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R-477

(Unclassified Title)

GUIDANCE AND NAVIGATION

SYSTEM OPERATIONS PLAN

APOLLO MISSION 202

January 1965

John M. Dahlen

Albrecht Kosmala

Daniel J. Liqkly

John T. Shillingford

Balraj Sokkappa

CAMBRIDGE

I NSTRU M E NTATI O N

LABORATORY39, MASSACHUSETTS

COPY# //__ OF _,°.3¢3 COPIES

THIS DOCUMENT CONTAINS I.,P_ PAGES

- +-""' " l - " 2_:I_'_2 '.7

r

ACKNOW LEDGMENT

This report was prepared under DSR Project 55-191,

sponsored by the Manned Spacecraft Center of the National

Aeronautics and Space Administration through Contract

NAS 9-153.

This documen_contains/ormation affecting

the national defe_ of_e United States within

the meaning of the_onage Laws, Title 18,

U. S. C., Sections 793_94, the transmission

or the revelation oflhich "_. a?y manner to an

unauthorized perso_ is prohi_d by law.

2

TABLE OF CONTENTS

Section

1

2

3

4

5

6

7

8

9

10

INTRODUCTION

G&N FLIGHT OPERATIONS SUMMARY

LOGIC AND TIMELINE FOR SPACE-

CRAFT AND MISSION CONTROL

GUIDANCE EQUATIONS

CONTROL DATA

G&N ERROR ANALYSIS

G&N CONFIGURATION

INSTRUMENTATION

G&N PERFORMANCE ANALYSIS

DISTRIBUTION

l° INTRODUCTION

i. 1 Purpose

This plan governs the operation of the Guidance and Navigation System and

defines its functional interface with the spacecraft and ground support systems on

Mission 202.

1.2 Authority

This plan constitutes a control document to govern the implementation of:

(1) Detailed G&N flight test objectives

(2) G&N interfaces with the spacecraft and launch vehicle

(3) Digital UPLINK to the Apollo Guidance Computer (AGC)

(4) AGC logic and timeline for spacecraft control

(5) Guidance and navigation equations _'"

(6) Digital DOWNLINK from the AGC

(7) G&N System configuration

Revisions to this plan which reflect changes in control items (1) through (7) require

approval of the NASA Configuration Control Board.

This plan also constitutes an information document to define:

(i) Trajectory uncertainties due to G&N component errors (Error Analysis)

(2) Trajectory deviations due to spacecraft performance variations and

launch vehicle cut-off dispersions (Performance Analysis)

(3) G&N instrumentation (PCM telemetry and on-board recording) exclusive

of AGC DOWNLINK

(4) External tracking data

Revisions to this plan which reflect changes in information items (1) through (4)

will not require approval of the NASA CCB.

1.3 Preliminary Data

Unapproved data are printed in red.

#To support these functions this document contains a Control Data section which

defines the reference trajectory, AGC memory data and applicable mission data

(mass, propulsion, aerodynamic and SCS data)

1-1

•" _,_T_.._._.._ _

2. G&N FLIGHT OPERATIONS SUMMARY

This section defines the mission plan as originated by NASA and summarizes the

manner in which the G&N system will operate to implement this plan as developed by

M[T in cooperalion with NASA and NAA/S&ID. This section is divided into three parts:

Par 2. 1 Test Objectives

Par 2.2 Spacecraft and Mission Control

Par 2.3 Mission Description

2. 1 Test Objectives

2. I. 1 Spacecraft Test Objectives which require proper operation of G&N

System:

i) Evaluate the thermal performance of the CM heat shieldablator

during a high heat load, long duration entry.

2) Demonstrate CM adequacy for manned entry from low earth orbit.

3) Determine nominal mode separation characteristics of the CSM

from the SIVB and the CM from the SM.

4) Demonstrate multiple SPS restart (after the second major burn,

two 3 second burns with i0 second intervals between burns are

required).

5) Determine performance of CSM systems: G&N, SCS, ECS (pressure

and temperature control), EPS, RCS and Telecommunications.

2. I. 2 Detailed G&N Test Objectives

i) Evaluate performance of the following integrated G&N/Spacecraft

modes of operation:

a. Boost Monitor

b. Thrust Vector Control

c. Orbit Attitude Control

d. Lift Vector Control

e. Unmanned Spacecraft Control

2) Determine accuracy of G&N system in computation of spacecraft

position and velocity during all mission phases.

2.2 Spacecraft & Mission Control

2.2.1 Spacecraft Control

Spacecraft Control is implemented by the Apollo Guidance Computer

(AGC) provided by MIT and the Mission Control Programmer (MCP) pro-

vided by NAA/S&ID. Basically, the MCP performs those non-guidance func-

tions that would otherwise be performed by the crew, while the AGC initiates

major modes which are dependent upon trajectory or guidance functions.

The function interface between the AGC and the MCP is complex and

its description is deferred until Section 3. The electrical interface is simple,

2-1

being relay contacts in the AGCDSKYwired to the MCP, and is describedin ICD MH01-01200-216. Thefollowing AGCoutputdiscrete signalsare

G&NATT. CONTR. MODESELECTG&NENTRYMODESELECTG&NAV MODESELECT+ X TRANSLATIONON/OFFCM/SM SEPARATIONCOMMANDFDAI ALIGNT/C ANTENNASWITCHG&NFAIL INDICATION

9) 0.05 g INDICATIONI0) GIMBAL MOTORPOWERON/OFFii) SPARE

2.2. 2 Mission ControlMission Control is provided by the Clear Lake Mission Control Cen-

ter (CLMCC)via the Digital CommandSystem(DCS),which has manydis-crete inputsto the spacecraft andanUPLINK to theAGC. The discretecommandsto the spacecraft andthe AGC UPLINK are described in Section3.

TheAGCUPLINKprovides the CLMCCwith the capability to enterthe AGC with any instruction or datawhich canbe enteredby the crew via theDSKYkeyboard. It is specifically plannedto use this link to provide the AGCwith several discrete commandsfor contingencies. This link will also beusedto updatethe orbit parameters in erasable memory with more accuratedata if it is available from ground tracking.2.2.3 GuidanceErrors

Theperformanceof the G&Nsystem for mission 202hasbeenestimatedassumingthat nonavigation datais inserted via the AGC UP-LINK.

provided:I)2)3)4)5)6)7)8)

Themost significant G&Nerror is that error in the critical pathangleat entry which is estimated to be 0. 165degreeona one sigma basis.Thenextmost significant error is manifestedin the CEPat splashwhich isestimatedto be 15.6 n.m.

A completebreakdownof G&Nerrors is given in Section8.

2.3 Mission DescriptionThepurposeof this section is to describe G&Nfunctions during eachmission

phase. Note that thesefunctions are describedin greater detail, sufficient todefineprogrammingrequirements, in Section3.

2-2• - , .Im

D

J

The reference trajectory is defined in Section S in sufficient detail to satisfy

MIT's requirements for development of guidance equations, spacecraft control

logic and determination of flight environment.

Section 9 presents those path and attitude characteristics resulting from

guidance control which are believed to have significant effects on other spacecraft

equipment and ground support systems.

The overall mission profile is illustrated in Fig. 5-i and Table 5-1 and

mi_:ht well be examined at this point.

2.3.1 Pre-gaunch

During this phase the IMU stable member is held at a fixed orientation

with respect to the earth. The X PIPA input axis is held to the local vertical

(up) by torquing the stable member about Y and Z in response to Y and Z PIPA

outputs. Azimuth orientation about the X axis is held by a gyro-compassing

loop such that the Z PIPA axis point downrange at an a_imuth of 105 degrees

East of True North. Initial azimuth is determined by tracking a ground target

with the G&N Sextant prior to closeout at -i i hours. Upon receipt of the

GUIDANCE RELEASE signal from the Saturn I.U. the stable member is

released to maintain a fixed orientation in inertial space for the remainder

of the mission. In this manner the Saturn and Apollo IMU stable members

retain a fixed relative orientation. Also at the time of GUIDANCE RELEASE

the G&N system_ starts its computation of position and velocity which corLtinucs

until first SPS burn cut-off.

2.3.2 SI Boost

The boost trajectory is described in Fig. 5- 2 . Upon receipt of the

LIFT OFF signal from the Saturn I.U. the AGC will command the CDUs to

the time history of gimbal angles associated with the nominal SI attitude

polynomials. The CDU outputs will then represent vehicle attitude errors

and will be displayed on the FDAI and telemetered to the ground. This SI

attitude monitor is a required element of the launch vehicle malfunction de-

tection scheme, and, in association with computed position and velocity,

constitute the Boost Monitor data provided by the G&N system during this

period.

2.3.3 Staging, Coast and SIVB Boost

The G&N system will not have the capability to control the SIVB.

During this period the G&N system will monitor IMU gimbal angles to detect

tumbling and will compute the free fall time to entry interface altitude

(300,000 ft.) from present position and velocity. These quantities are used

in the Abort Logic and, in association with computed position and velocity,

constitute the Boost Monitor data provided by the G&N system during this

period.

2-3

2. 3.4 Aborts from SIVBBoostAborts from the boostphaseare mechanizedin the sameway as

mannedflight aborts wheneverpossible. G&Ncontrol of CSMaborts fromSIVBboost is enabledby the MCP 2 secondsafter start of the MCPSIVB/CSMSeparationsequence. Uponreceipt of the SIVB/CSMSEPARATIONsignal from the spacecraft the AGC determinesa sequenceof eventsusingthe control logic givenin SectionS. Briefly, the sequenceof eventsisderived from three tests:

A. Has the AGCreceived the ABORT signal from the groundviathe UPLINK?

B. Dothe spacecraftbody rates exceedthe tumbling threshold?C. Doesthe free-fall time to entry interface altitude fall below

the abort Tf criterion of 160seconds?

For NOABORTandNO TUMBLING, the AGC commandsa normalseparationandSPSburn to the nominal First Burn aim point as describedmore fully below.

If the ABORT signal is received and there is NO TUMBLING, the

AGC commands an abort separation sequence followed by an SPS abort burn

to the downrange Atlantic Recovery Point. Landing area control capability

is illustrated on Fig. 5-1which shows a continuous recovery area and the

selected downrange Atlantic Recovery Point. This downrange point is the

splash point obtained if, after a nominaISIVB cut-off and separation sequence,

the SPS is fired for 7 seconds in the trajectory plane with the spacecraft

X axis 35 degrees above the visible horizon and the CM is oriented for full

up lift during the entry phase. The G&N system will control the thrust and

lift vectors to achieve this splash point with the constraints, (i) that the

spacecraft X axis be directed 35 degrees above the visible horizon during

thrusting and (2) that the entry point - splash point separation provide CM

lift sufficiently positive to reduce entry g's below a level acceptable for

human tolerance. A I0 g limit is incorporated in the entry program also to

minimize excessive g loads.

If the abort occurs too early in the boost phase or at an "unsafe"

flight path angle, the selected downrange Atlantic Recovery Point cannot be

reached because either (i) there is insufficient fuel in the SM tanks, or (2)

the booster cut-off conditions are such that the spacecraft would dip into the

atmosphere while thrusting. These two conditions are avoided by test C

which is mechanized as an interrupt. If the free-fall time falls below 160

seconds so that test C results in a NO answer, the AGC will command engine

shutdown and a CSM attitude maneuver to the CM/SM separation attitude.

2-4

When the free falltime to entry interface altitude falls below 75 seconds the

AGC will command CM/SM SEPARATION and CM orientation to the aero-

dynamic trim attitude. The liftvector will be up during the entry phase.

Note that "early" aborts result in splash points within the continuous recovery

area.

If TUMBLING is detected, the AGC will start the SPS 2.5 seconds after

separation. This will result in stabilization by the SCS rate loops, and SPS

cutoff by the AGC when it senses that spacecraft body rates have dropped

below the tumbling threshold. Following SPS shutdown the AGC will estimate

the maneuver time, TM, required to orient to the abort SPS burn attitude

(X axis 35 degrees above the visible horizon). If the free-fall time to entry

interface altitude is greater than T M + 160, the AGC will command the CSM

to the abort SPS burn attitude, command engine on at T M and guide to the

downrange Atlantic Recovery area. Again as in the non-tumbling abort case

the engine will be shutdown if free-fall time drops below 160 seconds. If

after tumbling arrest burn shutdown the free-fall time is less than T M + 160,

the AGC will command the CSM to the CM/SM SEPARATION attitude. Abort

area control is illustrated in Fig. 9-4.

