-

c - 67-37 Y ‘f

A LLO GUIDANCE AND NAVIGATION

TYPE I DOCUMENT APPROVED BY NASA

Revision 1 R-537

(Unclassified Title)

GUIDANCE SYSTEM OPERATIONS PLAN AS-501 VOl. I I

CONTROL DATA AND ERROR ANALYS IS December 1966

I N S T R U M E N TAT I 0 N LAIBORATORV

C A M B R I D G E 39, M A S S A C H U S E T T S

COPY# 50 OF 0 COPIES THIS DOCUMENT CONTAINS =;& PAGES

ACKNOWLEDGEMENT

This report was p repa red under DSR Pro jec t 55-23850, sponsored by the Manned Spacecraft Center of the National Aeronautics and Space Administration through Contract NAS 9-

4065.

This document contains information affecting the national defense of the United States within the meaning of the Espionage Laws, Title 18, U.S.C.,

Sections 793 and 794, the t ransmission o r the revelation of which in any manner t o an unauthorized person is prohibited by law.

ii

(Rev. 1 - 12/66)

Table of Contents

Volume I Operations

Section 1 Introduction

2 G&N Flight Operations Summary

3 G&N System Description 4 Mission Logic and Timeline 5 Guidance Equations for CSM

Volume I1 Control and Error Data

Section. 6 Mission and Vehicle Control Data 7 G&N Error Analysis

iii

I

6 MISSION AND VEHICLE CONTROL DATA

6. 1 Scope

Section 6 presents a summary of all data that e i ther have a n effect on AGC

programming o r a r e required f o r simulation and verification of AGC programs.

not be found in the memory explicitly as defined. units corrected, o r combined with other data in the most convenient and/or

ec onom ic a1 fashion.

Numerical values a r e recorded in the most widely accepted units and may These values are often rescaled,

Apollo mission and vehicle data f o r Flight AS-501 have been collected under

Apollo Mission Data (Sec. 6. 2) establ ishes the outline of the miss ion in t e r m s the following he ad ing s:

of t ra jec tor ies , profiles, etc. verification of computer programs.

This information is required f o r simulation and

AGC Memory Data (Sec. 6. 3) contains mission and vehicle dependent data that are writ ten directly into the memory of the AGC. r e f e r r e d to in Sections 3 , 4 , and 5.

p r imar i ly f o r s torage of computational variables. change during flight have been assigned to the fixed section of the memory.

before the launch date.

Other memory data are The l imited e rasable section is r e se rved

Those pa rame te r s that do not

Exceptions have been made f o r data that w i l l not be available until short ly

Spacecraft Vehicle Data (Sec. 6.4) includes configuration, m a s s propert ies ,

propulsion, and aerodynamics data. With f e w exceptions, this information will not appear directly in the AGC program. and program verification.

These data will mainly be used f o r simulation

Physical Constants (Sec. 6.5) will be used directly in the AGC programs as w e l l as program verification. kilogram, meter , and centisecond (10 sec) . Conversion to other units is accomplished by use of the fac tors defined in th i s section.

The AGC is programmed in the met r ic set of - 2

6-1 (Rev.1 - 12/66)

6.2 Apollo Mission Data

6. 2. 1 Mission Tra jec tory

Nominal mission profile Fig. 6. 1, 6. 2

Nominal Saturn boost profile Fig. 6. 3, 6. 4

Sequence of events

The information in th i s section is taken and compiled from:

Table 6. 1

(1) boost phase; NASA/(MSC & MSFC) Tra jec tory Document #66-FMP-2 “AS-501/CSM-017 Joint Reference Tra jec tory” dated 10 May 1966.

( 2 ) spacecraft; NASA/MSC Simulation Run ”Nominal AS-501 Mission Pro- file with AS-202 Guidance Logic ” Run Date 10 September 1966.

6. 2. 2 Nominal CSM/SIVB Separation Attitude Conditions

Xsc ax is

Ysc axis

Zsc axis

Roll ra te 0 degree/ second Pitch r a t e 0 degreelsecond Yaw rate 0 degreelsecond Dispersions ( 3 Sigma ) f o r Nominal Separation Attitude Conditions Xsc axis attitude 2 degrees

Ysc axis attitude 2 degrees

Zsc &.xis attitude 2 degrees

Rol l r a t e res idual 0.2 degreelsecond Pitch r a t e res idual 0. 2 degreelsecond Yaw rate residual 0 .2 degreelsecond

in plane of the t ra jectory, and in the direction of the forward horizontal

is along the momentum vector, R X V

points up, and para l le l t o the geocentric radius vector

- -

6. 2. 3

6. 2.4. SrVB Engine Shutdown Trans ien ts

Thrust decay f rom 10070 to 570 ra ted thrus t

Thrust decay f rom 570 ra ted to ze ro thrus t Cutoff impulse f rom mainstage cutoff t o 5 -percent th rus t is derived

Fig. 6. 5

Fig. 6. 6

by multiplying the thrus t level (at engine cutoff signal) by 0. 224 second.

The deviation about the cutoff impulse is the thrus t level (at the cutoff signal) t imes + - 0.030 second. derived by multiplying the thrus t level (at engine cutoff signal) by 0.0235

second.

Data Exchange Program, MSC-S-10 submitted 23 August 1966.

The cutoff impulse f rom 5-percent th rus t is

The information contained in section 6. 2. 4 is taken f rom the GN&C

6 -2 (Rev.1 - 12/66)

I

6 - 3

(Rev. 1 - 12/66)

I I I I I 1 0 - 0 hl

0 0 9 m 0 0

9 v)

6 -4 (Rev. 1 - 1 2 / 6 6 )

TIME AFTER LIFTOFF - SECONDS

Fig. 6. 3 Longitude, Latitude, Altitude, and Range During Boost t o Parking Orbit

38

37

36

35

I 34 p:

0 Z

t-

v) w w

8 33 w n

n

32 I=

I

u1

3 c

4

31

30

!9

1

52

56

50

54

I- ln

58 ln LU Lu (IL

2 72 0

n

'6 2

I

w

3 c -

9

IO

14

18

'2

6 -5 (Rev. 1 - 12/66)

INERTIAL VELOCITY / INERTIAL FLIGHT PATH ANGLE

I I INERTIAL AZIMUTH /

- !! 1

I I I I I I I 100 200 300 400 500 600 700

TIME AFTER LIFTOFF - SECONDS

06

02

98 v) w w (IC

W (3

94 u 2

0 Z a

90 E I a

86 2 2

z

3

- N

I- o? w

8 2

78

74

Fig. 6. 4 Inertial Azimuth, Inertial Ve loc i ty , and Inertial Pa th Angle During Boost to Parking Orbit

6 - 6 (Rev. 1 - 1 2 / 6 6 )

I

$2 I

m CD

G 0

V a, * a, a, *

.r( w

0 v) 0 rl

cu 0

2 0 m a .rl d 4

w a,

a, k a, cu a, k k a, s V m

x

E 5 a,

a, > 0 P cd -0 a, k 3

a, : E a, k cd m a, -0 3 .s 2 .. w b 0 z

o o o o o o m I 0 0 0 0 0 0 b m N

c- W 0

0 L n

m m M

0 m 4 m d-

0 0 0 0

0. d. C D N d - N

m c-

d- 0-

rl

1 0 0 I O 0

I 0. d 0 0 o m 0- m- O N v

3 0

3 0 d-

3 0 c- N r- N

m 4

ow El 0 0

m o W d

m o w

0 W W co N M d

c- a, 0 co

c- 0 N

m cu 0-

N

a m 0

R O4

z m m

0

m co N d

a 0 m

a 0

CO R

0 0

R 0 v M CD

4 m

z 0 d W CD

N m

Z R 0 0 c o c o O N c o c o odd N N

n a, a, k M a, -0 Y

a, a u3 M m c- N

c- 0 co

d 4 a, m

m 0

m cu

m CO

d m

Ln M

N rl

co- i cu cu 4

d- d

Y 0- c!

co v

M co CD CO cu

4 4

c 0 . .r( Z G

6 -7

I I I I I I I I I

8

6 -8 (Rev. 1 . - 1 2 / 6 6 )

+ [I) I L c b

3 a W rn

W

W

6-9 ( R e v . l - 1 2 / 6 6 )

6. 3 CMC Memory Data

6. 3. 1 Prelaunch

.

6. 3. 1. 1 Launch Pad #39A Memory

Geodetic latitude E Longitude E Geocentric radius E

Geo-entric radius t o G&N - F i s c h e r Ellipsoid radius E

Geoidal separation(ht. of MSL above ellipsoid) - Attitude of pad above MSL - Inertial reference plane azimuth E

Opticai targei H I : azimuth E elevation E

Optical target #2: azimuth E elevation E

The geophysical data in th i s section are taken f rom NASA document M -DE -8020-008B, "Natural Environment and Physical Standards f o r the

Apollo Program", Apr i l 1965.

6. 3. 1. 2 Cold Soak Attitude

The cold soak attitude will be a fixed attitude with respect t o the

stable member. prelaunch erasable memory load.

The desired attitude is defined as th ree angles included in

Angle X E 37. 81762010° Angle Y E -108. 9135267'

Angle Z E - 5 . 050401772°

The above data are taken f rom MIT/IL Record of Change F o r m #501-8.

6. 3. 1 . 3 Prelaunch Erasab le Memory Load Table 6. 2

The following is a list of the t e r m s , t he i r definitions, SOLRUM 5 5

addresses , scale f ac to r s and units, decimal values, ard oc ta l equivalents f o r the erasable data load f o r digital simulations. will undoubtedly change before flight t ime. (Table is dated 26 October 3966)

Some of these t e r m s

~

Value

28' 38' 50.93" N

279'21' 51. 93" E 6373283 m e t e r s

not available not available 0 m e t e r

not available 72OE of N

291. O O o

-015.02' 253.00'

-014. 90'

6-10 (Rev.1 - 12/66)

* J. c * u * o *

a *

* * c * *

w , 3 * -I* u * > *

” -

f- N m 10 f- n

0. 0. *

n r - J 0

N O 0 N 0 I O

U u u W Ul w

0 0 N N N

. a

a

c. CI 0 m

U U1 0

c. 0. N

I

U W

*

a

El 0

4 4

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c. Iv 0 t

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n N Q

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10 4 r- 0

W I

U w 0

N

U W 0

N 1

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U w

u W 0 0

U 2 2 W

I * c In - 3 3 c - 1 2 2 0 0 W w > U w w

N C I -

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r. 2 1 4 2 3 I O f ’

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k a * -I*

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10

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m n

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a- n

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4 - m w

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6-11 (Rev.1 - 12/66)

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w w 0

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2 0 I 2 e C - c s r.

