Delta 113 Postflight Report

8/9/2019 Delta 113 Postflight Report

1/217

,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


2/217

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


3/217

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


4/217

Attachment Ito:

A3-262-AMOO-M75-509

ATTACHMENT

1:

SECTION

1.

SYSTEM

PERFORMANCE-

COS-BMISSION


5/217

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


6/217

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


7/217

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


8/217

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


9/217

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13/217

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.


14/217

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


15/217

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


16/217

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


17/217

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


18/217

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


19/217

---

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


20/217

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


21/217

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


22/217

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


23/217

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


24/217

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.


25/217

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


26/217

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.


27/217

____

- -- - -- --------

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EXTENDED LONGTANKDELTACOS-B

MISSION FIGURE

I

UPPERAIR

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WESTERN

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RANGE VANDENBERG AFB

8 AUGUST

1975

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28/217

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TANK

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COS-B

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FIGURE

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1975

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29/217

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30/217

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

.--

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RUST_4>-Tm

>1-_

----o----.CTU.-.--LE

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=-

......

...........

.

.

.

'

a- 0

200 300 400 500

600 700.

TIME.FROM

LIFTOFF,

t

(SEC):

1-26


31/217

Attachment

2to:

A3-262-AMOO-M75-509

ATTACHMENT'2:

SECTION

2. PROPULSION

SYSTEMS -

COS-H MISSION


32/217

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


33/217

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


34/217

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


35/217

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


36/217

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

2-5


37/217

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.

2-6


38/217

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.

2-7


39/217

--

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

2-8


40/217

+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.

2-9


41/217

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


42/217

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

- - -


43/217

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

Delta 113 Postflight Report

Documents