2.3.5 CSM/SIVB Separation

There are two CSM/SIVB separation sequences, a normal sequence

and an abort sequence used if tumbling or the abort signal is present. In the

normal sequence the SPS is ignited by the AGC a fixed time delay of 12.7

seconds after it receives the CSM/SIVB SEPARATION signal. This time delay

permits the RCS ullage thrust to build up enough separation distance to prevent

the SPS from damaging the SIVB or upsetting its attitude. On the other hand

the time delay is not so long as to cause an unjustified ZXV penalty. After

separation the AGC computes the initial SPS thrust attitude and commands

the required attitude maneuver. If the spacecraft is not completely oriented

at the end of the fixed time delay, the SPS is started anyway and orientation

is completed during the first few seconds of the burn. Only when large rates

and/or large negative pitch attitude dispersions exist at SIVB cut-off will the

fixed time delay be too short to permit completion of spacecraft orientation

before SPS ignition.

In the abort separation sequence, the SPS is ignited by the AGC a time

delay of 2.5 seconds after it receives the CSM/SIVB SEPARATION signal.

This time delay is made as short as possible to minimize the probability of

CSM-SIVB re-contact or loss of IMU reference in the tumbling case and to get

the CSM away from the SIVB as quickly as possible in any abort case.

2-5

2.3.6 SPS First Burn

First burn thrust will be controlled by the G&N system to achieve the

reference trajectory major axis and eccentricity at cut-off. The trajectory

plane at cut-off will include the Pacific Recovery Point at nominal splash

time. The nominal attitude, flight path angle and altitude histories are

given by Fig. 9-1. Section 9 also contains tables which show the effects of

CSM performance variations and launch vehicle cutoff dispersions. The

steer law used in this maneuver is given in Section 4, where are found all

the CSM guidance equations for Mission 202. It will be noted that the

universal cross product steering law for Apollo is used whenever possible,

specifically, for this mission, in all cases except tumbling arrest and the

short third and fourth burns.

2. 3.7 Coast Phase, First Burn Cut-off to Second Burn Ignition

Following first burn cut-off the AGC will compute and command a

spacecraft attitude maneuver to align the X-axis to the local vertical, nose

down, and the Y-axis to a fixed angle of 0 degrees from the angular momentum

vector R * V. Simultaneously the AGC will establish the second burn

ignition point by a process of precision numerical integration.

When the inner gimbal angle reaches degrees the AGC will com-

mand FDAI ALIGN for i0 seconds thereby resetting the backup attitude

reference to correct for its accumulated drift error. The CSM attitude during

this interval will be within 1 degree of a pre-determined attitude with respect

to the IMU stable member.

After a time interval of 2006 secs. from first burn cut-off the vehicle

attitude in tracking the local vertical will come closest, in the nominal case,

to the second burn ignition attitude. At this time the local vertical mode will

be terminated and the AGC will command the vehicle to the second burn

ignition attitude, which it will hold inertially until ignition.

2.3.8 Second, Third and Fourth SPS Burns

Second burn ignition occurs after a fixed time delay of 3041 seconds

from first burn cut-off. The AGC will command + X TRANSLATION 30 seconds

before ignition to provide ullage. Thrust is controlled by the G&N system to

achieve the reference trajectory major axis and eccentricity at cut-off, and

a trajectory plane which includes the Pacific Recovery Point at nominal splash

time.

Second burn is terminated by the AGC six seconds before the required

velocity is attained. The spacecraft attitude at this time will be held until

fourth burn cutoff. During second burn the G&N attitude error signal will

develop a bias proportional to the e.g. shift from the engine gimbal trim

position set in prior to second burn ignition. After second burn cutoff the

2-6

CDUs will be moved off from their position at cutoff by a stored estimate of

this bias in order to minimize the attitude transient after engine shutdown.

The AGC will start and shutdown the SPS on a time basis so that the

last two burns are each of 3 seconds duration and so that the two short coast

periods are each of i0 seconds duration. The AGC will control the + X TRANS-

LATION signal so that the RCS will provide ullage thrust as well as attitude

control during the i0 second coast periods. Note that the SCS disables

+X translation during SPS firing.

The nominal attitude, flightpath angle and altitude histories are

given by Fig. 9-2. Figure 9-3 shows the slant range, azimuth and elevation

to the CSM from the Carnarvon tracking station. Tables in Section 9 show the

effects of CSM performance variations and launch vehicle cut off dispersions.

2.3.9 Pre-Entry Sequence

The fourth burn cutoff attitude is held until the free-fall time to entry

interface altitude drops below the normal Tf criterion of 160 seconds, when

the G&N system will s_ar_ pitching the spacecraft up to the CM/SM separation

attitude (+ X axis up in the trajectory plane and tipped forward in the direetion

of motion 60 degrees above the velocity vector. When the free-fall time drops

below 75 seconds the AGC will command CM/SM SEPARATION. After a 5 second

time delay to allow for separation and stabilization, the G&N system will start

orienting the CM to the entry attitude. The CM will then be at the aerodynamic

trim angle of attack with roll angle for down lift.

2.3. I0 Entry

The velocity and critical flightpath angle at entry are directly controlled

by the G&N system during the second, third and fourth burns. The nominal

entry trajectory provided by the guidance equations is illustrated in Fig. 9-4.

The entry guidance equations, which are given in Section 4, are designed to

provide a trajectory which will satisfy heat shield test objectives while con-

trolling the roll angle so as to splash at the designated Pacific Recovery Point.

2-7

3. LOGIC AND TIMELINE FOR SPACECRAFT AND MISSION CONTROL

3.1 Interfaces, Ground Commands and Constraints

3. 1. 1 G&N Interface with Spacecraft

The following interfaces will be effective on Mission 202/AF 011/

AGE 017:

3. I. I. I AGC Outputs to MCP

This interface is documented in ICD No. MH01-01200-216

and provides the following signals:

(i) G&N ATTITUDE CONTROL MODE SELECT

(2) G&N ENTRY MODE SELECT

(3) G&N AV MODE SELECT

(4) +X TRANSLATION ON/OFF

There is a requirement for this command (over and

above the translation requirement) to provide for termination

of Direct Ullage mode.

At SIVB/CSM Separation the AGC must command "+X

TRANSLATION ON" to key the MCP to terminate the "SIVB/

CSM Separate" command to the MESC, which in turn deacti-

vates the MESC-controlled "DIRECT ULLAGE" command.

The MESC will not terminate direct ullage earlier than 3.5

sec after receipt of "SIVB/CSM Separate" nor continue it

longer than 12 _ec regardless of whether the SIVB/CSM

Separate command is terminated or not.

(5) CM/SM SEPARATION COMMAND

(6) FDAI ALIGN

This signal will be initiated and held for 10 seconds soon

after orientation to SPS second burn attitude when the IMU

gimbal angles are near prescribed values. This will result

in FDAI ALIGN when the spacecraft is within 1-1/2 degrees

of a prescribed inertial orientation.

(7) T/C ANTENNA SWITCH

The AGC has the capability to switch the T/C Antennas

although the requirements for this function have not yet been

defined and thus not incorpor porated in AGC programming.

(8) G&N FAIL INDICATION

This signal is generated by the AGC based upon its as-

sessment of certain functions with the G&N. The AGC gen-

erated G&N FAIL INDICATION will be an "OR" of the fol-

lowing normal Block i FAIL indices;

3-i

IMU FAIL - an "OR" of IG servo errorMG servo errorOGservo error3200 CPS loss

wheel supply loss

ACCEL FAIL - an "OR" of

x PIPA error

y PIPA error

z PIPA error

CDU FAIL - an "OR" of CDU 25. 6 KC supply

CDU Motor excit, loss

Inner CDU error

Middle CDU error

Outer CDU error

Each of these three FAIL signals (IMU, ACCEL, CDU)

are subject to AGC program processing as there are certain

phases of normal G&N operation where the FAIL parameters

will exceed FAIL thresholds, thus requiring AGC inhibition

of FAIL indication. The G{N FAIL circuitry will be scanned

for evidence of failure every _50 ms.

As is apparent from the FAIL parameters the G&N FAIL

INDICATION signal is basically a monitor of the inertial sub-

system and not of the AGC. Thus, confirmation of an AGC

failure must be made by the ground by examination of the

DIGITAL DOWNLINK.

The G&N FAIL INDICATION can also be sent to the MCP

via the Up Data Link (UDL) based upon ground assessment of

tracking or telemetry data. Upon receipt of G&N FAIL INDI-

CATION the MCP immediately disables all mode commands

from the AGC and commands the SCS system to SCS ATTITUDE

CONTROL MODE. The attitude reference becomes the BMAG's.

The SCS system is now no longer responsive to any G&N

originated attitude signals, attitude error signals, engine on-

off commands (disabled by removal of AV mode), or AGC

commands via the MCP.

The MCP can be reset once to retransfer S/C control to

G&N, however, this command must come from the ground.

(9) .05G INDICATION

G&N will sense . 05G with the PIPA's, give this indication

to the SCS (via the MCP) and the SCS system will inhibit pitch

and yaw attitude control on the assumption that these axes will

3-2

i[

b_ st_Liiizcdby aerodynamic forces. Should the G&N .05G

indication not be received by the MCP/SCS this attitude con-

trol would not be inhibited, and if sufficient pitch and yaw

attitude errors are generated, ,RCS fuel would be wasted

throughout entry. The G&N entry program will attempt to

null the pitch and yaw error signals during entry based on its

estimation of the pitch and yaw trim angles of attack. MIT

estimates that the resulting pitch and yaw attitude errors will

not exceed the deadbands in the SCSo Should this be incorrect

IICS fuel loss will occur. The G&N 0.05G indication is not

used within the re-entry program, however, so should this

function be backed up by a redundant CM sensor or by the UDL

signal, no AGC confusion should result.

(10) GIMBAL MOTOR POWER ON/OFF

The AGC must terminate SPS GIMBAL MOTOR POWER

in order to key the MCP to select the appropriate SPS motor

gimbal trim inputs. The MCP does this sequentially and

therefore the AGC must terminate this command only once

after 1st SPS burn, (to select trim position for 2nd burn) and

once after 2nd SPS burn (to select trim position for 3rd burn-

ing). The trim position for the 1st burn is selected by MCP

upon keying from the SIVB/CSM Separate Command. The 3rd

burn trim position is also satisfactory for the 4th burn.

(ii) SPARE

This is a relay identical to those used in (i) through (I0)

and is identically wired to the MIT/NAA interface.

3. I. I. i. 1 Detailed Interface Operation

Certain additional facts are pertinent to the use

and comprehension of the AGC/MCP interface:

(I) The AGC must not command more than one SCS mode

simultaneously. This requires termination of each mode

before commanding the next; 250 ms has been established

as sufficient time interval between termination and selec-

tion.

(2) The response of the SCS system to the commands and/or

indication signals of the AGC via the 141CP are subject to the

arming of these command/indications by the MCP. Pres-

ently the arming logic for the G&N/MCP interface is as

shown in Fig. 3-i.

MCP_28 VDC

G&N FAIL: G&N _[__

G&N FAIL: GROUND[_

SPS GIMBAL MOTOR POWERCONTROL ARM

FDAI ALIGN ARM

T/C ANTENNA SWITCH ARM

SIVB/CSM SEP START-

SATURN I. II

LET JETT: SATURNI. U. OR GROUND

ABORT: GROIIND

CM/SM SEP C___MND: G&____N_

CM/SM SEP CMND. GttOUNDq_]

G&N ENTRY MODE

G&N AV MODE

G&NATT. CONT. MODE

CM/SM SEP COMMAND

+X TRANS ON/OFF

• 05 G IND. ARM

MCP +28 VDC

G&N FAIL INHIBIT: GROUNDG&N FAIL IND. ARM

Fig. 3-1 Arming Logic for G&N/MCP Interface

3-4

ICD. NO.

MH01-01024-416

MH01-01025-416

MH01-01038-216

MH01-01028-216

MH01-01028-216

MH01-01036-200

MH01-01278-216

MH01-01280-216

(3) In all case the MCP initiates SIVB/CSM Separation. For

normal cases its action is keyed upon notification from

the Saturn I.U.. For aborts the ground must command

the MCP to start the separation.

3. I. I. 2 Additional Interfaces

Pertinent G&N electrical signal interfaces with other S/C

subsystems are described in detail in the ICD's below.