6 - 1 2 (Rev. 1 - 1 2 / 6 6 )

0 0

t m N

Q Q

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r.

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c.

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n

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m

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CI

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v) U W ul I

c z w U

N

n

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c 3

a

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0 z -I

I < v)

a

u

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a

a

c e. n

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L

m " - e m a 0 U

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t

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r.

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v)

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m

a

a

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v u W W 0 0 N N

ul

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a

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N

N * n n n n c.c(

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m a 0 0 Y Y

N + c c

6 - 1 3 (Rev. 1 - 1 2 / 6 6 )

0 0 c\

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0 0 0 N t-

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U W U1 1 " c 2 W U

0 N N

r- 9 "

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c z u -1 -1s u - Y I Z

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c

m

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r- r- r- c c)

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a

b (1 v

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N Y,

Y) a

N + * W U

N

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c 2 0 m a z 0

c * c 0

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v) 0 In

L.

N W U

c m 3 a I c

-I a c - .- L e

n W c < U u W c z - CI m m a

c 0 c

c ul 3

I c d 4

e

z

0 W c a m- 0 w c z N v)

m

a

- - -

- a

N

c c a

z 3 a m N VI a v1

w c 3

0 U

w a

0 c

I v c

a

" a m

I U c D m N 4A

-

LL W a L Y

c LL

Y 0 0 N

w U

0 Y

0

I

a

c

U c - 3 m

I il e I Y,

Y W K

N .c Ln

c e

a

d m m c - W

W P

t3

I

a

a

c

U c w a VI

c1 Y

m 0 0 w a

a w A (3 I *

W c < n a 3

c . z 0

P

c m r.

W c u P 0

U 0 v I

0

a

I

c W I - c

w f a a c

c.

W

f c 2 a

6-14 ( R e v . 1 - 1 2 / 6 6 )

!!!!

r- r- 8 9

U W VI c. c 1: W u \ a W c y 1

b.9 L O 0 - w W P I de L L

e m - . 0

6-15 (Rev. 1 - 12/66)

6. 3, 2 Saturn Launch Vehicle Boost Phase Memory Value

Interval f rom liftoff t o LET jet t ison (assumed complete) E 184.0 seconds

Interval f rom liftoff t o s t a r t of ro l l maneuver E 8.1 seconds Duration of rol l maneuver E 1 8 . 0 seconds Total ro l l maneuver angle E 1 8 . 0 degrees

Maximum rol l maneuver rate E 1 degree/second

Interval f rom liftoff t o start of pitch maneuver Duration of pitch maneuver E 131.85 seconds

E -4.728695766 x Pitch polynom inal coeff ic ie nt s : E +4.874580890 x

E +8. 297180312 x 10-l ' E +8. 338368159 x E -1.745826987 x

E +l. 249994672 x E -3. 118989855 x

E 13 seconds

A O A 1 A 2 A 3

A 4 A 5 A6

NOTE 1 The form of the pitch polynomial is:

6 9 = Antn

n = 0

X - axis, in degrees .

where 0 = angle between iner t ia l horizontal at launch and the vehicle

t = t ime in seconds when t = 0 at liftoff +10 seconds.

6-16 (Rev. 1 - 12/66)

6. 3. 3 Attitude Maneuver Memory Data

Attitude maneuver constants wil l be found in the CSM guidance equation section of this document.

fixed memory.

These referenced values a r e found in t he

6. 3.4 Thrus t Vector Control (TVC) Memory Data

R e f e r to CSM guidance equation section of th i s document. These referenced values a r e found in the fixed memory.(For nominal and abort

missions. )

6.3.5 Programmed Time Delays The prese t delays between events are outlined in the mission logic

and t imeline section of th i s document. memory and o thers in e rasable memory. underlined.

Some of these delays a r e in fixed Erasab le memory delays a r e

6. 3. 6 Guidance and Navigation Constants

The constants used in the guidance and navigation equations are presented in Section 5 of th i s document.

portion of the memory.

6. 3.7 Re-Entry Memory Data

These constants are in the fixed

CSM attitude f o r CM/SM separation:

Xsc - axis is above the velocity vector by 60°

Ysc - axis along the momentum vector R X V - - - axis a b m e the velocity vector sc

CM Re-Entry Attitude:

Xsc - axis is above the velocity vector by 158'

Ysc - axis is along the momentum vector R X V

Zsc - axis above the velocity vector

(Assume a lift-vector up attitude)

- -

T r i m angle of a t tack 2 2 O

Nominal recovery point: Geodetic latitude 30.2075 ON

Long it ude 161. 103 OW

A complete l ist ing of (1) computer var iables in the CMC erasable memory, and (2) constants and gains in the CMC fixed memory may be found in the CM re-ent ry guidance equation section of th i s document.

6-17 (Rev.1 - 12/66)

6. 4 Spacecraft Vehicle Data

6. 4. 1 Apollo Vehicle Coordinate Reference System

Spacecraft CSM -0 17 reference dimensions Fig. 6. 7

The above figure is taken f r o m TRW Systems Document #2131-H009-R8-000 "Apollo Mission Data Specification D" dated 15 August 1966.

S M RCS, SPS, and fue l tank configuration Fig. 6. 8

Source unknown

6. 4. 1. 1 Specific Station Locations

RCS jet t h r u s t e r locations and vec to r s

The information in t h i s table is compiled f rom the GN&C Data R m - h a n g e Program, XAA-S-22 submitted 7 Apri l 1965.

Table 6. 3

SPS fue l and oxidizer tank dimensions and locations

The information in th i s table is taken f r o m the GN&C Data Exchange

Table 6 . 4

Program, NAA-S-68 submitted 11 March 1066.

SPS engine gimbal plane

XA location 833. 20 inches

YA location 0 . 0 inches

ZA location 0 . 0 inches

This information is taken f r o m the GK&C Data Exchange Program, NAA-S-68 submitted 11 Llarch 1966.

6-18 (Rev.1 - 12/66)

h4

t 4.00 RAD NOSE

LAUNCH ESCAPE SUBSYSTEM

XA 1141.25 (THEORETICAL CONE APEX)

--.-- XA 1133.5 (CM NOSE)

XA 1000 (CM HEAT SHIELD BOND LINE)

XA 833.2(ENGINE GIMBAL PLANE)

100.5 DIAMETER

SPACECRAFT LEM ADAPTER

XE 200.0 (ASCENT STAGE BASE)

XA 585.21 (LEM ATTACHING PLANE)

4 SEPARATION PLANE

FIELD SPLICE

XA APOLLO SPACECRAFT (CSM) COORDIMTE SYSTEM XE LEM COORDINATE SYSTEM Xc C M COORDINATE SYSTEM XL LES COORDINATE SYSTEM ALL LINEAR DIMENSIONS ARE IN INCHES

Fig. 6 . 7 Spacecraft CSM-017 Reference Dimensions 6-19

(Rev.1 - 12/66)

SEXTANT I x 50. SEXTANT H 8 70. SEXTANT P : 60° SEXTANT H = SOo SEXTANT P = 70. SEXTAkTZC 60'

QUAD A

'I

VIEW LOOKING

R E AC TI ON

ENGINE -E ( 4 PER QUADS,

4 QUADS)

SERVICE PROPULSION

SYSTEM

- Y

RCS JET THRUSTER RADIAL DIST. : 83.56"

E N G I N E GIMBAL PLAN E

- 154.00 I

V I E W A A

Fig. 6 . 8 CSM-017 RCS, SPS and Fuel Tank Configuration

6 - 2 0 (Rev.1 - 12/66)

10 k 0 0 c,

$ -0 C cd m C 0

cd .A c,

:

M

(D

C D C D C D C D lnlnlnln

m m t o m 0000 a m a m Q ) W c O Q ) I nmrnm m a a m

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c, 6. Q) C 0 a 0 u

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$

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c,

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-

T oh oh 0- oh o 0- on 0 0 0 0 0 0 0 0 rl r l r l rl rl rl d , d Y W W Y Y Y Y Y

m m m m m m m m o o O o o o o o 0 0 0 0 0 0 0 0 0 0 . 0 0 0 0 0 0

I I I 1

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9

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a,

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m cd -0 Q)

0 Q) a rn a, k cd 0 k 0 0 a, k- m 3 k

s

s

5

+

+.'

5 i k

6 - 2 1 (Rev. 1 - 12 /66)

rn C 0 .+ r cd V 0 ci -u C (d

rn C 0 rn C a,

.+

E b"

F

8

2

s c

&I a, N ..-I

2

a C rd

a, 4

rn UI

* a a,

PI

2 F

0 0 0 0

N N N m

N m m cc) W co co co

m m W W CD a m In W W W W co W m m m m

In In In In

m m PJ N N @3 N N

24 C F a 3 rn k a, N

E

.d

2 8 U U

0

N

II

0

cd k al &

.+ +.'

fi z *

6 - 2 2 ( R e v . 1 - 1 2 / 6 6 )

6. 4. 2 Apollo Vehicle Mass Property Data

CSM spacecraf t m a s s propert ies s u m m a r y Table 6. 5 SM vehicle m a s s propert ies summary Table 6. 6

CSM spacecraf t propellant loading summary Table 6.7 SM-SPS usab le propellant data Table 6.8 The information in th i s section is taken f r o m TRW Sys tems Document

#2131-H009-R8-000, "Apollo Mission Data Specification D AS-501" dated 15 August 1966.

