
available.
During
the
interim,
silica
chambers
are being
flown
on
stages
utilizing
the
Deltamani
fold
to
improve
stabilitymargins.
Anotheranomaly
withrespectto mixtureratio
andspecificimpulse
was
identified
and
is
discussed
in
detail in
AnomalyReport
No. T00168. Reconstructionof
inflightperformanceindicated anapproximate 0.010 mrushiftin mixtureratio
(M.R.)
startingduringthe
period
of
130Hz
oscillations.
The initial
re
constructionalso
indicated
anapparent
1%
lower
thanexpectedspecific
impulse
(Isp)
throughout
first burnengine
operation.
Thespecificcauseof
themixture
ratio
shifthas
notbeen
determined,but
is
consideredaneffect
of
the 130Hz
oscillations
due
to thesimultaneous
onset
times.
Possible
causesoftheM.R.shift are:
1) cracks
in
the oxidizer
pintle
slots
as
a
resultofthe
oscillations,or2) theeffect
ofoscillations
on
flowmetercalibration
(althoughthe
apparentincrease
in
oxidizerflowrate
is
notconsistent
withpostulated
flowmeterfailuremodes).
Theapparent
low
specific
impulse
has
beenattributed to
flowmetercalibration
error. A
detailed
evaluationof
propellantdepletioncharacteristics indicated
a
biasin the
oxidizer
flowmeterovertheentireengine burntime.
Withthe
bias
taken
out
ofthe
flovneterdata,thecalculated
specific
impulse
is
normal.
Normal
performancewas
alsoverifiedbystage
velocitydata;
therefore,
it
is
concludedthat theenginespecific
impulsewasnormal.
Information
presented
in
this
reportreflectsthe
correcteddata.
Predictedandreconstructed
values
forpropellant
consumption
werecomparable.
Approximately808pounds
ofusable
propellant
remained
on board
afterSECOM
No.
1
in
reservefor-the
secondand
third burns. Thisrepresentsa
firstburnpro
pellant
consumption (PC)
of
92.37
percent. Accordingto
the integrationof
flowmeterdata,
theaveragefirstburnmixture
ratiowas
1.584mru,less than
the1.598
predicted
but
withintwosigma
ofthepredicted
value.
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Reconstruction
of
thrust
from
acceleration
data
yields
an
average
specific
impulse
of 300.66
seconds
compared
toaBET predicted
specific
impulse
of
302.41
seconds.
As
depicted
inFigure
2-6, from
approximately
ignition
plus
200
seconds
the
reconstructed
specific
impulse
decreases
at
a
faster
rate,
movingfarther
away
from
the
predicted
level.
The
start
of this
decay
in
specific
impulse
(Isp) is coincident
with
the
beginning
ofthetailoff
portion
offlight.
Basedon chamber
pressure
data,
the
reconstructed
throat
erosion
was-1.2
percent
compared
to
apredicted
value
of 2.2percent.
2.3.2 Coast
The
secondstage
coastedfor
approximately
2688seconds
betweenthefirstand
second
burn. All monitored
pressure
and temperature
valueswere
acceptable
during
coast.
The
fuel
tank
pressure
increased
by
about
15 psi and
the
oxidizer
tankpressure
rose
about 27 psi
during
coast
due
to heating. Character
istics
of
the fuel
tankpressure
data
indicateproper
levels throughout
the
mission.
2.3.3 SecondBurn
(Experimental)
Thesecond
burn (as
reconstructed)
was
initiated
at3218.3
seconds
after
liftoff.
Theburnwas
preceded
bya
settling
period ofapproximately
15seconds.
No
actual
data wereavailable
forrestart.
Therefore,
no conclusion
can
be
made
as to theadequacy
of
the
settling
period.
The
restartsteady-state
burnduration
was
approximately
25.06seconds
which
was
5.79seconds
shorter
thanthepredictedvalue.
Restart
burn
times
usually
are
shorter
than
predictedbecause
heating
duringcoast
increasesthe propellant
tank
pressures resulting
in higher
thanpredictedthrust
andflowrate
levels.