TITLE

Attitude Error Signals

Total Attitude Signals

Engine On Signal to SCS

Central Timing Equipment

Synch. Pulse

G&N DATA Transmission

to Operational PCM

Telemetry Equipment

ACE Uplink/Spacecraft

Digital Up-Data Link toAGC

Launch Vehicle to G&N Inter-

faces (Block I Series i00)

Vehicle Separation Signals

to AGC (Block I Series i00)

SIGNALS INCLUDED

Pitch Error (Body & Body Offset)

Yaw Error (Body)

Yaw Error (Body Offset)

Roll Error (Body)

Roll Error (Body Offset)

Error Sign_.l Reference

(all signals go from G&N to SCS)

SIN AIG

COS AIG

SIN AMG

COS AMG

SIN AOG

i-I g'_kzK_2S A g'_g-_

Attitude Signal Reference

(all signals go from G&N to SCS)

Engine ON/OFF

(AGC command to SCS system - notvia MCP)

AGC synch, pulse to PCM telemetry

system

G&N analog data and AGC serial

digital data (AGC Downlink) to PCM

(Includes flight recorder aata).

Coded data input to AGC from ground

1. Liftoff

2. Guidance Reference Release

1. CSM/SIVB Separate

3-5

3.1.2 GroundCommands3. i. 2. 1 Digital UPLINK to AGC

By meansof the AGC Uplink, the groundcaninsert dataorinstruct the AGCin the samemanner normally performed by the crewusingthe DSKYKeyboard. The AGCwill be programmedto acceptthefollowing Uplink inputs:

(i) ABORTINDICATION(required for abort logic as describedearlier)

(2) LIFT-OFF (backupto discrete input)(3) SIV-B/CSMSEPARATION(backupto discrete input)(4) G&NATTITUDE CONTROLMODESELECT(5) G&NAV MODESELFCT(6) G&NENTRY MODESELECT(7) +X TRANSLATIONON/OFF(8) GIMBALMOTORPOWERON/OFF

(inputs (4) - (8)will causethe AGCto issue thesecom-mandsto the MCP)

(9) Position andVelocity data (provides groundcapabilityto updatenavigationdatain the AGC).

Operationalprocedures governingthe use of theseUplinkinputs mustbe developedto ensureproper operationwithin program con-straint s.

All information receivedby the AGCfrom the Uplink is in theform of keyboardcharacters. Eachcharacter transmitted to the AGCistriply redundant. Thus, if C is the 5 bit character code, then the 16bitmessagehasthe form:

ICCCwhere-Cdenotesthe bit-by bit complementof C. To these 16bits ofinformation the groundaddsa 3bit code specifyingwhich system aboardthe spacecraft is to be the final recipient of the dataand a 3 bit codeindicatingwhich spacecraft should receive the information. The 22 totalbits are sub-bit encoded(replacing eachbit with a 5bit codefor trans-mission. ) The rate of transmission is 1 K bits/sec, allowing for slightlyover 9 keyboardcharacters per sec. If the messageis receivedandsuccessfullydecoded,the receiver onboardwill sendback an 8 bit"messageacceptedpulse" to the groundandshift the original 16bits tothe AGC(ICCC).

3-6

3. i° 2. _ Discrete Real Time Commands to MCP

The following list details the real-time commands planned

for support of the Apollo 202 Mission. This list is restricted to com-

mands for the Command/Service Module Systems and is exclusive of

commands to the SIVB and AGC Uplink commands:

i. Abor_(Also Backup for SIV-B/CSM Separation Start)

2.

3o

4.

5.

6.

7.

8.

9.

i0.

ii

12

13

14.

15

16

17.

18.

19.

20.

21.

22-23

24-25

26-27

28-29

30-31

32-33

LET Jettison Start-Backup to onboard command fromS -IVB.

Thrust off- Turn off SPS engine; backup to onboard com-mand in case of malfunction.

Thrust On - Turns on SPS engine; backup to onboard com-mand in case of malfunction.

CM/SM Separation - Backup to onboard command fromthe G&N.

Lifting Entry - Necessary for no-roll entry in the SCSentry mode.

G&N Failure - Backup to G&N function.

G&N Failure Inhibit - Reset G&N failure.

Roll Rate Gyro Backup - Switches roll BMAG to rate mode

and uses this gyro for roll rate data.

Yaw Rate Gyro Backup - Switches yaw BMAG to rate

mode and uses this gyro for yaw data.

Pitch Rate Gyro Backup - Switches pitch BMAG to rate

mode and uses this gyro for pitch rate data.

Roll A and C Channel Disable - Disables the automatic

A and C RCS roll channels.

Roll B and D Channel Disable - Disables the automatic

B and D RCS roll channels.

Pitch Channel Disable - Disables the automatic pitchRCS channels.

Yaw Channel Disable - Disables the automatic yaw RCSchannels.

-Direct rotation + pitch

-Direct rotation - pitch

-Direct rotation + yaw

-Direct rotation - yaw

-Direct rotation + roll

-Direct rotation - roll

SM _uad A Propellant Off/On.

SM Quad B Propellant Off/On.

SM Quad C Propellant Off/On.

SM Quad D Propellant Off/On.

CM System A Propellant Off/On.

CM System B Propellant Off/On.

3-7

34,

35-37

38

39-40

41

42 -45

Direct Ullage (Also sets SCS AV mode)

Fuel Cell Purge (cell #1 - cell #3)

FDAI align

T/C Antenna Switch (-Z, +Z)

UDL/S-Band Comm. Switch

Cryogenic Heater Fan Switch (02#1, 02#2 H2#1, H2#2)

Commands 12-21 will be used to control S/C attitude in cases

where the G&N is not operable.

Of these commands only six are intimately concerned with

G&N operation; abort, thrust off, thrust on, G&N Failure Inhibit,

G&N Fail, Direct Ullage.

Abort: As discussed above, this abort command may be accom-

panied by an abort command to the AGC via AGC Uplink.

Thrust

Off:

The ground may thus inhibit starting of or may stop the SPS

thrust. Should AGC-controlled firing be inhibited or shut-

down the AV monitor logic would after i0 seconds exit from

thrust vector control and hold attitude until the free-fall

interrupt occurs.

Thrust

On:

AGC Engine On logic presently includes a monitor of AV to

ensure engine ignition. This monitor continues for i0 sec

after sensing no thrust during which time the ground might

start the SPS engine. If suitable AV has not been sensed

after i0 seconds the AGC would exit from thrust vector con-

trol and hold attitude until the free-fall interrupt occurs.

Should the ground successfully start the engine within i0 sec

the AGC will guide the burn normally. It must be assumed

however that as the AGC Engine On command did not work

correctly, AGC Engine Off will not either. The ground must

therefore command a timely "Thrust Off" compatible with

the AGC TVC calculations.

G&N

Failure:

This command is a ground backup for the G&N originated

command. As all control of the vehicle by G&N is thereby

inhibited, the resulting confusion of the AGC (it has no know-

ledge of the ground command) is interesting but irrelevant.

Direct

Ullage:

A backup command for ground use during a ground controlled

burn in the SCS AV mode. Its use during G&N controlled

flight would inhibit G&N attitude control with the possibility

of the G&N being unaware of the control loss.

G&N

Failure

Inhibit:

This command overrides the G&N FAIL signal.

3-8

3.2

3.1.3 BackupControl SystemsConstraintson G&NOperation3. I. 3. 1 BackupAttitude ReferenceSystem

The backupattitude reference system is _,heSCSBMAGsinconjunctionwith AGCU. G&Ncontrol of the CSMorientation is alwaysdonewith considerationfor the maintenanceandaccuracy of this sys-time. As the SCSsystemis presently designed,the BMAG's operateas free gyros in the C_N AV MODE; in other modes they are caged

through the AGCU.

As the mechanical stops of the BMAG's are at ±17 ° it is

apparent that during boost (MONITOR MODE) and attitude maneuvers

(G&N ATTITUDE CONTROL OR ENTRY MODES) both involving angular

changes of over 17 ° the BMAG's must be caged. In the G&N AV mode

however, attitude changes over 17 ° might occur.

The rate limits of the backup attitude reference system in the

caged mode are 5°/sec in Pitch and Yaw and 20°/sec in Roll. To pre-

clude controlling the S/C at rates beyond which the backup attitude

reference system can maintain its reference, the G&N will limit its

command rate to the CSM.

3. I. 3.2 Backup Entry Control

During the pre-entry coast the G&N system must orient

the CM for aerodynamic trim and lift vector down. Then, in the event

of G&N FAIL INDICATION, the MCP/SCS will hold this attitude until it

senses a prescribed "g" level at which time it will command a continuous

roll angular velocity.

Normal and Abort Mission Logic

The following pages describe the timeline and logic for AGC control of the

spacecraft.

3.2.1

3.2.2

Normal Sequence of Events

AGC Program Logic

3-9

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3-27

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3-28

3.2.2 AGC Program Logic, Mission 202

The following diagrams illustrate the AGC logic for Mission 202.

Each group of logical decisions and/or computations enclosed within a

dotted line represents a routine. Routines marked * are performed at

a specified time under AGC waitlist control. Such routines cannot be

interrupted by an other AGC activity and will run to completion.

The terminology used is defined as follows:

Call - Cause a specified routine (an AGC "waitlist task") to be

started at a specific time. Waitlist tasks are designated *

(a task called once at a specified time) or ** (a task

which, once called, continues until terminated)

Go To - Branch to another part of the program without return

Do - Branch to a routine with a return to the next operation in

sequence.

Set - A permanent change of state of a flag or register valid until

being reset.

T - Present time.

Store - Store indicated quantity in erasable for future reference.

3 -29

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3-30

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3-31

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3-32

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3-33

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3-34

4. GUIDANCE EQUATIONS FOR CSM

4.1 Powered Flight Guidance Scheme

The guidance scheme for Mission 202 is the same as that planned for

all Apollo CSM powered flights. It is based on the possibility of an analyti-

cal description of a required velocity (v r) which is defined as the velocity

required at the present position r, in order to achieve the stated objective of

a particular powered flight maneuver.

If v is the present velocity, then the velocity to be gained (Vg) is givenby

V -- V--g --r

Differentiation of both sides yields

where

- vu

--v g-a T--r

= b a T

(i)

(2)

(3)

(4)

b = v - gur

and g is the gravitational acceleration.

The steering command is developed by formulating a desired thrust

acceleration (aTD) as that which satisfies the equation

aTD* v = cb * v--g -- --g

where c is a constant scalar.

(5)

(6)

Indicates vector cross product

4-1

aTHencea measureof the error betweena_TD

is givenby

©

V * m

c IVgl

and the actual acceleration

(7)

where

e

m = cb - a T (8)

It can be verified that ¢0 is also the axis about which the thrust vector should-- e

be rotated to null the error. Hence co is used as the steering command.--C

Once a required velocity v r is defined satisfactorily, the procedure for

the generation of the steering command _--c is the same for all phases of

towered flight. The equations for the required velocity for the various

hases are described in the succeeding pages. Descriptions of the initial

alignment procedure, ignition and cutoff logic and implementation in AGC

are also included.

4.2 Nominal Mission

4.2.1 Required Velocity

The required velocity for the first and second burns of the

nominal mission is defined as that velocity which will put the vehicle

in an elliptical trajectory of predefined parameters (semi major axis

a, and eccentricity e). The values used are

First Burn Second Burn

a 2.22806 X 107 2.82776 x 107

e 0. 102415 0. 252865

These numbers correspond to the trajectory described in Section 5.

The value of c inEq. (6) is 1.

The required velocity can be written as

v = i + iH v H--r --r v rad(9)

4-2

where

.vrad: 2 1,I] (i0)

(11)

and

p = a (I - e 2)

r

i ---r

Irl

= * i )i H UNIT (i N --,

(12)

(13)

(14)

The positive sign is used in Eq. (i0) for the radial velocity during first

burn and the negative sign is used during second burn.

4.2.2 Yaw Steering

Plane control during the nominal mission is achieved by spec-

ifying the normal (i N) to the required plane appearing in Eq. (14). The

required trajectory plane is defined to be the plane containing the pres-

ent position vector (r_) and the landing site vector (rLs; 14.9N, 165.6E)

at the nominal time (5090 sec) of landing and is given by

i N _ UNIT (r * rLS) Sign [(r * [LS ) -iw] (15)

where i is the earth's polar unit vector. At cutoff the vehicle veloc---W

ity will be equal to v r, thereby ensuring the trajectory plane to be i N

according to Eqs. (9) and (14).

During the third and fourth burns, no computations are made

for v . The desired thrust direction is held fixed at the direction--r

computed at the end of the second burn.

4.2.3 Engine Ignition

In the nominal mission the engine is always ignited after a

fixed interval of time from a previous event. The first burn is

4-3

initiated 12.7 seconds after receipt of SIV-B/CSM separation signal,

the second burn 3041 seconds after first burn cutoff, the third

burn i0 seconds after second burn cutoff and the fourth burn

i0 seconds after third burn cutoff.