6. 4. 3 Apollo Vehicle Dynamic Data

6.4. 3. 1 Slosh - Mixture Ratio of 2. Of 1 .0

Oxidizer s losh frequency Fuel s losh frequency

Oxidizer equivalent s losh m a s s F u e l equivalent s losh mass

Oxidizer s losh C. G. X-location Fue l s losh C. G. X-location Oxidizer slosh C. G. Y-location Fue l s losh C. G. Y-location Oxidizer s losh C. G. Z -location Fue l s losh C. G. Z-location Oxidizer s losh damping ration

F u e l s losh damping rat ion

3.82 rad lsecond 4. 07 r ad / second

13.7 s lugs 970 t o 840 inches

974 to 840 inches

-48.3 to 48. 3 inches 14. 8 t o -14. 8 inches -6. 6 t o 6.6 inahes

47.8 t o -47. 8 inches 0.005

0.005

44.55 s lugs

The X-locations of the sloshing m a s s e s were obtained by eyeballing the SM diagram in NAA XTASI entry 13 page 3, together with XTASI

24 page 8.

P rogram, NAA-S-68 submitted 11 March 1966. Allowance hasbeen made f o r the fact that fue l and ozidizer tank assignments differ between Block I and Block 11.

The Y- and Z - locations are taken f r o m the GN&C Data Exchange

6-23 (Rev.1 - 12/66)

h

-.I v x

k rd E E 3 v) ro a,

k Q) a 0

.rl c1

c 2 m m

c, w rd k u a, V rd a v)

0 In I

4

Y u3

CD a,

2 H ;

c

* n m . g * m

4 8. " W "

c c

m In

- N O 0 . 3 - 0 . - r-

N

0 . r - O N O 0 . I ) u o . o 9 n a o 0

- 7 P l m N N N N - N

. . . . .

- m m m m

0 0 0 d 3 t l l l u u u 9 m 0 0 5

o m - - . . . . I I ,

g

m r- 0

N

o. r- 9

r-

N m N

0- - 0. 0

9

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m 9

9. 5 .n

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rn

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N 3

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-7

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6 -24 (Rev.l - 12/66)

a, .rl c, k

&

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co co a,

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6 -25 ( R e v . 1 - 1 2 / 6 6 }

0 . 0 . -

N 9 N 0 . 0

P N o N

N O N VI -

- V I - e

.D c VI N VI

m c I m a N

c 0 m * V I *

m m o m

c r- N 0 m

0 >

3

6 -26 (Rev.l - 12/66)

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d d d d d d d d d d d d d d d d d d d d d d d d d d d d

oooooooooooooooooooodOOOOooooooo d d d d d d d d d d d d d d d d d d d d d d d d d d d d

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d d d d d d d d ~ d d d d d d d d d d d d d d d d d d d I I I I I 1 I l l 1 I I I I I I

6 -27 (Rev.1 - 12/66)

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4 1 1 1 1 1 1 1 1 1 1 1 1 l 1 1 1 1 1 I I I I I d d d d d d d d d d d d d d d d d d ~ d d d d d "WJ 000000000000000000000000

4 1 l 1 1 1 1 1 1 1 1 1 l 1 l 1 1 1 1 1 1 1 1 1 1 d d d d d d d d d d d d d d d d d d d d d ~ d d 4

6 -28 (Rev.l - 12/66)

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6 -29 (Rev.1 - 12/66)

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d d d d d d d d d d d d d d d d d d d d d

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d d d d d d d d d d d d d d d d d d d d d

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d d d d d d d d d d d d d d d d d d d d d

2 2 2 4- 2 2 -= 2 2 NN NN ni N 6 $ "'

300000000000000000000

I+ d d d d d d d d d d d d d d d d d d d d d ;i 9 , I I I I

6 -30 (Rev.l - 12/66)

k a, N 2 .d x 0 H

3 V 0

x s 4 2

d rd

u a 0

d d

ca CD ci d n rd b

6-31 (Rev. l - 12/66)

A

* P

t2 v,

co

o o o o o o o o o o o a o o o o o o o o o

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

n q ~ $ o o o o o o o o o o o o o d o o o o o

o o o o o o o o o o o o o o a o o o o o o

I O J , I I I I , I , I I I , I I

6 -32 (Rev.1 - 12/66)

al M Id k 0 x k Q) N

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

d d d d d d d d d d d d d d d d d d d d d d O O d - c 1 -4 d .d Y

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

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~ 4 4 d d d d O d d d d d d d d d d d d d d d O ~ ~ I I J I J J I I I 1 I 1 , I I L I I L I I

52 0 c a

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I-i

6 -33 (Rev.l - 12/66)

k

k 3 N .I

v .I - 5 c-(

Y

x P

I

6 -34 (Rev.l - 12/66)

6.4.4 Service Propulsion System (SPS) Data

6.4.4. 1. Engine Physical Proper t ies

Mass 25 s lugs

The source fo r the above datum is unofficial, te lecon with

Mr. Jack Pot ts S&ID/NAA. 2

Engine iner t ia

Engine c. g. t o gimbal

Th i s information is taken f rom the GN&C Data Exchange P rogram,

246. 1 s lug feet 8 inches

NAA-S-46 submitted 8 October 1965.

6. 4.4. 2 Gimbal Actuator System (MOD 11) SPS thrus t vector orientation Fig. 6. 9 The above figure is taken f rom TRW Systems Document

#2131-H009-R8-000 "Apollo Mission Data Specification D A/S-501" dated 15 August 1966.

+ 6 degrees

- 4 degrees + - 7 degrees

0 degrees + 1/8 inch

- Pitch gimbal l imit

P i tch gimbal offset Yaw gimbal l imit Yaw gimbal off set

The above information is taken f rom TRW Systems Document Thrus t - t o - gimbal offset -

#2131-H009-R8-000 "Apollo Mission Data Specification D AS-501 ' I dated 15 August 1966.

Jet damping coefficient

Hose stiffness

17 1 foot-pound/rad/sec 285 foot-pound/rad

The above information is taken f rom discussions of S&ID/MIT Meeting #66B of 17 September 1963.

Thrus t angular misalignment 0 .5 degree

The above datum was taken f rom NAA XTASI # l o . 2 A ct uat o r iner t ia IA 65 slug - feet

Actuator lag WA 6.6 rad/second

Total amplifier - clutch gain KS-KT 20(3530) f t . - lb/rad.

Torque l imit LMT 1500 ft-lb.

Slew rate l imit LMR 0. 1 rad /sec .

NOTE: Because of the inclusion of a Block I1 actuator in the AS-501 vehicle, the above actuator parameters (LMT and LMR) are speculative.

In par t icular , a worst case LMR should be sufficient t o detect anomalies

causedby the Block I1 actuator. Reference MIT/IL MDRB Record of Change F o r m # 501-18

dated 31 October 1966.

6-35 (Rev. 1 - 12/66)

v, c z - A A s

I A

m m a0 It

?

a

$

X v a

E a, + rn h rn P 1 v1

2! .r(

l-4 rn 1 a 0 k P( Q, 0 > k a, v1

.d

6 -36 (Rev.l - 12/66)

6. 4.4. 3 SPS Engine Vacuum Performance

Turn-on-off s tep t ransient

Buildup impulse to 90% rated thrust Thrus t buildup vs t ime of engine-on signal

Specific impulse 1 steady -s t a t e operation i Thrus t

Propellant flow ra t e

Thrus t decay vs t ime of engine-off signal Tail-off impulse to 10% rated thrus t

0.27 second

300 + 200 lb - second

Fig. 6. 10 311. 4 + 1.5 seconds 21,500 + 215 lbs.

69.0 lb f second

Fig. 6 . 10 10,800 + 1200 lb. -sec.

-

- -

- The information contained in this section is taken f r o m TRW Sys tems

Document #2131-H009-R8-000, "Apollo Mission Data Specification D AS-501" dated 15 August 1966.

Thrus t Vector Control Autopilot block diagram F ig 6. 11

Reference MIT/IL MDRB Record of Change F o r m #501-18 dated

31 October 1966.

6-37 (Rev . l - 12/66)

h k (d

E . .

g W >

6 -38 (Rev.1 - 12/66)

A

4

..

r tx I-

-++-=! + 5

6-39 (Rev.l - 12/66)

6. 4.5 CSM Reaction Control System (RCS)

6. 4. 5. 1 RCS Jet Physical P rope r t i e s

RCS thrus t chamber configuration Fig. 6. 12

Th i s figure is taken f r o m TRW Systems Document #2131-H009- R8-000, "Apollo Mission Data Specification D AS-501", 15 August 1966.

Off set angle Cant angle Thrus t rad ia l a r m

7. 25 deg rees

10.0 degrees 83. 5596 inches

The above data are taken f r o m the GN&C Data Exchange Program,

NAA-S-22 submitted 7 April 1965.

6. 4.5.2 Jet Vacuum Performance

v s electrical pulse width Fig. 6. 13, 6. 14

steady- state operation Fig. 6. 14

Total impulse Propellant consumed

Specific impulse Thrus t

Specific impulse Propellant flow rate Thrust buildup t r ans i en t s Thrus t decay t r ans i en t s

not available not available

1 The f igures referenced in this section are taken f r o m TRW Systems

Document #2131 -H009-R8-000, "Apollo Mission Data Specification D AS-501" dated 15 August 1966.

RCS Autopilot block diagram Fig. 6. 15

Reference MIT/IL MDRB Record of Change F o r m #501-18 dated 31 October 1966.

6 -40 (Rev. 1 - 12/66)

I I , -‘A

VIEW 1

(TYP .)

LOCATION OF SWRCS CLUSTERS LOOKING FORWARD INTO SM

‘ZA

(IN THE DIRECTION OF THE +XA AXIS)

Fig. 6 . 1 2 SM/RCS Thrust Chamber Locations

6-41 ( R e v . 1 - 1 2 / 6 6 )

10’ (TYP .) I-

0.04

h m -1 v

p 0.03 z 2 0 U

5 0.02 W L s a /I 0.01 %

0.00

-

-

- h

g - v)

W v) 2 -

v

I I I I I I 1 1

G -2 .00 2

W w

1.80 2 x z

1.60 d 2

1.40 2 ci w N X 0

1 .oo 'I

1.20

2 80

260

240

220

200

180

160

140

120

100

10 20 30 40 50 60 70 80 90 100

PULSE WIDTH (MS)

Note: Data are high, low, and average values High and low values resulting from a large number of qualification tests.

shall be used as 30 values.