Approximately
55 poundsofpropellant
remained
on board after
SECOMNo.
2.
This corresponds
to
a
propellant
consumptionvalue
of
99.74percent
forthe
twoburns.
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--
2 3.6
Nitrogen
Auxiliary
Propulsion
-Systei
(APS)
Predicted
and
actual
impulse
usagesfrom
the nitrogen
APS
for
variousevents
are
tabulated
below:
Event
Usage
Predicted
Actual-Predicted
First
burn
18
16
2.3
Firstcoast
297
430
-133
Separation-and
retro
54
44
10
Plume
impingement
14
40
-26
Settling
125
115
10
Experimental
burn
1
1
0
Total
509
646
-137
Overall
APS
performance
was
normal.
Stage
Initiation
OF
POOR
41AI
.3.7 Second
Retro
OR I
p
Second
stage
retro
initiation
occurred
at3027.29
seconds
after
liftoff.
Duringretro,
bottlepressure
decayednormallyfrom
190psia
to
approximately
zero
psia.- Based
on
the
DIGS systemintegrated
velocity
value,
the
separation
distance
at
thirdstage ianition
was approximately
42.3
feet, compared
to
the
minimum
required
of
25 feet.
2.4
THIRD
STAGE
PERFORMANCE
2.4.1 Spin
Motors
The
spin
table
microswitch
data indicate
that
the
eight
spinmotors
were
fired
at3025.30
secondsafter
vehicle
liftoff
and
produced
aspin
rate
ofapproxi
mately
39.3
rpm
(versus
predicted39.8
rpm)
at
thirdstage/spin
table
separation.
It
is concluded
thatall
eight
spin
motors
performed
satisfactorily.
2.4.2 Third
StageMotor
Performance
ofthe third
stage
motor
(TE-M-364-3,
S/N00025)
was satisfactory
based
on
theaccuracy
of
thespacecraft
orbit.
Thechamber
pressure
data
:exhibited
a-10
psi shiftat23
seconds
afterignition,
and
acomplimentary
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+10
psi
shift
at
motortailoff.
The
shape
of thepressure-time
curve
appeared
to be
normal bothprior
to
and after
the pressureshifts.
Accelerometer
data
did
notexhibit
corresponding
shifts
in
the acceleration
level,
and
itishere
fore
believedthat
the
chamber
pressure
data shifts are
not
indicativeofactual
motorperformance.
Immediatelyafter
third
stagemotor
tailoff,
lowlevel
oscillationswerenoted
on thespacecraft
attachfitting
accelerometers. The
oscillationshad
amaximum
amplitude
of2.3G's
O-peakatafrequency
of 800-1000
Hz,
and
lastedfor 18
seconds. During
this period
of
time,
the motorchamber
pressure
did
notregisterany
activity. The
cause of
the oscillations
isunknown
at
this
time,
but
isunder
investigation.
Predicted
third
stagesolid motor
performance
parameters,
obtainedfrom
Reference
2-6,
are listed
in Table
2-7.
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REFERENCES
2-1
Memorandum
A3-250-AD03-74090,
dated21 March
1974
2-2
Memorandum
A3-200-AAC3-11-75-357,
dated
June
23,
1975
2-3 Memorandum
A3-200-AAC3-M-75-454,
dated
6
August
1975
2-4
'Memorandum
A3-226-AD03-75191,
dated10 July
1975
2-5
MemorandumA3-226-AD03-75190,
dated2
July1.975
2-6 Memorandum
A3-226-AD03-75178,
dated
23
June1975
2-10
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TABLE2-1
VEHICLE
AND VEHICLE COMPONENT:
IDENTIFICATION
SUMMARY
(COS-B)
Item
Launch Vehicle
First
Stage
Main
Engine
Vernier
Engines
(two)
Solid
Motors
(Set
No.
I three)
Solid
Motors
(Set
No.
2
-
three)
Solid
Motors
(Set
No.