4.2.4 Engine Cutoff

During all the burns a time to cutoff (T g) is continuously be-

ing estimated from the equation

Tg = [Vg/[aTI

The accuracy of Tgincreases as Tg-*0, because aSlVg [ -*0,Ibis-0.

For the first burn, when T falls, for the first time, belowg

the computational repetitive interval of At, the clock is set to turn off

the engine T seconds later. The second burn is turned off the mo-g

ment T falls below 6 seconds for the first time.g

In the third and fourth burns the engine is turned off 3 seconds

after ignition.

4.3 Aborts During Boost

The guidance equations for aborts during boost have been designed to

meet the following constraints that have been imposed on the spacecraft atti-

tude.

The visual horizon is to be kept on a hairline on the forward window

during the entire powered flight and this line should be independent of the

time at which abort is initiated.

The window geometry indicates that this requires the thrust direction

to be between 4 ° and 36 ° to the line of sight to the visual horizon. Within

this limitation, the larger the angle, the greater is the interval of time be-

fore nominal SIV-B cutoff during which the capability exists to reach a par-

ticular recovery area in the event of an abort. Hence a thrust angle of 35 °

to the line of sight to the horizon is used. (See Fig. 4.1 ).

4.3.1 Required Velocity

The definition of a required velocity, in the usual sense, con-

sistent with the direction of thrust pre-speeifled as above, is not pos-

sible.

(16)

Hence, a pseudo required velocity is defined for aborts, which,

4-4

i_._ _I-

I0-_- 0_- I ,-- oz" I N

....._ _-_%._....

Z _0 -N _

0'T

"r

zoN

0

,.J

@

I

/ I

/

0

0

.r-I

!

. r-q

4-5

when incorporated into the general steering scheme, will satisfy not

only the constraint on the thrust direction but also permit recovery

from a specified landing area.

Let r be the entry position (300,000 ft) eorresponding to a--e

free fall from the present position. Then we can write

and

sin @f

r e cot 7 + r cot 7 e

r - re

2x

x2+ 1

2x -i

cos 8f = 2x +1

(17)

(18)

(19)

(20)

where

cot 7

v.i-- --r

v-i H

cot 7 e = r/p [e 2 -(r_e--e- 1) 2]1/2

(2i)

(22)

' * iI H' = ip --r

(23)

i = UNIT (r * v)--p

(24)

Of is the free-fall central angle to the entry point,

r is the radius at 300,000 ft altitude,e

7 e is the flight path angle w. r.t. the local vertical at entry

is the present flight path angle (w. r.t. vertical)

The entry-point is given by

r = r (i r cos ef+i H, sin Of)--e e(25)

4-6

Now, let r T be the desired entry point (target vector). The error d

can be written as

d°IrT re l (26)

The target vector is the inertial position of 14, 3926°N latitude

and 313 ° longitude at 975 seconds from lift-off. This choice corre-

sponds to minimum plane change for aborts at 617.4 seconds from the

nominal boost trajectory.

The rate of change of this error is computed by differencing

r as--e

(27)

Ire - r I /Atn --en-1

(28)

where the subscript n denotes the nth computational repetition

Observing that d/d is a measure of the time to cutoff (Tg) and

that Tg according to Eq. (16) iSIvg I / IaT lin the general scheme, the

magnitude of v is defined as--g

or

Izgl= d I TId

Izgl-- __-ald L_v[

(29)

(30)

where Avis the velocity increment measured with the aceelerometers

in the interval At.

Now consider Eq. (6). Set c = 0; then

_ = 0aTD * Vg

(31)

If the direction of v is chosen as the desired and known direction of--g

a T , the specified constraint on the spacecraft attitude will be satisfied.

4-7

Figure 4-I shows the geometry of the spacecraft window.

The angle ¢ between the thrust and r is given by

RvhI (32)

where e is the specified angle (35 °) to the horizon and Rvh

to the visual horizon.

From Eq. (32) and Eq. (30) we can define v as,--g

d JA___vJ

Vg-- _-d (-cos fi r + sin_ ill,)

is the radius

(33)

4.3.2 Yaw and Roll Steering

The development of Eq. (33) is based oni r andiH, which are

both in the present trajectory plane according to Eq. (23). However,

normally, a plane change will be required to reach the same landing

site from different points of aborts on the boost trajectory.

Let the plane containing the present position r and the target

vector (See Section 4.3. l) [T be defined by

IN= UNIT (r * r T) Sign [(r_ * rT).lw] (34)

The velocity increment along_ip (normal to v_) to null the error

between i andi N is given by (See Fig. 4-2).--p

Av N '= [vi (ip * i N ). i(35)

The acceleration along ip required to accomplish the plane

change is given by

aN = i--p

Av N

T + 6g

(36)

where 6 is a small scalar (5 seconds). In order to prevent large yaw

rate commands, a limit of 5 ft/sec 2 is imposed on I aN - aN n 1 I"n -

4-8

as

Equation (33) can be now modified, to include yaw steering,

v = iT I vl--g Ad(37)

where

i T =[UNIT - i_r cos_b + UNIT (ill,a T +an)sin _] (38)

and aT is the magnitude of the thrust acceleration.

The required velocity is given by

V = V + V--r -- --g

(39)

where v is given by Eq. (37). With the required velocity so computed--g

and with e = 0, the same steering (Eq. 6) as for the nominal mission

is used.

The rate command resulting from the required velocity v has--ronly pitch and yaw components. However, the vehicle must be rolled

such that the pitch axis is in the horizontal plane (See Fig. 4-1) This

is achieved by generating a roll command (_R) from

_R =-(-JR ipitch) iroll (39a)

The negative sign is the result of the desired orientation in which the

spacecraft z - axis is pointed up.

The roll rate command is added to the rate command genera-

ted from Eq. (7). Note that the roll rate computed according to Eq. (39a)

must not be commanded unless the spacecraft z-axis is within 90° of

local vertieal.

4.3.3 Engine Ignition

In the ease of a non-tumbling abort the engine is ignited 2.5

sees after receipt of the SIV-B/CSM separation signal.

Iftumbling has been detected by the time the separation signal

is received, the engine is ignited 2.5 sees later and is shut down when

tumbling has been arrested. Ifthe capability of landing area control

4-9

-'r

HORIZONTALPLANE

_.___ i4IN

- ,_.p

_ _iT

ff

Fig. 4-2 Computation of a n andi t

4-10

|exists, the engine is re-ignited after a time interval calculated to be

sufficient to orient to the desired initial thrust direction.

4.3.4 Engine Cutoff

When T falls below At, the clock is set to turn off the engineg

T seconds later under normal area control. However, the engine willg

be turned off if any one of the following violations has occurred before

T < At.g

a) Free-fall time to 300,000ft is below 160seconds

b) --eris beyondr T. That is,

r. r e < r. r r(40)

I

ft should be pointed out that the estimate of T is very poor ing

the early part of the burn for long burns. Hence its value at ignition

cannot be used in back-up systems.

4.4 AGC Computations

Since the information about the thrust acceleration comes from the ac-

celerometers in the form of velocity increments (Av), the computations in

the AGC are in terms of increments of velocity rather than instantaneous

acceleration. The repetetive guidance computations are shown in the form

of a block diagram in Fig. 4-3. The computational blocks are common to all

powered flight maneuvers except the computation of v described in the pre---r

ceeding sections.

4.4.1 Average g Equations

The vector position and velocity are updated in each computa-

tional cycle with a set of equations based on the average gravitational

acceleration written as

gn- %itr n

n

(46)

g •+

g+ --n-_ --n At + Av (47)

v n= v n-i2

4-11

>I

U

31

z,,, 0

_I "_

ouJ

IOl [:

O

©_9_DC9

°_

_9

©

h_

c9o

!

.r-i

4-12

and

(r = + At + At + --

--n rn-1 Vn-1 gn-1 2 2

where the subscript n denotes the nth computational repetition.

4.4.2 Steering Command

The vector bwas defined in Eq. (5) as

b = v -g

(48)

(5)

In the AGC (as shown in Fig. 4-3), the increment (b At) is

computed as

bAt _ Av - gAt

Then the steering command in Eq. (7) can be written as

(49)

where

v * Am-g

A0 - At

--c Iv_llZXm I(50)

A0 = co ZXt (51)--e --c

Am = c b At - A___v (52)

4. 4. 3 Orbital Integration Equations

Position and velocity during the free-fall phases of the mission

are calculated by a direct numerical integration of the equations of

motion. Since the disturbing accelerations are small the technique of

differential acceleration due to Encke is mechanized in the AGC, as

described in MIT Report R-467, The Compleat Sunrise.

4.5 Initial Thrust Alignment

Before the engine is ignited for any particular maneuver, the vehicle

should be oriented so that on ignition the thrust is in the desired direction at

that point. Since the time of ignition is known beforehand, the position and

velocity at ignition can be computed prior to the arrival of the vehicle at that

4-13

point. By integrating over At seconds from that point, the vectors v and--g

bat can be computed as shown in Fig. 4-3.

The desired thrust direction can be now calculated (prior to arrival at

the ignition point) as

i T = UNIT (cb + (q - i • cb) i- --g -- --g

where

i = UNIT (Vg)--g

q = (aT 2 - (cb) 2 + (_.ig • cb_)2) 1/2

(53)

(54)

(55)

and a T is an estimate of the magnitude of the thrust acceleration.

Once_i T is computed from Eq. (53), the vehicle is oriented prior to

arrival at the ignition point such that the thrust axis is alongi T.

4. 6 Entry Mode

Included in this section is a set of flow charts that describe the logic

and equations that control the entry vehicle. Figure 4.4 shows the overall

picture of the sequence of operations during entry. Each block in Figure 4.4

is described in detail in subsequent charts. Table 4-1 defines symbols which

represent computed variables stored in erasable memory. The value and

definition of constants is given in Section 5.

Every pass through the entry equations, which is now anticipated to be

once every 2 seconds, is begun with the section called navigation. (See

Figure 4.5). This integrates to determine the vehicles new position and

velocity vector. This sub-routine is used by other phases than entry and

will probably be operated during all of flight 202.

Next, the targeting is done. This updates the desired landing site

position vector and computes some quantities based on the vehicles position

and velocity and the position of the landing site. (See Figure 4.6).

The next sequence of calculations is dependent upon the phase of the

entry trajectory that is currently being flown. First is the initial roll angle

computation. (See Figure 4.7). This merely adjusts the initial roll angle

(180 ° for flight 202 is now planned) and tests when to start the next phase.

4 -14

Fig. 4-4 Re-Entry Steering

4-15

@

vEs_

-Icr= u_ rr (_.'), _z

I:_ E_AD & CLE/_I_,,,P|P/_$

5_VE VEL. INC_:_M_.NT

AT _a_ 2

_. - _-_ ZX-T -_-_

_j_I_.T _ _IkV IG/kT I O_

Li

1TE/x,,', + ,c_-r- _ + ix-t- _ J

T 2 I

I

Fig. 4-5 Re-Entry Steering - Navigation

i

4-16

I v:_,_,w,_ _v_ IvsQ: vZ/v s_,--F _ _ IROOT: _. Wa_T (_a} .I

LE(:_-- (,VSG-_-I') _S,

DC)LD = D

D = ABVAL (._') /AT

[RELVELSW: 1 It

t

I LATAIx3G = UI_ IT ( __.T _) • _ N

I'T._rA: co_ -_ (,U_,T(_-T_-U_,T(_./)t

t

Fig. 4-6 Re-Entry Steering - Targetting

4-17

N'E_ NO 1

YEs I e,aO

I T_.C)LL C.-- C_ IC:)

I._=° I!

o',-+v _.o,..,',-,-..o,_N_:_

I_°_°_°II_:_ II GO -TO _40_r-_S -T I

Fig. 4-7 Re-Entry Steering- Initial Roll

4-18

The next phase maintains a constant drag trajectory while testing to

see if it is time to go into the up-control phase. The testing is presented

in Figures 4.8 and 4.9. The eonstant drag equations are given in Figure 4. i0.

The other phases (up-control, ballistic and final) are listed in Figure 4. ii,

4.12 and 4.13. The final phase is aeeornplished by a stored reference

trajectory. Its eharaeteristies as well as the steering gains are stored as

shown in Figure 4. 14o The routine that prevents excessive acceleration

build-up (G limiter) is given in Figure 4.15. And finally, the section that

does the lateral logic calculations and computes the commanded roll angle

is shown in Figure 4. 16.