Fig. 6 . 1 3 SM/RCS Vacuum P e r f o r m a n c e Data for Pulse Wid ths Less than 100 ms

6 -42 ( R e v . l - 1 2 / 6 6 )

v)

w v) -1

Y

3 n

u LL

n

r Y v)

I1 eL m -

0 . 4 ~

0.3

0.2

0.1

0.0-

300

260

240 -

- h

Y v) I

m 2

w m -1

-- 3 n z 2 -1

100

90

80

70

60

50

40

30

20

10

t r

c / STEADY STATE PERFORMANCE

(PULSE WIDTH > 1 .O SEC) - . AVE

LOW 98.0 275.6 0.352 1.979

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 .o PULSE WIDTH (SEC)

N o t e : Data are high, low, and avera e values resulting from a large number of qualification tests. High andyow values shal I be used as 3y values.

Fig. 6.14 SM/RCS Vacuum Performance Data for Pulse . Widths Greater than 100 ms

6 -43 (Rev.l - 12/66)

i

6 -44 (Rev. 1 - 12/66)

6.4. 6 CM Vehicle Data

6. 4. 6. 1 Apollo C M Coordinate Reference System

Spacecraft CSM-017 reference dimensions CM axes and notation system

Fig. 6. 7 Fig. 6. 16

The f igures shown above a re taken from TRW Sys tems Document

#2131 -H009-R8-000, "Apollo Mission Data Specification D AS-501", dated 15 August 1966.

6.4. 6. 2 Specific Station Locations

IMU location 1056. 6 inches

T h i s number is taken from NAA-MIT/IL ICD MH01-01301-116

RCS je t t h r u s t e r locations Table 6. 9

T h i s table is taken from TRW Sys tems Document #2131-H009-R8-000, "Apollo Mission Data Specification D AS-501", dated 15 August 1966.

6.4. 6. 3 Apollo CM Mass Proper ty Data

CM m a s s propert ies summary CM sequential m a s s propert ies

Table 6. 10 Table 6. 11

The data contained in this section are taken f rom TRW Systems Document #2131 -H009-R8-000, "Apollo Mission Data Specification D AS-501", dated 15 August 1966.

Thrus t Chamber

5

7 8 6 3 1

10,11

9,12 2 4

Table 6. 9 RCS Je t Thrus te r Locations

C, of Thrust Exit P lane on Outer ML

A

B C D E F

(I)G

(1I)H J K

L1

xC

27.6750

27.6750 27.6750

27.6750 27.6750 2 7 . 6750 32.3000

32.3000 85.5000

91.1250

yC

72.2846

72,2846 -72. 2846 -72. 2846

-4.4211 4.4211

-51. 9826

51. 9826 0

0

zC

-4.4211 4.4211

4.4211 -4.4211

-72. 2846 -72.2846 -50.3454

-50.3454

-35. 5483 -31. 9616

RC

72.4197

72.4197 72.4197

72.4197 72.4197 72.4197 72.3661

72.3661

35.5483 31.9616

6-45 (Rev.1 - 12/66)

+c N

cos a7 = cos a COS P tan = tan D

sin a

CA AXIAL FORCE COEFFICIENT (BODY AXIS), AXIAL FORCE/qmS

ROLLING MOMENT COEFFICIENT ABOUT C G (BODY AXIS), ROLLING MOMENT/kSd ROLLING MOMENT COEFFICIENT ABOUT THEORETICAL C O N E APEX (BODY AXIS), ROLLING MOMENT/q=Sd PITCHING MOMENT COEFFICIENT ABOUT CG, PITCHING MOMENT/q,Sd PITCHING MOMENT COEFFICIENT ABOUT THEORETICAL C O N E APEX, PITCHING MOMENT/qapSd

NORMAL FORCE COEFFICIENT (BODY AXIS), NORMAL FORCVqmS YAWING MOMENT COEFFICIENT (BODY AXIS), YAWING MOMENT/qSd

S13E FORCE COEFFICIENT (BODY AXIS), SIDE FORCE/qS

ctA

MA

C N

‘n

Cy ‘M “M.

‘n, + ‘“6

CM C

PITCH DAMPING COEFFICIENT , PER RADIAN

YAW DAMPING COEFFICIENT , PER RADIAN q a

ia ANGLE OF ATTACK, DEGREES aT

TOTAL ANGLE OF ATTACK, DEGREES

9 ANGLE OF SIDESLIP, DEGREES ROLL ANGLE DEGREES REFERENCE L ~ N G T H = 154 INCHES REFERENCE AREA = 129.35 SQUARE FEET FREESTREAM VELOCITY, FEET PER SECOND DYNAMIC PRESSURE, POUNDS PER SQUARE FOOT

@ d 5 V q, 9 PITCHING RATE

Fig. 6. 16 Command Module Axes, Aerodynamic Coefficient, and Notation System

6 -46 (Rev.1 - 1 2 / 6 6 )

cn Q) .rl c, k

0 !L

u M

0 v) I

2 0 v-l

0 N

N ic -

0.

0 N ci

f

- 0 ;c

f

t f

+ -r

* 0

x - - -1. 0 -

0

0 - -

6 -47 (Rev.l - 12/66)

m

k a m

6 -48 (Rev.1 - 12 /66)

6.4.6.4 CM Re-Entry Aerodynamic Data

Reference area Ref ere nce d iameter

Heat shield cant Moment re ference center: X-component

Y - c om pone nt

Z -component

A e rod y nam ic coefficients Lift and drag charac te r i s t ics f o r M = 6+25 T r i m L/D v s C. G. locations

2 129. 4 feet 145 inches

0. 1300 degree

1141.25 inches

0. 0 inch 0. 0 inch

Table 6. 12 Fig. 6. 17

Fig. 6. 18

The information contained in th i s section is taken f rom TRW Sys tems Document # 2131-H009-R8-000, "Apollo Mission Data Specification D AS-501. 'I

dated 15 August 1966 as amended by United States Goverment Memorandum P M 3/M-170/66 dated 14 November 1966.

6 -49 (Rev. 1 - 12/66)

TABLE 1

A L P H A

110. 1300 115. 1300 120. 1300 125.1300 130. 1300 135. 1300 140. 1300 145. 1300 150. 1300 155.1300 160. 1300 165. 1300 170.1300 175. 1300 180.1300 185.1300 190. 1300

110.1300 115. 1300 120.1300 125. 1300 130.1300 135. 1300 140.1300 145. 1300 150.1300 155.1300 160.1300 165.1300 170. 1300 175. 1300 180. 1300 185.1300 190. 1300

110. 1300 115. 1300 120.1300 125. 1300 130. 1300 135.1300 140.1300 145.1300 150. 1300 155. 1300 160.1300 165. 1300 170. 1300 175. 1300 180.1300 185. 1300 190.1300

Command Module - Spacecraft 017

Aerodynamic Coefficients

c M cN M = 0.4

-0.1021 0.2359 -0.0866 0.2186 -0.0590 0.1902 -0.0289 0.1529 0.0037 0. 1076 0.0392 0.0554 0.0770 -0.0018 0.1068 -0.0509 0.1210 -0.0810 0.1186 -0.0921 0.0950 -0.0782 0.0630 -0.0531 0.0333 -0.0331 0.0068 -0. 0181

-0.0198 -0.003 1 -0.0467 0. 0129 -0.0739 0.0279

M = 0.7

-0.1455 -0. 1070 -0.0690 -0.0377 -0.0085 0.0177 0.0383 0.0432 0.0267 0.0116 0.0097 0.0208 0.0304 0.0242

-0.0096 -0.0357 -0.0443

0. 3139 0. 2775 0. 2391 0.1988 0. 1544 0.1111 0.0700 0.0478 0.0518 0.0497 0.0287

-0.0084 -0.0424 -0. 0524 -0. 0223 -0. 0003 -0.0034

M = 0.9

-0.2519 0.4947 -0.1902 0.4213 -0.1357 0. 3529 -0.0918 0.2845 -0.0600 0.2232 -0.0288 0. 1630 -0.0087 0.1168 -0.0076 0.0976 -0.0167 0.0935 -0.0224 0.0844 -0.0057 0.0424

0.0094 0.0024 0.0178 -0.0276 0.0145 -0. 0406

-0.0046 -0.0276 -0.0207 -0.0156 -0.0217 -0.0226

cA

-0. 0405 -0. 1605 -0. 3604 -0. 5003 -0. 6102 -0. 7001 -0.7800 -0. 8449 -0.8998 -0. 9398 -0. 9498 -0.9449 -0.9299 -0. 9200 -0. 9150 -0. 9170 -0. 9201

-0.0607 -0. 2006 -0. 3805 -0. 5505 -0. 7004 -0. 8203 -0. 8902 -0.9501 -0. 9851 -1.0101 -1.0281 -1.0380 -1. 0399 -1, 0379 -1.0200 -1.0300 -1. 0500

-0.1111 -0. 2960 -0.4808 -0. 6406 -0. 7805 -0.8904 -0.9753 -1.0402 -1,0902 -1. 1302 -1. 1501 -1. 1600 -1. 1549 -1. 1349 -1. 1249 -1. 1350 -1. 1550

6-50

(Rev. 1 - 1 2 / 6 6 )

TABLE 1 (Cont'd)

ALPHA

110. 1300 115. 1300 120.1300 125. 1300 130.1300 135.1300 140. 1300 145.1300 150.1300 155.1300 160.1300 165. 1300 170.1300 175. 1300 180.1300 185. 1300 190.1300

110.1300 115.1300 120.1300 125.1300 130.1300 135. 1300 140.1300 145. 1300 150.1300 155.1300 160.1300 165.1300 170.1300 175. 1300 180.1300 185.1300 190.1300