3
-
three)
Manufacturer
MDAC
MDAC
Rocketdyne/A
Division
of
Rockwell
International
Corporation
(RD)
RD
Thiokol
Corporation
(T)
TC
TC
Second
Stae
Propulsion
System
MDAC
Engine
TRW
Third
Stage
TC
Spin
Motors
AtlanticResearchCorporation
(ARC)
Model
DSV-3P-11B
DSV-3P-lA
RS2701A
LR-lOl-NA-11
TX354-5
(Castor
II)
TX354-5
(Castor
II)
TX354-5
(Castor
II)
DSV-3P-4B
TR-201
TE-M,364-3
ID00399-529
Serial
No.
20018
20020
(602)
0017
No.
1:
No.
2:
338180
338181
No. 1:
No. 2:
No. 3:
473
474
475
No.
4:
No. 5:
No.
6:
480
535
489
No.
7:
No.
8:
No. 9:
494
498
548
20020
1016
00025
- - -
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TABLE
2-2
FIRST
FLIGHT ITEMS
(COS-B)
1.Castor
IImotors
with
bimodal
oxidizer
propellant:
CastorIIMotors,
S/N4's
473,
474,
475,
480, and
489,
were
loaded
with
bimodal
oxidizer
propellant
aft
of the
aftpropellant
slot. This
isa
departure
from
the
standard
trimodal
oxidizer
propellant.
2. Castor
IIdirect-mountpressure
transducer:
Previously,
the solid
motor
chamber
pressuretransducerwas
mounted in themotor
forwarddome
with
a
shorthardline
connection
to
the
motorpressure
port.
Thechange
to
the direct-mountconfiguration
was
accomplished
to
reduce potential
hardline leak
paths
and
facilitate
launch
siteinstallation.
3.
Castor
II
solid
motoraluminumwiring
tunnel:
The previous fiberglass
tunnel cover
has been
replaced
with
asimilar design
fabricated
of
aluminum including
cork
sheet
interior
insulation.
The
change
to the
aluminum
tunnel
was
implemented
to
reduce
cost.
4.
Over-age
TE-M-364-3
third
stage
motor: This
was theoldest
third
stage
motor
to
be
flown
(44
months
old at
launch). The
oldestmotor
previously
flown
was S/N00026
on
SKYNET-IIB
(35
months
old).
5.Second
stage
fuel and
oxidizertank
shutoff
valves:
The
Fuel Tank
Shutoff
Valve
(FTSV), P/N
IU96916-1,
and Oxidizer
TankShutoff
Valve
(OTSV),
P/ti
1B96916-501,
replace
the
FTSV, P/N IB95417-507,
and
OTSV,
P/N 1B95417-509,
effectiveDSV-3P-4B,
S/14
20020,
and subsequent.
The Pneudraulics,
Inc., P/N
9386,
restrictor
check valve
(PRCV)
was
replaced
by the1B97422-1
PRCV
in order
to
obtain
proper
OTSV-FTSV
differential
opening
time
with
the
newTSV's.
Physically,
the
two
valves
are
identical except
forthe
sizeof
the
restrictor
flow
orifice.
The incorporation
ofthe
new
TSV's
also
required
minor
modificationof
the
interconnecting
tubes
between
the PRCV
and
the
TSV's,
andachange
in
-thepressurization
sequence.
6.Second
stage
fuel
pressurization
fittingmodification:
The 1294500-l
fuel
tank
pressurization
fitting
located
on
the fonard
domeof
the
SSPU
fuel tankwas
modified
by
having
its
sense port
increased
from
0.098 inch
diameter
to
0.300
inch
diameter.
In
addition,
a
0.012
to
0.013
inch diameter
hole
was added
between the
fitting
inlet pressure
port
and
the
fuel tank
top pressure
sense
port.
7.POGO
accumulator
vendor
change:
The
1B89068-507
POGO accumulator
(manufactured
by
Solar)
was replaced
by IB96342
(manufactured
by Coast
Metal
Craft).
The
vendo