4-19

S _4UI,4"I'INC) _'b,l_

lax-aa lT

(1_ DO -- C'9 NEG'_

k) o Y_S_

J

ALP = 2_C)

VL = V.%

_:EPLE_. _.AP,_G.F-- CALCt_ L_-r'l_l_

SlIJ_ = S'l_ (I_.D_TL/VL.) z

I

iS VL--VLlV_I_'-J k3E.C3 _"_--_0Y_# .

tI_,SV-Ep : V-EPLE_ RANGE I

_, _.{7 V_ / _J_..I

I AsP 3 : q5 ((_6-r_PCrTL), GI%I_N_ CO_.R._X_

_ i_ Tt-4ETI_/_--A5 P SJE_. _l

Fig. 4-8 Re-Entry Steering - Huntest

4-20

ILv=,,_,T.<_o.__oo-,_/V,,,_,-_Ii

)

I,,,=w+ v _o_. i v_ =vl+ ,ooo Iv,_.v,.+vco,_l lGo To _O_'re..ST_i

'q "4l v_ = v, -vco_/=_l

FACT z: _,'--_(_P-,)/Ao_,EL.IE::C.'T¢ot_-... --- _-)¢::_C-O)_"T'_.OL..

Go To UPc_.OK.)TI:_::_.

Fig. 4-9 Re-Entry Steering = End Huntest

4-21

i

i

Fig. 4-I0 Re-Entry Steering- CONSTD

4-22

Go "TO _'J_-_WES'-_ I

Fig. 4-1 1 Re-Entry Steering - UP CONTRL

4-23

-T N -If VX

EGsw = I I P,'_H " _ k_1

SELECTo_x= P_EDIC_r

Y_,w = 0

Fig. 4- 12 Re-Entry Steering - Ballistic

4-24

(DOLD- D)

(DOLD* _') /e.

PI_EDAN C-:-:-bL= _TO_O (V') ¢ F2 (V)(DADVt - DAI::)VR.(V)_

+_-_ (v)(D-D_E_-_(V))

L/D = LO D + 4- (T,ETNN,, -- pIP_,._= DAI',aGL) _y (. V)

Fig. 4-13 Re-Entry Steering- Predict 3

4-25

p_o

p_

n4

(D

_Z0 e4 O0 C'Q _ D-- O_ ,-_ ¢Q c'q co LO _

(J

I" I" I 16 I" n° l" I _', , , , _, _

A

•"4 I I I I

_J

0cO

¢)

c_

,--Ic_

.r-I

1--II

.,-I

4-26

IS GN_,A>(.-_:::_ PO_:_

yES_

IGc>"to3,oI

-,(aH_ _,,,,_,xl,,) _" I

I_,-,_R.,-v_.QI

_)lO(D

IL/D--O.Z (R DOTC_- - R_Z:_C:_-")ID -

I_°_°_,°I,

Fig. 4-15 Re-Entry Steering - GLIMITER

4-27

NO

Fig. 4- 16 Re-Entry Steering - Lateral Logic

4-28

TABLE 4-i

VAR_BLES FOR RE-ENTRY CONTROL

q, T_

!'7

\/

O

I

OT_"

!,T _

mT

'!_,! f

_,L._

A gK FD

Eqr_l

^ SD! 'P

r_

m_

D_FF

F2

r4CT!

F._r T2

v 2or'Ll__

I.^TANC

LF©

Lrhr

LV

OP,_T

Prh_TR_F

OrJL t..r"

r_D_TL

T

Tl-lc T A

TM_'- TN M

\/cOpo

V

Vt

VL

VP='F

VBAPS

V£1")

!aT

Y

HUNT[Nh

HIMD

qFLV_!_ gW

_GSVl

V-"L _r T TY V_CTC 'r

or_c T T T _,_ V_:qTpO

\l_CTqP F#.CT A T TNTTTA L T^_..r-_T

Mt_O_AI_ T _ DT c ^n, lh II z

T ,_._r: _"T VCr- TOo

IJ_uIT k'C)'_,;AI. TC TP,_.JF('TrhPY PLY, _r

T_IITIAI _ePC, F:©o L.:DC_'T_t_

r'n,',_qT FOP UPC*_'Tr_[

_rPLrP _AN_F"

FT_.'AL PH, ASF PANG c

GAM_4A ('©OPFCTTPL'

Om_DICTEF _AN rq_ = _SKEP+ACD]+ASPLJP+,_SP3

TqTAt. ECCCl =re#Tit')K,

cml,,TOrH I.. _'r'. cr_4CT mm,_C

_CFEP_.'C c D DP,_G/DV (_Tn!aL p_^SF)

RFFSRF_ CE [;RAG

r-qCCNTq lqI TY

r)_ANG. F/C DPh, G (FINAL PP'AS_)

,m_ f, _ r- P / _A O r_ T "_c, (F ! N_ L PH ,_.qc)..

r_*,_CT _0 ° tJDC©_.ITRL

r_!qT FC.o LIDr_'T_I.

I,! C I: Im t_,l I ID (- f",_, T _ I

I_'PTCATO9 Fr'm '_r'[! c','TT(-U (])

I ATFPAI_ _,':tN,'-_

-_XCFSS C,F. qvFr_ C.m,:V =lV_r__l )re

l_/r_ C_OPCTT _ FP_I-" i:r_-r_,'TOl._

h _r t.Mr- m / r,V

OP_I'-TF;-; Pa,_Ii_V (FTNaL pH_SF)

A!..T [ Tr'_F D':TF

D-'zI:_:D'-z'_,.I,r'P ,2..lh (q _("D ..D'- _1_ T q I..

_XTT P'b0T F__P 'D(-C'*'Tr:'i

RAK, G,F TO _r, (FT_.'AL a_-,'.S;l

T f Mr

V_.LOCITy CO_PFCT'[,O_,I _'__o UI:_r"('MTOL.

Vr[.r_q T Ty H ' _/u ! Tt !F_":

INITTt. L VFL:_,C]TY #fhq IIp,[,,'!Tfl_

taXi T VrL_CTTy F,_ _iP,",'t, tT'_".

RF-FFr_m_,CF VTL'D, CTTY FC': ',JPCr'_,T_i.

2 2

Vl_ /\/c,& T

2 ?

N_F_MALI qpr_ V_'-i_C'CTT Y _'::'t_AO::i "_ = 'v' ',,'Ct, T

#'A PTH _ t_T _ X TTM"

c,!,.dTTC# TO C, FLFCT C(_K,!CT I_ L_Vr--I. r',

SWITCH TO ITFR!_'TF tN {-;I_NT[ CT r

_L VFLOC!TY F'v,'!Tq'-' 0

qW ] ] :-H :!

4 -29

5. MISSION AND VEHICLE DATA

5.1 Scope

Section 5 is a summary of all Flight 202 mission and vehicle data that have

an impact on AGC programming. Data have been collected under the following

headings:

Section 5.2 Mission Data. Establishes the outlines of the mission in terms

of trajectories, profiles etc. Includes performance figures for Saturn boost phase

inasmuch as they affect conditions pertaining at take-over of control by G&N system.

Section 5.3 Memory Data. Contains all mission- and vehicle-dependent data

that are, in one form or another, written directly into the memory of the AGC. In

a wired-memory computer such as the AGC, the very limited erasable section is

intended primarily for storage 6f computational variables. An attempt has been

made to consign those mission parameters that do not change during flight to the

fixed section of the memory. Some exceptions have had to be made in the case of

the Saturn boost polynomials and SPS aim-point criteria, since these will not be

available until shortly before the flight.

Section 5.4 Vehicle Data. Contains information that will mainly affect simula-

tions and rope verification and will not, with only one or two exceptions, appear

directly in the AGC program.

Section 5.5 Physical Constants. These definitions will be used in AGC programs

and verification work.

Numerical data are presented in the most convenient and widely accepted units.

The AGC is, however, programmed in the metric set of kilogram, meter, and

centisecond (10 -2 sec). Conversion to other sets of units is done by use of the factors

defined in Section 5.5.2.

Points on the surface of the earth are defined in terms of geodetic latitude and

longitude referred to the Fischer ellipsoid of 1960, and geocentric radius.

5-1

5.2 Mission Data

5.2.1 Mission Trajectories

Referencetrajectory(Saturnboost, SPS1,coast, SPS2)Referenceentry trajectory(Pre-05g to touch-down)Nominal mission profileMajor eventsduring nominal missionNominal Saturnboostprofile

inot available

not available 1

see Fig. 5.1

see Table 5.1

see Fig. 5.2

Note I. No officialtrajectories issued. All dependent data in this report are derived

from MSC SA-202 optimum trajectory dated September 1964, and an undated

entry reference trajectory communicated on ii November 1964.

5-2

Z z z

J

o

•" , ." "::7..

4

S

;

- f

Z

o

t

/

i

.I

./- ,/

/./

\\

\.\.

e.¢

//

i

ill

I =,i_-I

L

/

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/u_

_J

uJ

uu

oo o °o _ oo o°o o o_ oo oooo o o° o o

(133-I) 3(:] f11117_ I

I_l_lrll _l lli_illl]i lilllii lili_lllll IIIIIIII Illlliii lill

°o o° o o °o °o o0 0

" _

(SON033S) 31NI1

H

0

@g.

0

_ow3I---I

- LO

u_

I

. . • .Im,_lll_l "

I 1 I I I I

(93G) IX0

I Io _

I i ' /I d

I I I I I0 0

o o,0

I I I I I8 o

I I I I I

I I I I I

I I I I I I I0 0 0 00., _ o o

(:)3s) 3wllI I I I I I I

0 0 0 00 0 0 0

( J.:l 000'001. NI ) q

I I I I I I I

(Sd:l 0001 NI) _'A

I I I I I I I0 0 0 0

( 1-1 000'00£ 01 -]Wll 11V=1:13;,1:1 ) _1

I I I I I I I I_, :_ _0 ® o ,,, ., .o

(z3_lS/l:l) xV

m

I o

I

I

I

Iao o

©

0

c_

E_

0©_Q

cdo9

c_l

LE3

.t--I

5-5

TABLE 5-I. MAJOR EVENTS, MISSION 202

_-VENT

Lift- off

SIB c/o

SIVB Ign

LES Jett

iSIVB c/(

SPS Ign.

SPS c/o

Apogee

Ullage

SPS Ign.

SPS c/o

SPS Ign.

SPS c/o

SPS Ign.

SPS c/O

Entry

End of

Entry

t , V i

(sec) ( fps_

0

145.5

151

161

617.4

628.4""

882.0

12528.

3893.4

3923.4

4014.

4024.

7,022

6,953!

7,037

121,848

21,834

I

.25,632

122,713!

!24 837I

124,923

i127,739

27,647I

4o27 27,751

037 j27,789 039. £'L27,878

i319.6 28,690

5087.6 1,791

AZi

(deg)

o 9o. ooi

24.02 i02.43i

I22.80 102.49

21.01 102.66

2. 635_112.05

2. 395}112.31i

5.77 Ll17.70

-0.006105.89

-5.80

-5.82

-7. 67

-7.53

-7.56

-7.44

-7.46

-3.51

64.61

63.99

62.18

ALT. N. GEOD.

LAT.

0 28.53

175,508 28.23

190,757 28.21

216,802 28.17

544,569! 23.43

554,8931 23.20

918,95_ 16.46

3,606,67@27.99i

1,620,50Q-19.95

1_43, 64fl-19. 24t

1243,964 -16.26!

1218,301 -16.15

1,207,390 -16.05

1,171,147 -15.72

1,161,917 -15.60

397,603- 5.16

49,613 14.9

E. LONG:WEIGHT

1

279.42 1,322 530

279.98! 434,058

280.06 320,240

280.21 306,928

294.61 83,099

295.18 45,701

309.35 27,956

29.56 27,956

109.25 27,942

110.95 27,942

116.48 21,589

116.63 21,589

116.82 21,379

117.43 21,379

117.59 21,200

134.85

165.6

End of Entry for No SPS AV1469.4

End of Entry for No SPS &V 24977.