Command Module - Spacecraft 017

A e rodynamic Coefficients

c M

-0.2942 -0.2548 -0. 2075 -0. 1529 -0. 1046 -0.0561 -0.0083 0.0169 0.0136 0.0083 0.0046 0.0027 0.0046 0.0069

-0.0078 -0.0192 -0.0082

CN

M = 1.1

0.5113 0.4670 0.4126 0. 3502 0.2798 0.2025 0.1173 0.0582 0.0421 0.0311 0.0190 0.0050

-0.0141 -0.0330 -0.0250 -0.01 60 -0.0321

M = 1.2

-0. 3004 0. 5093 -0.2591 0.4629 -0.2103 0.4056 -0.1537 0. 3382 -0.0932 0. 2608 -0.0384 0. 1725 -0.0016 0. 1023

0.0186 0.0492 0.0209 0.0232 0.0143 0.0121 0.0100 0.0031 0.0126 -0.0120 0.0161 -0.0280 0.0138 -0.0340 0.0101 -0.0381 0.0082 -0.0411 0.0064 -0.0430

M = 1.35

110.1300 115.1300 120.1300 125.1300 130.1300 135.1300 140.1300 145.1300 150.1300 155.1300 160.1300 165.1300 170.1300 175.1300 180.1300 185.1300 190.1300

-0.3589 -0.3202 -0.2534 -0.1862 -0.1179 -0.0579 -0.0096

0.0272 0.0415 0.0428 0.0361 0.0270 0.0189 0.01 51 0.0126 0.01 14 0.0102

0.5994 0.5590 0.4776 0.3922 0.3038 0.2215 0.1422 0.0779 0.0388 0.0148 0.0008

-0.0113 -0.0213 -0.0303 -0.0383 -0.0453 -0.0523

cA

-0. 2962 -0.4411 -0.6009 -0. 7908 -0.9506 -1.1005 -1.2003 -1.2401 -1.2751 -1. 3001 -1.3250 -1. 3400 -1.3450 -1.3399 -1.3379 -1. 3350 -1.3499

-0.3112 -0.4711 -0.6309 -0.8058 -0.9706 -1. 1004 -1. 1902 -1.2401 -1.2551 -1.2750 -1.3000 -1.3200 -1.3349 -1.3399 -1.3449 -1.3449 -1.3349

-0.2864 -0.4563 -0. 6311 -0.8009 -0.9707 -1.1205 -1.2303 -1.3452 -1.4001 -1.4200 -1.4300 -1.4400 -1.4400 -1.4399 -1.4399 -1,4399 -1.4399

6-51

(Rev. 1 - 12/66)

TABLE 1 (Cont'd)

A L P H A

110.1300 11 5. 1300 120.1300 125.1300 130. 1300 135.1300 140.1300 145. 1300 150. 1300 155. 1300 160. 1300 165.1300 170.1300 175. 1300 180.1300 185.1300 190.1300

110.1300 115.1300 120.1300 125.1300 130.1300 135.1300 140.1300 145. 1300 150.1300 155.1300 160.1300 165.1300 170.1300 175.1300 180.1300 185.1300 190.1300

110.1300 115.1300 120.1300 125.1300 130.1300 135.1300 140.1300 145.1300 150.1300 155.1300 160.1300 165.1300 170.1300 175.1300 180.1300 185.1300 190.1300

Command Module - Spacecraft 017

Aerodynamic Coefficients

c M cN

M = 1.65

-0.3294 0.5493 -0. 2976 -0.2552 -0.2103 -0. 1624 -0. 11 60 -0.0713 -0.0350 -0.0083 0.0081 0.0161 0.0150 0.0130 0.0102 0.0083 0.0078 0.0097

0. 5130 0.4696 0.4173 0. 3569 0.2926 0.2243 0. 1620 0.1089 0. 0628 0.0277 0.0067

-0.0094 -0. 0204 -0.0284 -0.0344 -0.0424

n4 = 2 . 0

-0. 3199 -0. 2776 -0.2372 -0.1954 -0. 1564 -0. 1185 -0.0837 -0.0514 -0.0241 -0.0050 0.0097 0.0195 0.0211 0.0155 0.0109 0.0063 0.0027

0. 5293 0.4860 0.4407 0. 3903 0. 3390 0.2877 0.2364 0.1822 0.1289 0. 0798 0.0367 0.0016

-0.0204 -0.0283 -0.0334 -0.0354 -0.0374

RI = 2 .4

-0. 2974 -0.2579 -0.2203 -0.1833 -0.1492 -0. 1165 -0.0852 -0.0566 -0.0336 -0.0156 -0.0046 0.0047 0.0087 0.0073 0.0058 0.0079 0.0112

0.4944 0.4541 0.4127 0. 3704 0.3281 0. 2828 0.2355 0.1872 0. 1420 0. '097 8 0.0597 0.0246

-0.0014 -0.0154 -0.0264 -0.0364 -0.0454

A

-0.31 12 -0.4612 -0. 6111 -0.7709 -0. 9308 -1.0707 -1. 1955 -1. 3004 -1.3753 -1,4251 -1.4551 -1.4700 -1.4780 -1. 4800 -1.4819 -1.4799 -1.4779

-0.291% -0.4451 -0. 5870 -0.7269 -0. 8608 -1.0007 -1.1305 -1. 2504 -1. 3503 -1.42u2 -1. 4601 -1.4800 -1. 4950 -1.4999 -1.4999 -1.4999 -1.4999

-0.281 1 -0.4160 -0.5559 -0.7008 -0.8307 -0.9606 -1.0955 -1. 2154 -1.3253 -1.4002 -1. 4451 -1.4801 -1.5050 -1. 5200 -1. 5199 -1. 5149 -1.4999

6-52

(Rev. 1 - 1 2 / 6 6 )

TABLE 1 (Cont'd)

A L P H A

110.1300 115.1300 120.1300 125. 1300 130.1300 135.1300 140.1300 145.1300 150.1300 155. 1300 160.1300 165.1300 170.1300 175.1300 180.1300 185.1300 190.1300

110.1300 115.1300 120.1300 125.1300 130.1300 135.1300 140.1300 145.1300 150.1300 155.1300 160.1300 165.1300 170.1300 175.1300 180.1300 185.1300 190.1300

110.1300 115.1300 120.1300 125.1300 130.1300 135.1300 140.1300 145.1300 150.1300 155.1300 160.1300 165.1300 170.1300 175.1300 180.1300

Command Module - Spacecraft 017

Aerodynamic Coefficients

hl c N M = 3.0

-0.2624 0.4394 -0. 2223 0.4001 -0.1840 0.3598 -0.1499 0. 3205 -0. 11 91 0. 2832 -0.0952 0.2479 -0.0694 0.2086 -0.0481 0. 1694 -0.0275 0.1291 -0. 0113 0.0929 -0.0056 0.0668 -0.0067 0.0467 -0.0112 0.0326 -0.0110 0.0146 -0.0078 -0.0034 -0.0010 -0.0214 0.0102 -0.0384

-0. 2332 -0.1891 -0.1494 -0.1147 -0.0866 -0.0647 -0.0482 -0.0345 -0.0233 -0.01 1 7 -0.0044 -0.0001

0.0015 0.0001 0.0007 0.0028 0.0098

T

-0.2150 -0.1660 -0.1229 -0.0841 -0.0597 -0.0460 -0.0383 -0.0244 -0.0154 -0.0066 -0.0017 0.0032 0.0064 0.0079 0.0087

M = 4 . 0

0.3935 0.3492 0.3069 0.2686 0.2323 0.2020 0.1737 0.1465 0.1203 0.0921 0.0659 0.0428 0.0237 0.0077

-0.0084 -0.0234 -0.0383

M = 6 4 2 5

0.3701 0.3197 0.2724 0.2241 0.1998 0.1835 0.1703 0.1440 0.1218 0.0976 0.0764 0.0543 0.0322 0.0131

-0.0019

-0.2510 -0.3909 -0.5208 -0. 6607 -0.8006 -0.91 56 -1.0505 -1. 1604 -1.2653 -1.3502 -1.4102 -1.4601 -1.4901 -1. 5100 -1. 5150 -1.5050 -1.4899

-0.2109 -0.3508 -0.4907 -0. 6256 -0.7505 -0.8805 -0.9954 -1. 1003 -1.2003 -1.2852 -1. 3552 -1.4101 -1.4501 -1.4750 -1.4850 -1.4800 -1.4599

-0.1978 -0.3407 -0.4806 -0. 6195 -0.7435 -0.8734 -0. 9894 -1.0928 -1. 1863 -1, 2772 -1.3552 -1.4238 -1. 471 1 -1.4940 -1. 5000

6-53

(Rev. 1 - 12/66)

180

160

140

120

100

80

6C

4c

2c

C

1.8

1.6

1.4 n U

n z 1.2

0

V 0.8

(3

n 2 n z Q 0.6 I- LL

--1 -

0.4

0.2

I 6

L I F T -DRAG RATIO,

Fig. 6.17 CM (AF-017,020) M = 6+25, Lift and Drag Charac ter i s t ics

6-54

(Rev. 1 - 12/66)

10.0

9.0

-

I,

LL L Lu N - 6.0

- 5.c

Y L Lu N

4.c

3.0

L / ~ trim

PRIMARY TRIM LIMIT LINE

1036 1038 1040 1042 1044 1( X C g (INCHES)

Fig. 6.18 Command Module (AF-017 ,20) L/Dtrim A s A Function of Center-of-Gravity Location Using

M = 6-25 Data. Heat Shield Cant is 0.13'.

16

6-55

(Rev. 1 - 12/66)

6. 4.6. 5 CM Reaction Control System (RCS) Data

RCS thrus t chamber configuration Fig. 6. 19

Tota l impulse Propel lant consumed vs electrical pulse width Fig. 6 . 2 0 , 6. 2 1

Specific impulse

Thrus t Specific impulse steady state operation Fig. 6. 2 1

Propel lant flow rate Thrus t buildup t rans ien ts not available Thrus t decay t rans ien ts not available

The f igures referenced in th i s sect ion are taken f r o m TRW Systems Document # 2 131 -H009 -R8-000, "Apollo Mission Data Specification

D AS-501, dated 15 August 1966.

6-56 (Rev. 1 - 12/66)

U x w

2

0, C aJ

.. v)

c a 0 z .

6-57

(Rev. 1 - 12/66)

5 2.20 w pc

2 2.00 X - x

~ 1.80 w

L 3

? 1.60

h 0.04 8 240- vl

w vl

v

5 220 - = 9 3

- u 200- 2 M 2 .

E?0.03 a.

U - 180 - 0

II U c 0.02 Z

-! 1 6 0 - 4 n

-1 d W

140- 8 a.

3 )I 0.01

0.0-

8

'

-

IO

- 9

8

- - 7 M

& 6

vl I m

W vl a - 2 5 I 2 4

- c 3

0 c - II

2

L O h

I I 1 I I I I

PULSE WIDTH (MS) Note :

Data are high, low, and average values resulting from a large number of qualification tests, with the exception of w , where only average values are available. High and low values s h a l l be used as 30 values.

F'g. 6.20 CM/RCS VQcuum Performance Data for Pulse Width Less than 100 ms.

6-58

(Rev. 1 - 12/66)

W

I- 5 2.0- x = 1.8- A W

II

Y 0

0.40-

m

e 0.30 -I Y

5

z' 0.20- 4

P

v)

Z 0 U

-1 W n

n ( 1 0.10- i3

0.00-

- 100

90

80

3 70

5 60 W v) I

Y

W v)

3 50 3 3 40 2

30

20

10

0

-

'SP L-

TEADY STATE PERFORMANCE

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 PULSE WIDTH (SEC)

Datoare high, low and average values Note:

resulting from a large number of qualification tests, with the exception of w , where only average values are available. However, high and low values of w are presented for steady state operation. High and low values shal I be used as 30 values.

Fig. 6.21 CM/RCS Vacuum P e r f o r m a n c e Data f o r Pulse Widths G r e a t e r than 100 ms.

6-59

(Rev. 1 - 1 2 / 6 6 )

I

7. G&N ERROR ANALYSIS

7. 1 Introduction

The resu l t s of a revised G&N e r r o r study fo r the 501 miss ion a r e given

This study w a s pr imar i ly concerned with the effects of IMU component here . uncertaint ies on t ra jec tory uncertainties fo r two update cases . These were:

1) Navigational update 5 minutes before injection burn ignition (2nd SIVB

burn). Referred to as update 1 in tables.

2 ) Navigational update 13 minutes before 2nd SPS burn ignition (22. 2 min of f ree- fa l l coast to 400,000 ft) . Referred to as update 2 in tables.

The c a s e where update was made 20 minutes before injection burn ignition (considered in the previous revision) w a s not included in the present study, s ince

the T case .

uncertainties for this c a s e were not markedly different f rom the update 1 ff

The simulation of the s ta te vector, or navigational R , V update, included

the effects of tracking uncertainties. Updating did not apply to the alignment of the IMU Stable Member, since the SM is not realigned during the 501 flight.

The e r r o r studies assume a prelaunch Stable Member orientation as shown i n Fig. 7 . 1. XSM is up along the local ver t ical a t launch instant, while ZsM is

horizontal down-range at the nominal azimuth. axes are col inear with Stable Member axes.

The gyro and acce lerometer input

Block 1 IMU uncertainties were assumed for these s tudies (see section 7 . 9 These were the same as those assumed for the previous revis ion of for data).

July 1966, with one important exception: 3-meru bias drift uncertainty is used in th i s e r r o r study, whereas in the previous revision 2-meru bias drift w a s inadvert-

ently used.

7. 2 Significant Results of E r r o r Study

Of p r imary concern in the e r r o r s tudies were the effects of IMU uncertain-

t i e s on:

1) Uncertainty i n computed free-fall time of flight, (U)Tff, t o the reent ry

s t a r t altitude of 400,000 feet.

2) Flight path angle uncertainty, (U)yAA, at reent ry s ta r t .

3) CEP at reentry end.

7 - 1

(Rev . l - 1 2 / 6 6 )

1 ~~~

I I 4x1

HORIZONTAL PLANE AT LAUNCH INSTANT

'5 M

X I lYI I Z I - LAUNCH INERTIAL A X E S

XSM IYw,Z5M - I M U STABLE MEMBER AXES

Note: NOMINAL AZIMUTH IS 729

Fig. 7. 1 Coordinate Axes for 501 Launch Configuration

7 -2 ( R e v . l - 1 2 / 6 6 )

~

I

Table 7.1 gives these data in summary form. A l l uncertainties r ep resen t

the combined effect of 10 Block 1 IMU uncertainties and tracking update uncertainties.

Tables 7 . 7 and 7 . 10 give detailed data on IMU component contributions t o (U)Tff. Table 7. 13 gives detailed data on contributions to (U)yAA for the two update cases. The r e a d e r should consult Section 7. 11 and Fig. 7. 2 on the definition used for flight-

path-angle uncertainty.

7 . 3 Accelerometer Inputs to AGC

F o r normal free-fall flight periods the AGC does not receive accelerat ion information from the IMU accelerometers . on AGC inputs.

Accelerometer bias has then no effect

F o r the 501 mission AGC programming s imi la r t o that for the 202 mission

w i l l be used. outputs f rom SIVB cutoff to 1st SPS burn ignition.

mission the AGC w i l l accept output AVls f rom the acce le romete r s f rom 1s t SIVB cutoff through the whole parking orbit and from 2nd SIVB burn (injection burn) cutoff f o r about 12 minutes to 1st SPS burn ignition.

For the 202 mission the AGC was left sensit ive to acce le romete r

Correspondingly, for the 501

A f t e r 1st SPS burn cutoff the AGC is then insensit ive to acce le romete r outputs until 30 seconds before 2nd SPS burn ignition (t ime of ullage s ta r t ) . The AGC then accepts accelerometer data the rest of the flight including the coast to r een t ry s ta r t .

Table 7. 2 gives indication uncertainties before update at 5 minutes before These relatively la rge uncertainties show the effect of the AGC's injection burn.

continuing sensitivity to accelerometer bias while in parking orbit. before injection burn ignition is imperative; otherwise, computed Tff uncertainties

a f t e r the burn would be excessively large.

An update

7 . 4 E r r o r Table Description

The following tables , summarizing the resu l t s of the e r r o r studies, a r e

t o b e found a t the end of this section. tables showing individual contributions by IMU component uncertainties.

Tables 7.3 through 7.15 a r e detailed e r r o r

7. 2

7. 3 7. 4 7. 5 7. 6 7. 7 T uncertainties at injection and SPSl burn cutoff

7.8

Summary of 501 flight uncertainties

SIVB cutoff indication uncertainties IMU SM misalignments at SIVB cutoff

Injection burn cutoff uncertainties (Update 1) SPSl burn cutoff uncertainties (Update 1)

ff Uncertainties at 13 mins. before SPS2 ignition (Update 1)

7 - 3

(Rev. l - 12/66)

m a,

c (d

k a,

.d Y

.d

+J

x 3 +-’ c (d 0 G .3 c

3,

.3 rn 0 c

$ E 1

CT)

d Id 0

(d M >

.3 c,

.3

z c 0

c, c a,

&

0 V a, a, m VI

W t- 4 d

0 0

I

W c- 4 d

I I

d d

0 0 V V a, a, a, a, m m m rn

k c

+ 0 3 V

c k 3 P c 0

0 a, c

.rl c,

.- Y

%+ c

+ 0 3 0 c k 3 P m m rn

a c,

r(

h

k c - E t- A d . . O Z

n k E

c o t - c o . . r - M W

Y

c m

w 4 .r(

E d 0

m Lo *

7 -4

(Rev.1 - 12/66)

a C D

CDO . .

N CDO

m 0 u3 M

. . - -

$ z -IN

- CD 0 03

CO N

DCO

D r l D

. .

- -

-IN

- M CO rl

0 M

s Y

a, > cd M a, c -a c cd a, >

.rl +

.d + .rl

m 0 a k 0 rcl ' tr' h

5 Y

1c

a, (d '0

+

3 I

M 31 a CO rl

--ttt--

h k

a , n

+

ffirz

(Rev. 1 - 12/66)

7 . 9

7 . 1 0

7 . 1 1

7 .12

7 . 1 3

7. 1 4

7. 1 5

7. 1 6

7 . 1 7

7 . 18

Uncertainties a t 2 mins. before S P S 2 ignition (Update 1) Tff uncertainties at 1 3 and 2 mins. before SPSB ignition Reentry s t a r t uncertainties (Update 1 )

Reentry s t a r t uncertainties (Update 2)

Flight path angle uncertainties a t reent ry s t a r t (Updatr:s 1 Ri 2 )

Reentry end uncertainties (Update 1) Reentry end uncertainties (Update 2)

Stable Member drift angles and misalignments Propagation of Initial Condition Er rors f rom SPSl cutoff Propagation of Initial Condition E r r o r s f rom Update 2 point

A l l tables for position and velocity uncertainty and for computed t ime-of- flight uncertainty ((U)T ) a r e f o r indication uncertainties (i. e. indicated-minus-

actual) . However, Table 7. 1 3 f u r flight -path-angle uncertainties a r e ior actual- minus -indicated flight -path angle.

f f

Most tables give uncertainties relative to local axes a t event t ime. Posit ive

local axes a r e defined a s follows:

Altitude - outwards along at event t ime.

Track - along X

Range - along &t X Track

7. 5 Initial E r r o r Condition Propagation Tables

The last two tables at the end of this section a r e given t o show how unit initial-condition e r r o r s propagate, f i r s t , f rom SPSl cutoff conditions, and second,

f rom update t ime 1 3 minutes before SPSB ignition. The f i r s t table , 7. 1 7 , shows how velocity magnitude and angle e r r o r s a s well a s altitude e r r o r s propagate during the long coast . e r r o r s propagate f rom the update 2 point, and may be useful for decisions relat ive to th i s update.