9.25 320.6

4.7 148.4

Data from this time on is from MIT 202 performance simulations

Fourth Burn cutoff due to propellant depletion in this simulation

5-6

5.2.2 Nominal SIVBSeparationAttitude Conditions

X-axis in planeof maneuver, forward of localvertical by

(Y-axis alongmomentumvector R* V

Z-axis abovelocal horizontal)

Roll rate

Pitch rate

Yaw rate

67.20°

0°/sec

0°/sec

0°/sec

5-7

5.2.3 3a Dispersions from Nominal at SIVB Separation

X-axis attitude dispersion 2°

Y-axis attitude dispersion 2°

Z-axis attitude dispersion 2°

Roll rate residual 0.2°/sec

Pitch rate residual O. 2°/sec

Yaw rate residual O. 2°/sec

5-8

5.2.4 SIVBEngine-off Transient

Decaytime 100%-10%

Decaytime 10%-0%

Tail-off impulse 100%-10%

Tail-off impulse 10%-0%

not available

not available

not available

not available

Q

5-9

5.3 Memory Data

5. 3. 1 Prelaunch

Launch position: Latitude

Longitude

Radius

Inertial reference plane (IMU) azimuth

Memory

Type

F

F

F

F

Value

28.53253 ° N

279.41701°E

6,373,305.2meters

105.0000 ° E of N

O

5-10

5.3.2 Saturn Boost

Roll polynomial coefficient (s)

Pitch polynomial coefficient(s )

Heading polynomial coefficient(s )

Interval: Lift-off-SI attitude monitor

terminate

Interval: Lift-off-LET jetison

assumed complete

Memory

Type

E

E

E

F

F

Value

not available

not available

not available

150 sec

171 sec

5-11

5.3.3 Attitude Maneuvers

MemoryType

Value

Limit: commanded S/C angular

rate:

Roll (CSM) F 7.2°/sec

Roll (CM only) F 15°/sec

Pitch, Yaw (CSM, CM) F 4°/sec

Interval between attitude updates (CSM)F 1 sec

(CM only) F 0.5 sec

Interval for stabilization after

maneuver

(CSM) F 6 sec

(CM) F 3 sec

5-12

5.3.4 TVC (Normal mission)

Memory

Type

CSM c.g. displacement in X-Y plane:

(SPS 1 ) F

CSM c.g. displacement in X-Y plane:

(SPS 2) F

CSM c.g. displacement in X-Y plane:

(SPS 3) F

CSM c.g. displacement in X-Z plane:

(SPS I) F

CSM c.g. displacement in X-Z plane:

(SPS 2) F

CSM c.g. displacement in X-Z plane:

(SPS 3) F

Mass loss rate of SPS engine F

Initial mass of CSM + propellants F

Tailoff impulse (mean) of SPS engine F

Minimum AV criterion for thrust monitor F

Interval for thrust monitor F

Interval between steering updates F

Steer law gain F

Steer law velocity bias F

Steer law coefficient (C) F

Interval: freeze CDUs to engine-off command F

Value

6.60 ° 1

3.35 ° 1

0.30 ° 1

2.25 ° I

0.60 ° 1

-0.70 ° 1

2. I75 slug/sec

1,428 slugs

8,400 lb-sec

1 ft/s/s

10 sec

1 sec

0.25

160 ft/sec

1.0

1.8 sec

Interval: SIVB/CSM Sep. - SPS 1 ignition F 12.7 sec

Interval: SPS 1 cut-off - SPS 2 ignition E 3041 sec

Interval: SPS 2, 3 cut-off - SPS 3, 4 ignition F I0 sec

Note 1: Figures derived from data in Section 5.4.1 using weight data in Table 5-1.

5-13

Interval: SPS 3, 4 ignition - SPS 3, 4 cut-off

Interval: + X translation - SPS 2 ignition

Interval: between SCS mode change eommands

Interval: Gimbal mot. power ON - Enginestart F

SPS 1 aim-point criteria

Semi-major axis E

Eccentricity E

SPS 2 aim-point criteria:

Semi-major axis E

Eccentricity E

Interval: Lift-off - touch down. (Nominal

mission) E

Memory

Type

F

F

F

Value

3 sec

30 sec

0.25 sec

2 sec

2.22806x107ff

0.102415

2.82776x107ft

0.252865

5090 see

5-14

5. 3.5 Entry (Normal mission)

CSM attitude for SM/CM Separation:

X-axis above velocity vector by

(Y-axis along monentum vector (R * V),

Z-axis above velocity vector)

CM Pacific pre-entry attitude:

X-axis below velocity vector by

(Y-axis along momentum vector (R * V),

Z-axis below velocity rectory. A lift-

vector down attitude )

Trim angle of attack

Interval: SM/CM Sop. - start maneuver

Pacific recovery point: Latitude

Longitude

Constant on ALP

Initial shaping roll

Constant drag gain (on drag)

Constant drag gain (on RDOT)

Lead velocity for up control start

Minimum constant drag

Minimum D for lift up

Minimum drag to start Kepler

Minimum drag to end Kepler

G-limit

Minimum drag for lift up if down

Up control gain, optimized

Up control gain, optimized

Lateral switch gain

Time of flight calculation gain

5-15

Symbol

C1

C10

C16

C17

C18

C19

C20

DMIN

DMIN2

GMAX

KA

KB3

KB4

KLAT

KTETA

Memory

Type

F

F

F

F

E

E

F

F

F

F

F

F

F

F

F

F

F

F

F

F

F

Value

60 °

160 °

20 °

5 sec

14. 000°N

165. 600°E

1.25

0

0.01

0. 0002

5o0 ft/s

35 ft/s/s

200 ft/s/s

6 ft/s/s

6. 5 ft/s/s

lOg

O. 2g

O. 0034

"3. 4

O. OO75

1, 500

Max L/D

LAD cos (15°)

Up control L/D

Final phase L/D

Final phase range

Final phase dR/dV

Final phase initialvelocity

Final phase dR/dRDOT

Final phase initialRDOT

Minimum drag for up control

Minimum RDOT to close loop

Minimum VL

Normalization factor, acceleration

Atmosphere Scale Height

Normalization factor, velocity

Nominal earth's radius (entry only)

MemorySymbol Type

LAD F

L/DCMINR F

LEWD F

LOD F

Q2 F

Q3 F

Q4 F

Q5 F

Q6 F

Q7 F

VRCONTRL F

VLMIN F

GS F

HS F

VSA T F

RE F

Value

0.3

O. 2895

0. i

0.18

641 n.m.

O. 07 n. m/ft/s

23,500 ft/s

O. 3 n. m/ft/s

820 ft/s

6 ftls/s

700 ft/s

18,000 ft/s

32.2 ft/s/s

28,500 ft

25, 766. 197ft/sec

21,202, 909ft

5-16

5.3.6 TVC (Abort)

Criterion for tumbling detection

Symbol

Memory

Type Value

F not

available

Interval: SIVB/CSM Sep. - SPS ignition

(tumbling and abort) F 2.5 sec

Interval: Time-to-go bias

Interval: between steering updates

Thrust attitude:

X-axis above visual horizon by

(Y-axis normal to local vertical,

Z-axis above local horizontal)

Limit: commanded change in yaw accelera-tion

Abort alm-point: Latitude

l__ngitude

Interval: Lift-off - abort aim-point

(Abort from nominal mission (See Section 4.0))

Mean geo-centric radius of visual horizon Rvh

F 5 sec

F 2 sec

F 35 °

F

E

E

E

F

5ft/s/s

14.3926°N

313.0000°E

975 sec

not

available

5-17

5.3.7 Entry (Abort)

CM Atlantic pre-entry attitude:

X-axis above velocity vector by

(Y-axis along neg. momentum vector (V'R)

Z-axis above velocity vector

A lift-vector up attitude)

Atlantic recovery point: Latitude

Longitude

MemoryType

F

E

E

Value

160 °

not avail-

ablenot avail-able

5-18

5.3.8Free-fall time (Tf) monitor

Entry interface altitude

Abort Tf criterion (A) to start orientation

to CM/SM Separation Attitude

Normal Tf criterion (N) to start orientation

to CM/SM Separation Attitude

Interval: rain Tf to start CM/SM

Separation

Interval: between Tf updates

Memory

Type

F

F

F

F

F

Value

300, 000 ft

160 sec

160 see

75 sec

1 sec

5-19

5.4 Vehicle Data

5.4.1 CSM Data

Weight empty

Weight of initial fuel load

Variation

Variation

Variation

Variation

Variation

Variation

Variation

Variation

Variation

of principal inertia with mass

of principal inertia with mass

of principal inertia with mass

of product of inertia with mass

of product of inertia with mass

of product of inertia with mass

of C.G. X-location 2 with mass

2of C. G. Y-location with mass

of C.G. Z-location 2 with mass

MS 21, 200 lbs

ML 24, 500 lbs

IXX Defined in Fig. 5.4

IYY Defined in Fig. 5.5

IZZ Defined in Fig. 5.6

IXY Defined in Fig. 5.7

IYZ Defined in Fig. 5.8

IZX Defined in Fig. 5.9

CGX Defined in Figs.5.10, 5.11

CGY Defined in Figs.5.12, 5.13

CGZ Defined in Figs.5.14, 5. 15

MF 14.3 slugs

MO 44.6 slugs

RF 958 ins. (ApolloRef. )

RO 966 ins. (ApolloRef. )

WF 4.07 1 rad/sec

ZF .005

WO 3.82 i rad/sec

ZO .005

LT 7.1 feet

LE 833 ins. (ApolloRef. )See Fig. 5-3

Fuel equivalent slosh mass

Oxidizer equivalent slosh mass

Fuel mass C.G. X-location

Oxidizer mass C.G. X-location

Fuel mass natural frequency

Fuel mass damping ratio

Oxidizer mass natural frequency

Oxidizer mass damping ratio

RCS thruster moment arm

Engine hinge point location

Spacecraft Launch Configuration

NOTE: 1. Data corresponds to initial thrust acceleration of 20.9 ft/sec 2

(W2aT 2and the relation )t = (W a T) initial is assumed.

2. Angles given as positive rotations of ¢ngine hinge'point to c. gJ

line about positive CSM Y and Z axes.

5-20

x L - 406.5

X o - 1490.0

LAUNCH

ESCAPE

SYSTEM

COMMANDMODULE

X a = 1gO0.0 ----

Xc=0

SERVICEMODULE

l ZL

133.5"

TO HEAT SHIELD

STRUC1 URE

162.0'

i

SPACECRAFT

DIMENS IONAL

DIAGRAM

XL=G

X c = 83. S

X o = I083,5

154.0' DIAMETEr,'

X a = 833.2 ENGINE GIMBAL PLANE

SPACECRAFT

LEM ADAPTER

X a = .502.0- --

Z o

_SD_ Stabilizing _'embers

Fig. 5-3 CSM Launch Configuration

5-21

44000

40000

36000

Z

2

Z

70

_ 32OO0

28000

....................... i ......

• , .... ° .... ° ....... "_

_l:_:.,............. i.......... ,................

12000 14000 16000 18000 20000 22000

IXX ROLL MOMENT OF INERTIA (SLUG FEET SQUARED)

Fig. 5-4 IXX Moment of Inertia against CSM Weight

5 -22

44000

40000

3,5000

2

Z

320000

28000

24000

36 000 40000 44000 48000 52000 56000

IYY PITCH MOMENT OF INERTIA (SLUG-FEET SQUARED)

60000

Fig. 5-5 IYY Moment of Inertia against CSM Weight

5 -23

Ir,_lFJ.l_ : _ -.

¢'1

Z

oa.

z

o

44oo0¸ _ L I ......._- __._44I •

• I _

36000" i .foI• . , . ..., 1 I .

...... t

_--t _ /,• [ L '

ooo i...... t !:- _/4

z,ooo l ....... _ '

36 000 40000 440 O0 48000 52000 $ 6000 60000

I .....

. [ . .

IZZ YAW MOMENT OF INERTIA(SLUG FEET SQUARED)

Fig. 5-6 IZZ Moment of Inertia against CSM Weight

5 -24

44000 - .

40000

36000

z 32000

0

z

0

28000

24000

J _ ....

t ii!iliilit ........... i iii

, , _ ,

; , - ; '

44000

4200O

40000

51000

-4000 - 3000 -2000 - 1000 0

_XY PRODUCT OF INERTIA ( SLUG FEET SQUARED)

Fig. 5-7 IXY Product of Inertia against CSM Weight

5 -25

44000

40000

3_000

32000

28000

24000

1

I

-400 0 40o BCC 12oo 16o0

IYZ PRODUCT O_ i/',_RTiAISLUG-FEET _f_UAREDi

2000

Fig. 5-8 IYZ Product of Inertia against CSM Weight

5 -26

44000

40000

3O000

ZD

28000

I

" i .....

400

Fig. 5-9 IXZ Product of Inertia against CSM Weight

5 -27

-- CO:t!FZ:;_IT:,kL -

44000

40000

3_000

32000

28000

24000

94,4 948 952 956 960 964 968 972 976

X. CENTER OF GRAVITY - APOLLO S/C STATIONS

Fig. 5-10 C.g. X-Axis Coordinate against CSM Weight

5 -28

CO ..... !" : '"-" " '-

v1..--.

Z

0

z

o

1,1.1

44000

40000

36000

32000

28000

24000

I "'1';I'1;I '1 ;II I I L, 1 Ill! I i _ t t ; ; ; I 1 ] ! .....