7. 6 T Computation Uncertainties

The second table, 7. 18, shows how unit position and velocity

f f The uncertainties in computed free-fal l t ime of flight, (U)Tff , were calcu-

lated by perturbing the equation for ‘r velocity uncertainties due to each IMU component uncertainty. A s indicated in Section 5 . 7 the approximate equation f o r Tff is used when e >, 0.8 and a > 5 X 10

This condition ex is t s at SPSB cutoff, although it does not exist at S P S 2 ignition.

( s ee Section 5. 7 ) with the position and f f

7 ft.

Tables 7 . 1 and 7 . 2 give rss data on the effect of IhlU component uncertainties

on the computation of free-fall t ime of flight t o the reent ry s t a r t altitude of ~ ~ ~ , o ~ ~ feet.

event t imes .

Tables 7. 7 and 7. 10 give detailed data on contributions to (U)Tff at var ious

7 -6

(Rev.1 - 1 2 / 6 6 )

(

The relatively la rge increase in (U)Tff between injection burn and 1st SPS

The AGC accepts accelerometer outputs during th i s t ime interval.

During most of the long coast the rss uncertainty in Tff (with update 5 min

The magnitude of

burn cutoff is pr imar i ly due to the effect of acce lerometer bias during the intervening

coast .

before injection burn) remains almost constant a t 178 seconds.

(U)Tff is approximately the same f o r positive and negative IMU uncertainties.

e v e r , by the t ime that Update 2 t ime is reached (13 minutes before SPS2 burn ignition) the T s iderably as can be seen f rom Tables 7 . 1 and 7 . 2 .

during the long coast a r e X - and Z - acce lerometer b iases and Y- and Z - gyro bias drift .

How -

uncertainties for positive and negative IMU uncertainties have diverged con- f f The pr ime contributors to (U)Tff

7. 7 Flight -Path-Angle Uncertainty

Data on flight -path-angle uncertainty a t reent ry s t a r t is given in Tables 7. 1

and 7.2. Section 7. 11 and Fig. 7 . 2 should be consulted for definitions.

Tables 7. 13 gives detailed data on IMU component contributions to (U)yAA.

Of p r imary interest a r e the data f o r the Update 2 case. Table 7. 13 shows that Y -gyro b ias dr i f t uncertainty (NBDY) is by far the most important contributor

to the rss uncertainty of 0. 175O ( 3 . 05 m r ) . (3. 01 mr), while negative NBDY is responsible for -0.168' (-2. 94 mr).

Posit ive NBDY is responsible fo r 0. 1 7 2 O

Table 7. 13 shows that for NBDY the flight-path-angle uncertainty relat ive

t o spacecraf t actual location axes a t nominal t ime, (U)yAI, is 0.96 m r .

t ime , assuming that BDY is the only uncertainty, the S/C is 10,725 feet above the nominal altitude of 400,000 feet.

flight-path angle t o altitude change is 0. 188 angle f rom nominal t ime to the t ime that the S/C actually reaches 400,000 feet is then 2.02 m r (0.116 deg).

(U)yAA of 2. 98 m r . the alt i tude dependency of (U)yAA f o r this situation.

At nominal

The coefficient, dy/dAlt, relating change in

mr / fee t . The change in path

Summing this with (U)yAI gives an approximation to This is close t o the (U)yAA given above, and c lear ly shows

7 . 8 Navigation Update Conditions and Uncertainties

Because the AGC accepts accel.erorneter outputs during the parking orbi t ,

by the time of injection burn ignition relatively la rge position and velocity indication uncertaint ies develop due to the presence of acce lerometer bias uncertainties.

Because of th i s , a navigational update before injection burn ignition is provided for . w a s assumed.

F o r the e r r o r studies an update 5 minutes before injection burn ignition

A 2nd navigational update 13 minutes before the 2nd SPS burn ignition ( o r 22. 2 minutes of f ree-fal l coast to 400,000-ft altitude) is a l so made. At th i s t ime

7 -7

(Rev.1 - 12/66)

( -~

is nominally 2 minutcs f f ( min) t h e Tff to Tff(min) comparison is s ia r ted , where T before SPS2 burn ignition.

A t both update t imes the uncertainties in position and velocity a r e reduced

to those represented by t h c tracking update uncertainties.

Thc navigational updates would be performed on the bas i s of orbit computa- t ions made using observations by the MSFN (Manned Space Flight Network). tracking-station-computed position and velocity vec tors would be subject to uncer - tainties because of noise and bias in tracking measurements .

The

In t h i s IhlU e r r o r study, simulation of tracking update Uncertainties was based on data available i n MSC Internal Note No. 6 6 - F M - 4 6 , " E r r o r Analysis of

IVlSFN Tracking Data for AS-501" by P. T. Pixley and M. L. Alexander. uncertainty covariance niatr ices for t imes j u s t before injection burn ignition and the 2nd SPS burn ignition were available i n th i s report . The one-sigma position and velocity uncertainties for the two update t imes relative to local ver t ical axes were a s follows:

Tracking

Posit ion Uncertainty Velocity Uncertainty

Alt. Track Range Alt. Track Range (in nautical mi les ) ( in f t / s ec )

5 min before Inject . Burn Ignit. 0 . 10 0.04 0.41 2 . 6 0. 6 0 . 4

13 min before 2nd SPS Ignit. 0 .02 0 . 0 7 0. 08 0 . 3 0. 6 0 . 1

In the e r r o r tables , a f te r updating t ime , the uncertainties include the effects of both navigational update and IIVlT' uncertainties.

7. 9 IMU E r r o r s and Uncertainties

The AGC w i l l be able to provide compensation fo r the measured average

values of the following IRlU component e r r o r s :

1) acce lerometer bias e r r o r (ACB)

2 ) 3) gyro b ias drift (NBD)

4)

5)

acce lerometer scale factor e r r o r (SFE)

gyro input ax i s acceleration sensit ive drift (ADIA) gyro spin reference axis accelerat ion sensit ive drift (ADSRA)

Since the average IMU e r r o r s w i l l be compensated by means of AGC programs

during prelaunch and in flight, it is the actual unpredictable deviations f rom the measured average e r r o r s that constitute the IMU component uncertainties.

The e r r o r tables here employ a definition of sca le factor e r r o r whose

polarity is effectively the r eve r se of that fo rmer ly used. being adopted, since it is consistent with that employed in component and sys tems t e s t s fo r some t ime past .

The new definition is

7 -8

(Rev. l - 1 2 / 6 6 )

!

The Block I IMU e r r o r uncertainties ( see a l so ME1 No. 1015000-Part I) for

the present e r r o r studies a r e a s follows:

Block I One-Sigma IMU Er ro r Uncertainties

Inrmt A x i s

Accelerometer Accelerometer Accelerometer Gyro bias drift

X - bias (ACB) 0.40

sca le factor (SFU) 150 nonlinearity (NC) 10

(B D) 3

Gyro input axis accel. sens. drift (ADIA) Gyro spin axis accel. sens. drift (ADSRA)

Gyro accelerat ion squared sens. drift Accelerometer I. A. misalignments

Non-orthogonality X to Z Non-orthogonality X to Y Y about XsM

About SRA About OA

Gyro I. A. misalignment

8

5 0 .2

0. 14

0. 14 -

0. 50

0 . 50

Z - Y - 0.40 0.40

150 150

10 10

3 3

8 8

5 5

0 . 2 0 . 2

0.50 0. 50 0.50 0. 50

Units

c m / s e c

P P M

I-rg/g

2

2

meru m e r u / g m e r u / g m e n / g 2

m r m r

m r

m r m r

Certain IMU component uncertainties affect both the pre -launch alignment of the Stable Member and the in-flight computation by the AGC of position and

velocity. The acce lerometer bias uncertainties affect the ver t ical erect ion of the Stable Member about YI and ZI (see Fig. 7.1). sensi t ive drift r a t e uncertainties affect the azimuth alignment of the SM through the i r effect on the gyro-compassing loop during pre-launch alignment. shows the effects of the various gyro drift uncertainties on azimuth alignment

uncertainty.

The gyro bias and accelerat ion

Table 7 . 4

This table shows that the rss azimuth alignment uncertainty is 3 . 5 m r .

The IMU uncertainties affecting p re -launch SM alignment and the in-fl i&t Their effects on position navigation computations a r e assumed to be correlated.

and velocity uncertainties a r e accordingly summed arithmetically in the e r r o r

tables.

7. 10 Stable Member Orientation

The orientation of the IMU Stable Member axes (XSM, YSM, ZsM) relat ive

The X, Y, Z acce lero- t o launch iner t ia l axes (XI, YI, ZI) a r e shown in Fig. 7.1. m e t e r and gyro input axes a r e colinear with corresponding Stable Member axes.

The launch inertial axis ZI at the nominal launch azimuth of 72

is in the horizontal plane of launch instant and oriented 0 from north. The XI - Z plane w i l l be the I

7 -9

(Rev. l - 12 /66)

( ~-

initial pitch plane as w e l l as initial reference t ra jec tory plane.

is not realigned during flight.

The Stable Member

Table 7. 16 gives data on S M drif ts and misalignments throughout the flight. Table 7. 4 gives detailcd data o n SM misalignments a t SIVB cutoff.

7. 1 1 Flight -Path-Anglc and Altitudc -Ratr Uncertainty Definitions

Fig. 7 . 2 defines the th ree flight-path-angle uncertainties, (U)yAI, ( U)yAIN Data for (U)yAA a r e given only for reentry s t a r t (at 400,000-ft and (U)yAA.

alt i tude) in the summary tables , 7. 1 and 7. 2 , s ince the flight-path-angle uncertainty with the spacecraft actually at -100,000-ft altitude is the desired paramettlr . all other t imes during the 501 flight, data a r e given for (U)y

F o r

AI' A s the range a n g k uncertainty, (U)Rge/R, i nc reases ( a s it will for prolonged,

non-updated orbital missions, sincc, (1J)Rge is unbounded), the uncertainty, (U)yAIN,

w i l l increase correspondingly, since y

horizontal axis.

is measured relat ive to the nominal A IN

Data in a l l e r r o r tables for RSS position and velocity uncertainties a r e given relat ive t o nominal local ver t ical axes ( see Fig. 7. 2 ) .

compute (U)yAIN. be computed from the tabulated position and velocity uncertainties.