.....I....I[.....i I....I .....!.....I I........./,

t .....[...................t......I....r....I.-::!_i,.........!'I!t! ..........l/l,_...,_....I!.....I_":,,_::::_:,:: : i .... :;t _ t_..._":::: : ......... "t

iiiit_[ii_itii itii tii i[iii;I_tiiii_tii:_tiii_{iiiiitii_itii

i ....:.... ::::! ........ _r::::":::r:"::) :l::}:I:i::t::=:I::

..... t ..... t .... t .... 1- .... t ' . • _÷ .... ,t ..... _...... _...

...... : ........... , I TTF , FT, -: _" '

..... : t : i [ ] ! i[_ } I , i [

.............. _' " ±-" _-'_ r I ! l t i T I ] I ' I ; ; ; ! i IIl ! _--i--i---;" : ' '

. .- i . • : :: : : i . ! . . . i : ! ,, ; ;.. _ .... , ] . . . , . .

. °

2 3 4 5 6

SPS TRUE GIMBAL ANGLE FROM THE X AXIS, DEGREES

Fig. 5-ii SPS True Gimbal Angle from X-Axis against CSM Weight.

5 -29

z

oa.

z

I

o

44000

40000

36000

3 2000

28000

...... I ....

.... ....

t

i

I

2_0C_

_. 6 8 10 12

CENTER O_: GRAVITY- _NCHES

Fig. 5-12 C.g. Y-Axis Coordinate against CSM Weight

5 -30

44000

40000

36000

a

Z

o

z

32000O

uJ

i - -

28000 i

1

24000

.... t .....

10 20 3.0 40 5.0 60

SPS ENGINE GIMBAL ANGLE IN X-Y PLANE, DEGREES

Fig. 5-13 SPS Gimbal Angle in X-Y Plane against CSM Weight.

5-31

I

m

44000

J ....

40000

3000O

Z

0

Z

_ 32000

0

28000

24000

-40 -3.0 -20 -I.0 0 1.0

Z CENTER OF GRAVITY- iNCHES

2.0

Fig. 5-14 C.g. Z-Axis Coordinate against CSM Weight

5 -32

44(

Z

o 32000

Z

0

28(

240

o .

' !1!! I!!!l!i!!l!!!/!!!t!!!l!!!t!!!i!!!t!!!t!!!l!!!l!!!,!-!!l!!!'f

:: :::: ....................... ' ...... ri ......t-t- r-:-::.._i_...t.::t:::l!:I I!_:]!:_t!:!t!_]_.! _!_-_.:_y_................ _ ................ Ii ..... : ......... & ...... ,•..1...... t'-' 1'-'1'"t....t,"'!....t'"t ...... 1_'.,...... t......•-..-...I......i ............:....... ....,7......i -I.i!iilii i iliiiliiiliiil iliiiliiiEiiiliiiliiilHiiliiiliiilii

•.. ........................... ._,.!._.: .... ,./,,... .... ,,.,_.. .........A '

2iiti212ilii21:i21ii21ili i1222_Z,22;_ii211ilil?iil_iti2it?]/

- • • ! .... • . .

............ • ...:-. :..._....J ...r..,..: ...,:..,....-... ....

...... i " It ..... .;I ..... I' I .... I .... 1 I"'1 .... I ............• . . ,... *...a.-. _.- .f... t...t.., t...i.., v...v...i...i...| ....

• .I...I...I... _.,...I...1...1...1...I...i..-.I .... 1...1 ..........

• il;,".,;_'/_i .... ; ;";' ;; '; .... ;; "; ....... 1"

i --._,_" ................. :. :'.l..:l::.i_..l..:i.:,i:.:i ............................ '''I''" I'''I '''I''.I...I...I .... ; • i • .

-0.4 0 0.4 0.8 1.2 1.6 2.0 2.4

SPS ENGINE GIMBAL ANGLE IN X-Z PLANE, DEGREES

Fig. 5-15 SPS Gimbal Angle in X-Z Plane against CSM Weight.

5 -33

5.4.2 SPS Engine Data

Item

Mass

Inertia (IY = IZ = IR)

Hinge to c.g. radius

Vacuum thrust

Specific impulse

Maximum start and shutdown transients

Mean thrust-off impulse

Displacement, thrust vector from engine

gimbal axes intersection

Misalignment, thrust vector from engine

mount plane normal

Symbol

ME

IR

LE

TF

ISP

Value

20 slugs

213 slug ft 2

8.0 inches

2. 19 x 104 Ibs

(+ I% after 30 sec)

+10%(-1% after 750sec)

318.7 sec (3avalue

after 750 sec)

See Fig. 5.16

8,400 Ib-sec

<0. 125 inches

<0.5 deg.

t

5-34

i __m

I I I I 1 I I 1 I o0 0 0 0 0 00 oO ,.0 _ c'_

ISA_HI (]31V_ %

O_

o°t-IO_

0"o

o_

"o

GI

o

oelb.O

'elI

.r-I

5-35

5.4. 3 TVC Autopilot Data

TVC Autopilot Data

Configurat ion

Attitude error gain

Attitude rate gain

Rate command limit

Symbol Pitch (Y) Yaw (Z) Units

Defined in Fig. 5. 17

KA I. 00 rad/rad

KR 0. 500 rad/rad/sec

L 0. 140 rad

(effectively 16°/see)

Art. rate filter lead time constant _-I

Art. rate filter lag time constant T 2

Forward filter gain KE

Commanded position breakpoint LMP(1)

Commanded position limit LMP(2)

Clutch servo amplifier gain KS

Clutch servo amp. lead time const. T 3

Clutch servo amp. lagtime const. I"4

Clutch servo current limit LMI

Clutch gain KC

Actuator moment arm RA

Clutch lead time constant T 5

Clutch lag time constant T 6

Total actuator load inertia JT

Actuator load time constant 77

A ctuator load natural freqUency WB

Actuator load damping ratio

Engine rate limit LMR

Engine position limit (pitch) LMY

Engine position limit (yaw) LMZ

Position feedback gain KD

Position pickoff frequency WD

Rate feedback gain KG

Rate pickoff frequency WC

5 -36

1.00

0. 125 sec

0. 042 sec

1.50

0. 105 rads(6 °)

0. 227 rads(13 °)

20.0 Amps/rad

0. 025 sec

0.029

0.600

3,530

0.022

0.029

281

0.150

104

0.104

0.300

±0.105

1.05

287

0. 154

81.7

0. 137

+0.218

-0.078

1.00

63.0 46.2

0.090

48,1 40.0

sec

Amps

s/amplb

feet

see

sec

slug-ft 2

tad/see

rad/see

rad/sec

rad(i6 °)

.,+12 50_

raat_4. _o /

rads/ft/sec

rads/_

rads/sec

_-+ ], +_ ,

_1 "'_

r

"il++Ii

I

Z

_0

00

<_Z *

oo

Ni,I

I'+v

I--

_,o I

t:_

@e..-i

©.,-I

04_

!

5 -37

0

Z

<

0

_T0 i

0 0

.E"au o

o

vi ii II I

p..

d

d _

d

_ 0 0 I 0

0J

.d

0 0 0

o _ ,.g

_ t:_

t'-.-o

I I, , c; o"I

<

V

0 0

o _

I i, , o" d

I

<

b-0

,4 c; dI

<

i 0i

t_

c_I

<

<

@

b_

• ¢_ o9

,-_ 0 0 =

0 _:

@

0c;

j_

0 ,_

.0G_

_r.)

N_

I:_ r./'/

._ oO

0 _

0 . _b_ 0

••,-I r/_

_t ._.-I r---I r---t

g s ss_ % N% N N

5-38

r- ----

_w<(

1U.I

I

o I

"1L

I-- I.J

_- kJ --

LU US

uJ 0_J

¢10

I

IL

III

F'I

I--

I,-- t'v

II oII

_1_1

ly

U.I

C3 _D 0

%

-1

-- i

I!

W

1!

°1

! i

-- !II UJ

--I

III

Z

_- 0

c_

c_

cio

o

o

<m

-?

.,-4

5 -39

5.4.5

Item

RCS Reaction Jet Data

Configuration

Nominal vacuum thrust

Specific impulse (steady)

Minimum impulse

Thrust rise lag

Thrust rise time constant

Thrust decay lag

Thrust decay time constant

Duration, minimum impulse

electrical signal

Units

Ibs

secs

ib-sec

millisec

millisec

millisec

millisec

millisec

Value

SM

(see Fig. 5. 19)

I00+ 5

300 + 5

0.5+0.1

<12. 5

2. 0 (exp)

<6. O

2. 0 (exp)

18.0±4.0

CM

(see Fig. 5.20)

91+3

270 + 4

2.0+0.3

<13. 0

2.0 (linear)

<4.0

5. 0 (linear)

18.0±4.0

5-40

_e v ..... %rklTl A I

TENSION TIE &

UMBILICAL (3 PLACES)

RCS ENGINE

( 4 PLACES )I

SEXTANT I -- 50"

S EXTANT TT = 70"

SEXTANT 111"=60*

SEXTANT 1_7 =50*

SEXTANT _ =70"

SEXTANT _7T =60*

12o45 ,

÷Z

I

RCS l

ENGINE 7.1El

ct.MEAN

RADIUS

T_r

EQUIPMENT BAY

Trr

(empties

first )

DIZER TANK

( empties

first)

)Xl D I ZE

Irr

(emptie

÷Y

FUEL

( emlotie

EQUIPMENT BAY

I

15'

Fig. 5-19 (1) CSM Reaction Jet Positions

5 -41

Fig. 5-19 (2) CSM Reaction Jet Positions

5 -42

F

ji!

<

O

br_

__T_ _

_z /=o- •; _, /,_

,,'" t..,_-o------

-Qo

>

0

z

02

E

\fq

\

_ Z

.zC_

.o <

Z

_ u_

) i , I t i i z • _

w

"' "_ : oO . ; z

, _;z z :_o ._ z

r_

0

-r'4

o

o

c_

r..)

o

I

5 -43

5.4.6 CM Data

Control Weight

Principal inertia (IXX)

Principal inertia (IYY)

Principal inertia (IZZ)

Product of inertias (IXY)

Product of inertias (IYZ)

Product of inertias (IXZ)

CG X-location

CG Y-location

CG Z-location

Aerodynamlc

Aerodynamlc

Aerodynamlc

Variation of coefficients with

Mach number

reference area

reference diameter

coefficients

11,000 lbs

5065.0 slug-ft 2

4491.3 slug-ft 2

3973.5 slug-ft 2

-1.7 slug-ft 2

-43.5 slug-ft 2

-291.8 slug-ft 2

43.4 inches ] from CM

0.5 inches / origin =5.3 inches S/C sta. 1000

129.4 square feet

154.0 inches

see: Table (5.2), Fig. (5. 21)

see: Fig. (5.22)

5-44

Table 5. 2

Aerodynamic Coefficients Against Angle of Attack

for the Command Module with Protuberances

_,deg.

140.465

145.465

150.465

155.465

160.465

165.465

170.465

175.465

C M

0. 03282

0. O2686

0. O1851

0. 00779

-0. 00268

-0. 01411

-0. 02601

-0. 03708

C N

0. 13187 -0

0. 10490 -i

0. 07990 -I

0. 06223 -i

0. 05562 -I

0. 04354 -1

0. 01772 -1

-0. 00144 -i

C A C L

99218 0.52987

10571 0.54042

20796 0.52595

29105 0.47950

36511 0.40405

42967 0.31666

47186 0.22634

50081 0.12010

%

0.84915

0.97033

1 09038

1 20032

1 30513

1 39484

1 45446

1 49600

L/D

0.62400

0.55695

0.48236

0.39947

0.30958

0.22702

0.15562

0.08082

NOTES: i.

2.

Above Table for Math I0.0

Coefficients for Moment Center at

X = 1043.1 inchesC.g.

Y = O. 0 inchesc.g.

Z = 5.4 inchese.g.

5-45

000

_b,,I

"i,

0

.r'4

©

E

q_

I

.r--_

5 -46

200

150

Qtrim

100

5O

0 I I2 4

MACH

I I6 8

NUMBER

I10

0.4

0.3

L/D trim0.2-

0.1-

0 I I I I2 4 6 8

MACH NUMBER

I10

o.sI-

0.6 L

C L trim / _

0.4 /0.2

0

2.0

1.5

CD trim1.0

0.5

I I I I t o2 4 6 8 10

MACH NUMBER

NOTE" COEFFICIENTS FOR MOMENT

Xc.g. = I043.1 ins

Yc.g. :0.0 ins

Zc.g. : 5.4 ins

I I I I 12 4 6 8 !0

MACH NUMBER

CENTER AT

Fig. 5-22 Experimental Trim Values for Block I CM with ExternalProtruberances

5 -47

5.5 Physical constants

5. 5. 1 Geophysical constants

Earth's gravitation constant

Gravity potential harmonic coeff.