These data may be used to Unless appropriate t ransformations a r e made, (U)yAI can not

7. 1 2 E r r o r Computation Procedure

Posit ion and velocity uncertainties given in the tables were computed a s

follows. component e r r o r on t ra jectory position and velocity. 1) that the e r r o r s were smal l rt.lative i.0 the pa rame te r s being nw asu red , and

2 ) that the IMU component e r r o r s \vere statist ically independent of each other .

The e r r o r equations took i n t o account the effect of the IRIU e r r o r s on gravity vector computation. requi re nominal t ra jec tory acceleration and position vectors ( re la t ive to fixed

iner t ia l axes) as inputs at d i scre te t ime intervals. w a s generated in a separate program. detailed e r r o r printouts were made giving the position and velocity uncertainties

due t o each IMU uncertainty relative to nominal local ver t ica l axes .

Approximate e r r o r equations were derived for the effect of each IMU The assumptions were:

The computation program incorporating the e r r o r equations

The nominal t ra jec tory itself A t significant events , such a s SIVB cutoff,

7 . 13 E r r o r Table Explanation

To make up a one-page e r r o r table format some of the more insignificant

e r r o r sou rces were omitted.

misalignments and gyro accelerat ion and accelerat ion squared sensit ive dr i f t uncertainties.

These included cer ta in acce lerometer input ax is

7-10

(Rev. 1 - 12/66)

I NOMINAL POSITION OF S PAC EC R A F T

DESIRED ALTITUDE

NOMINAL LOCAL

ACTUAL POSITION OF SPACECRAFT WHEN INDICATED ALTITUDE EQUALS DESIRED ALTITUDE

OF

DESIRED ALTITUDE

NOTE: THE FLIGHT PATH ANGLES, YAlNt YAI, AND Y A A ARE ALL SHOWN NEGATIVE. THE ANGLES 4, AND 42 ARE SHOWN POSITIVI

Flight Path Angle and Velocity Uncertainty Equations

Altitude Rat e Uncertainty Equations 0 - - -

(U)AltAIN = VAI s h y A I N - V (U)AitAI (U)AitAA = vAA sinyAA - v N s inyN

s h y N z (VAI - V,) * R /R N N N = vAI sinYAI - v N s inyN

Note that yAIN is measured with respect t o the s a m e nominal horizont.31 axis as yN, while yAI and yAA a re measured with respec t to thr>ii-

par t icular local horizontal axes.

Fig. 7-2 Flight Path Angles

7 - 1 1

(Rev.1 - 12/66)

The principal IMU component uncertaint ies are denoted by the following abbreviations:

2 ACB - acce lerometer bias. Units are c m / s e c . SFU - acce lerometer sca le fac tor uncertainty. Units a r e pa r t s pe r

million.

NC - acceleration squared indication uncertainty. Units a r e micro-g -6 2

per g2 (10 g /g )

NBD - gyro null bias drift uncertainty. Units a r e m e r u (mil l i -ear th ra te units)

ADIA - gyro input ax is accelerat ion sensit ive drift . Units a r e meru/g .

ADSRA - gyro spin reference axis accelerat ion sensit ive drift . Units a r e meru/g .

2 AD - gyro accelerat ion squared sensit ive drift . Units a r e m e r u / g (e. g . , ADMX - gyro input axis accelerat ion squared sensit ive drift)

The symbols X, Y, Z denote the input ax is of the IMU gyro o r acce lerometer t o which the uncertainty applies. axes .

IMU component input axes a r e identical with SM

Initial SM Misalignments (Uncorrelated) - The misalignment uncertainty about X (azimuth) of 0. 5 m r is pr imar i ly due to gyro input axis misalignments I re la t ive t o SM axes.

independent of the SM misalignments caused by the effect of IMU component un- cer ta inty on SM vert ical erect ion and azimuth alignment. These p re -launch SM misalignments a r e partially given elsewhere in the e r r o r tables .

Table 7.4

This misalignment a s well a s those about YI and ZI a r e

For SIVB cutoff, gives the individual SM pre-launch alignment uncertainties.

The subscript "INIT", used with some of the IMU uncertainties such a s ACBX, denotes those uncertainties (position and velocity) due to the initial p r e - launch SM misalignment caused by the par t icu lar IMU uncertainty. The subscr ipt

"FLGT" denotes those uncertainties due to the effect of the IMU uncertainty on in- flight navigation computations. F o r gyro drift t e r m s it a l so includes the effect of gyro drift since launch.

the above two effects.

The subscr ipt "COMB" denotes the ar i thmetic sum of

7-12

(Rev . l - 12/66)

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(Rev. 1 - 12/66)

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7-28

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R-537

DIST RIRUTION LIST

Internal

R. Battin

D. Bowler

E. c o p p s s. c o p p s

* J. Dahlen

*G. Edmonds S. Felix P. Fellemen J. Fleming (4)

F. Grant D. Hoag

* M . Johnston

J. Kernan A. Kosmala

A. Koso * W. Kupfer (20)

L. Larson G. Levine

* D. Lickly (15)

H. Little G. Mayo

R. McKern *H. McOuat

J. S. M i l l e r * J. E. M i l l e r

* R . Morth

J. Nevins *P. Peck

MIT Instrumentation Laboratory (4) c / o No. American Aviation, Inc. , S&ID 1 2 2 1 4 Lakewood Boulevard Downey, California 90241 Attn: (TOD)

MIT Instrumentation Laboratory ( 2 ) G&N System Laboratory c / o Grumman Aircraf t Engineering Corp. L M Projec t - Plant 25 Bethpage, Long Island, New York Attn: (TOD)

R. Ragan *J . Rhode

R. Scholten

N. Sea r s *P. Seligson (4) * J. Shillingford *: M. Sullivan

J. Suomala R. Weatherbee

R. Werner * W. Widnall

L. Wilk

R. Woodbury Apollo Library (10)

MIT/IL Library (3)

MIT Instrumentation Laboratory (3) Code/EG MIT Building 1 6 NASA Manned Spacecraft Center Houston, Texas 77058 Attn: (TOD)

NASA Manned Spacecraft Center Houston, Texas 77058 Attn: Mr. S. Laquidara - FC

(1 5)

Bldg. 45, Room 442

MIT Instrumentation Laboratory (5) P.O. Box 21025 Dennedy Space Center , Florida 32815 Attn: (TOD)

* Names marked with a n a s t e r i sk (defining "501 distribution") w i l l receive all

changes a s they a r e made.

i s sued o r completely revised. A l l GSOP recipients w i l l receive the periodically i ssued record of revisions frontispiece to Vol. I.

Others w i l l receive documents only when initially

D-1

(Rev. 1 - 12/66)

External

A. Metzger (NASA/RASPO)

AC Electronics , Div. of GMC

Kollsman Instrument Corp. Raytheon Company Major Wheeler (AFSC/MIT) NAA RASP0 National Aeronautics and Space Administration

Resident Apollo Spacecraft P r o g r a m Office North American Aviation, Inc. Space & Information Systems Division 12214 Lakewood Boulevard Downey, California

F 0: National Aeronautics and Space Administration, MSC (3) Florida Operations, Box MS Cocoa Beach, Florida 32931 Attn: Document Control

HDQ: NASA Headquarters 600 Independence Ave., SW Washington 25, D. C. 20546 Attn: MAP-2

AMES: National Aeronautics and Space Administration Ames Research Center Moffett Field, California Attn: L ib ra ry

National Aeronautics and Space Administration L e w i s Research Center Cleveland, Ohio Attn: L ib ra ry

LEWIS:

FRC: National Aeronautics and Space Administration Flight Research Center Edwards AFB, California Attn: Research Library

LRC:

GSFC:

National Aeronautics and Space Administration Langley Research Center Langley AFB, Virginia Attn: M r . A.T. Mattson

National Aeronautics and Space Administration Goddard Space Flight Center Greenbelt, Maryland Attn: Attn: M r . Paul Pashby Code 554

Manned Flight Support Office Code 512

D-3

(Rev.l - 12 /66)

External (cont 'd)

ERC:

GAEC:

NAA :

GAEC RASPO:

AC RASPO:

WSMR:

MSFC:

KSC:

MSC :

National Aeronautics & Space Administration Electronics Research Center 575 Technology Square Cambridge, Massachusetts Attn: M r . R. Hayes / Mr. A. Colella

Grumman Aircraf t Engineering Corporation Bethpage, Long Island, New York Attn: M r . C. Brown (1)

Mr. R. Adarnato (1) M r . B. Sidor (1) Mr. H. Sherman (1) Mr. E. Stern (1R) Mr. R. F le i s ig (1) Mr . G. Smith (1) Mr. R. P r a t t (1)

North American Aviation, Inc. Space & Information Systems Division 12214 Lakewood Boulevard Downey, California 90241 Attn: Apollo Data Requirements

Dept. 096-340 Bldg. 3, CA 99

National Aeronautics and Space Administration Resident Apollo Spacecraft P r o g r a m Officer Grumman Aircraf t Engineering Corporation Bethpage,. Long Island, New York

National Aeronautics and Space Administration Resident Apollo Spacecraft P r o g r a m Officer Dept. 32-31 AC Electronics Division of General Motors Milwaukee, Wisconsin 53201 Attn: Mr. W. Swingle

National Aeronautics and Space Administration Pos t Office Drawer MM Las Cruces , New Mexico Attn: B W 44

National Aeronautics & Space Administration George C. Marshal l Space Flight Center Huntsville, Alabama Attn: R-ASTR-5

Flight Crew Training Mail Code HW Kennedy Space Center , Florida Attn: Mr. Frank Hughes

National Aeronautics and Space Administration Manned Spacecraft Center Apollo Document Control Group ( P A 2) Houston, Texas 77058

(Above includes information copy f o r PD-6, s e e l e t t e r P P - 7 -66 -775)

(1)

( 7 + 1 R )

(1 + 1 R )

BELLCOM ( 3 )

D-4

(Rev.1 - 12/66)