Earth's mean equatorial radius

Earth's sidereal rate

Reference ellipsoid

Symbol

MUE

J

H

D

RE

WIE

Value

3. 986 032 233 x 1014

meters3/sec 2

1.62345x 10 -3

-5-0. 575 x i0

0.7875 x 10 -5

6. 378 165 x 106 meters

-57. 292 106 35x i0

radians/sec

Fischer, 1960

5 -48

5.5.2 Conversion Factors

International feet to meters

Pounds to newtons

Slugs to kilograms

Nautical miles to kilometers

Statute miles to kilometers

Slugs to pounds (g)

Multiply by

0.304 8

4. 448 221 530

14. 593 902 680

1. 852

1. 609 344 000

32. 174 048 000 ft/s/s

5 -49

6. G&N ERROR ANA LYSIS

This section provides the results of G&N Error Analysis. Table G-1 summarizes

the one-sigma total error at each major event time and breaks these down into the con-

tributions of IMU errors accumulated during each powered phase. Tables 6-2 through

6-16 break down each line of Table 6-1 into the contributions of each IMU sensor error

term.

On the basis of these data the following key errors are estimated:

Entry 3' i (one sigma)

Entry V i (one sigma)

CEP at Pacific Recovery Point:

, 0. 165 degree

18.0 feet per second

15.6 nautical miles

The following comments expla_," the terminology, method of analysis and the basic

assumptions used.Xsm _ SM_ STABLE

l) The IMU Stable Member axes are aligned prior J MEMBER/

to launch relative to local vertical axes as in- |]

dicated in sketch. XSM is up along local verti- [

cal at instant of launch, while ZSM is along _/_/_ll _ _ Zsmlocal horizontal pointed down-range at an _l',._'kocol _-.,.'_--Earth

Verticalazimuth of 105 degrees. V__i

1

2) The data in the error tables are given relative to local vertical axes (altitude,

track, range) at the particular event designated.

ol _nly ,he stgnlllcan_ error figures have been listed in the error tables.

4) No realignment of the Stable Member was assumed.

5) Accexerometer bias errors affect indication errors in two ways. First, they

affect the initial pre-launch alignment of the Stable Member. Second, they

affect the in-flight computation of position and velocity. The two effects are

summed in the tables, since the accelerometer bias error prior to launch is

assumed to be correlated with the bias error during flight.

6) Accelerometer inputs to the AGC are not used during the free-fall phases of

the trajectory.

7) "Initial S.M. Alignment Errors" includes only the uncorrelated alignment

errors. They do not include the alignment errors due to accelerometer bias

errors. The azimuth alignment error (about XSM) is affected principally by

by Z gyro drift effect on the gyro-compassing loop. Since there are other con-

tributing factors to azimuth misalignment, this alignment error has been

°=s .... ,_+,,_._ °+_+_=+_,._11y independent of Z ..... A_¢¢+

6-1

8) The position and velocity errors given in the tables for the various IIV[U

sensor error terms are indication errors. No steering error was assumed.

The indication errors in position and velocity were computed separately for

each sensor error term using an array of error equations and the input position

and acceleration trajectory data. These equations take into account the effect

of the platform error on the gravity vector computation. For each trajectory

run the position and velocity errors due to each platform error are computed

simultaneously and printed in a summary table for all trajectory events of

interest.

6-2

u nun N "

All .... . w

TABLE 6-1

Time

fromEvent

start

(mlns)

SIVB

Cutoff 10.3

SPS 1 st

Burn 14.7

Cutoff

Coast

Apogee 42.2

Coast i_na

End (SPS2nd Burn 65.4

Ignition)

SPS

2nd Burn

Cutoff67.0

EntryStart

72.0

EntryEnd

(at altitude

of 50, 000ft)

202 TRAJECTORY ERRORS

Position Error (n. miles)Type of Error

A It. Track Range

Velocity Error (ft/sec)

Alt. Tra ck Range

i) Total Indication Error 0.34 3.07 0.17 9.2 72.1 4.2

I) Total Indication Error 0.75 6. 35 0.53 13. 5 79. 5 7.72) Effect of IMU Errors

during SPS ist Burn 0.06 0.34 0.05 3.3 17.7 2. 3

i) Total Indication Error 2.68 ii. 14 5.84 33.8 46.2 12.22) Effect of IMU Errors

during SPS Ist Burn 0.74 2.72 i. 16 7.3 3.6 3.1

i) Total Indication Error 3.84 5.35 ii. 36 70. 3 73.5 16. 12) Effect of IMU Errors

during SPS ist Burn i. 46 0.08 2.99 20.0 17.1 6.2

i) Total Indication Error

2) Effect of IMU Errors

during SPS ist & 2ndBurns

3) Effect of !MU errors

during SPS 2nd Burn

1) Total Indication Error

2) Effect of IMU Errors

during SPS 1st & 2ndBurns

3) Effect of IMU Errors

during SPS 2nd Burn

3.78 6.51 11.80 73.2 74.1 17.7

I. 46 0.39 3.20

0.02 0.05 0.01

3.17 9.70 13.20

i. 24 I. 43 3.93

0.20 0.34 0.04

4.76 12.06 14.43

20.7 22.1 7. 7

3.1 6.1 1.6

82. 5 53. 5 18.0

24. 3 19.8 7. 5

3.9 5.6 0.6

116.0 38.7 28.61) Total Indication Error

2) Effect of IMU Errors

during SPS 1st & 2nd

Burns & Entry 2.253) Effect of IMU Error

during SPS 2nd Burn

& Entry 1.074) Effect of IMU Errors

during Entry only 1.70

3.60 4.92 56.1 50.3 ii. 9

1.64 0.62 36.8 49.0 9.7

1.66 0.31 45. 3 48.9 7.80

6-3

g, i,

m_

Ore

UO

Error

Position

A

T _-----_- R

/

(E)Xio

(F)YIo

(F)ZIo

RMS

Frror

0 ft

0 ft

0 ft

I

0

Velocity

Initial

S.M.

Alignment

l'rrors

Ace, 1. IA

Nonorthog-

onality

Bias

Error

Scale

Factor

Error

A eccl. Sq.Sensitive

Indication

]:] rror

Bias

Drift

Acceler-

ation

Sensitiw"

Drift

Acceler-

ation

SquaredSensitive

Drift

(I___) VXI o -4

(l')VyI o

0 it/see

0 it/see

{I')Vz[ 0 it/sec

A(SM)X I 3_ 6 mr

A(SM)Y I 0.04 mr

A(SM)ZI 0.07 mr

XtoY 0.1 mr

X to Z 0. 1 mr

0.1 mrYtoZ

Direct effect

ACBX Fff on Init Mira

Combined Fff

Direct effect

ACBY Fff on Init Mira

Combined Fff

Direct effect

ACBZ Fff on Init Mlm

Combined Fff

SFEX

SFEY

SFF, Z

0.2 cm/sec 2

0.2 cm/sec 2

0.2 cm/sec 2

87 PPM

87 PPM

87 PPM

i0 ,g/g2NCXX

NCYY 10_g/g2

NCZZ I0 ug/g2

BDX Direct effect 3.6 meru

BDY Direc( effect 3.6 rnPru

BDZ Direct effect 3.6 meru

A DIAX

A DSRAY

A DIA Z

15 meru/g

I0.5 meru /g

_)X

A2 D(SRA)(SRA)Y

2

,A A(tA)(IA) z

15 meru /g

1 meru/g 2

1 meru/g 2

1 meru/g 2

Root Sum Square Error (in ft and ft/se_')

Root Sum Square Error (in n. mi. and it/see)

Final Position Error

in Local Axes

(in feeG

Alt. ]Track Rangel

; II

Final Velocity Error

in Local Axes

(in it/see)

Alt. Track Range

0 0 0 0 0 0

-18,580 -71.61

153 - 283 0.75

I 403 0.89

-0.73

557 157 2.34 -0.61

-I,316 369 -4.65 1.18

,- i.194[

1,1731

359 ___ i-__, 142_

779 1,441

-1,138 29'_

573 15S

_____; R__

134 !

l95

-1.24

-'_ 83

07

__ -I. 62

0 lI

_30 1_ -C. 5_

28 -0.24

- 3.69

2.60

- 1.09

-3.52

3.73

0.21

0.38

0

-1.65

0.05

C 0

2i 68 -0.09 -0.27

333 - i.97

281 302 1.86 -1.29

213 0.74

- 1,594 F7.71692 642 -4.95 3. 19

608 2.47

....... ...........

94 ____-- - 88 __ 0.85__ ........... -0.42

58 O. 23

l2 t 069 18,667 1 t 028 9.20 72. 12 4. 16

0.34 3.07 0.17 9.2 72.1 I 4.2

Table 6-2 Total Indication Errors at SIVB Cutoff

6-4

_ _r-,c-,sdkjl_11k_ i_ iba ,s-._ _A a

i .

IL-,%,,_'. _ • . i,.F I.. I .i j w# _L

8_.

¢

0

P]z'ror

P_,sitior,

V,'locit V

Initial

S. lVl.

A!_gnment|"r'rors

A(','cl. IA

No rio rt ho R -

,reality

Bias

Error

Scale

Factor

I':r rot

Acrel. Sq.

Sensitive

IndJ cation

F rror

Bias

Drift

Acceler-

ation

Sensitive

Dr] ft

A ccelel'.-

ation

Squared

Sensitive

Drift

A

jf

( E)X Io

(F)YIo

(F)ZIo

V)VxI o

(l")Vyl o

(F')VzI o

A(SM)XI

A(SM)YI

A(SM)ZI

XtoY

XtoZ

YtoZ

Direct effect

ACBX Eff on Init Mlm

Combined Fff

Direct effect

ACBY Fff on Init Mlm

Combined Vff

Direct effect

ACBZ Elf on Init Mlm

Combined Fff

SFEX

SFEY

SFEZ

NCXX

NCYY

NCZZ

BDX Direct effect

BDY Direct effect

BDZ Direct effect

A DIAX

ADSRAY

A DIA Z

A 2 D_IA)(IA)X2

A D(SRA)(SRA) Y

2

A A(EA )([A) z

RMS

Error

O ft

0 ft

0 ft

0 ft/sec

0 ft/sec

0 ft/see

3.6 mr

0.04 mr

O, 07 mr

0.1 mr

0.1 mr

0; I mr

0.2 cm/sec 2

0.2 cm/sec 2

0.2 cm/sec 2

87 PPM

87 PPM

87 PPM

10 ug/g 2

10 ug/g 2

10 ug/g2

3. R meru

3.6 meru

3.6 meru

15 meru/g

10.5 meru /g

15 meru /g

I meru/g 2

1 meru/g 2

1 meru/g 2

Rcot Sum Square Error (in ft and ft/spr)

Root Sum Square Error (in n. mi, and ft/sec)

Final Position Error

in Local Axes

(in feet)

Alt. Track Range

O 0 0

Final Velocity Error

in Local Axes

(in ft/sec)

Alt. Track ]Range

r

o o o

-38,364 -78.85

240 558 0.83 -0.92

622 0.75

1,160 671 3.12 -!.47

-2,598 1,491 -6.97 3.2.t

- 2,307

1.812

495

-1,412

-1,224

-2_636

962

0

,- 4.72

2 20

- 2,52

-1,893 -3.61 -3.19

2,844 -4.22 4.67

951 -7.83 0.88

539 -2.07 0.82

0

h_;:4 L 762 -!.44 -I,54

152 84 -0.31 0. I]

0 0

89 120 -0.22 -0.22

938 -2.73

685 894 2.55 -2.26

41 0.77

- 3,81_ -9.27

-1,842 2,174 -6.92 5.96

I 1,294 2.71

, 363 -0.82245 290 0.91 -0,78

122 0.25

4,506 38,600 3,224 13.50 79.54 7.66

0.75 6.35 0.53 13.5 79.5 7.7

Table 6-3 Total Indication Errors at SPS 1st Burn Cutoff

6-5

M

O

O

<

Error

Position

Velocity

Initial

S.M.

AlignmentFFF_FS

Aecel. IA

Nonorihog -

onality

Bias

Error

Scale

Factor

Frror

A ccel. Sq.Sensitive

Indication

Error

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-

ation

SquaredSensitive

Drift

A

L_,/

(E)Xio

(F)YIo

(F)ZIo

(F)VxI o

(E)VyI o

(F)VzI o

A(SM)XI

A(SM)YI

A(SM)ZI

XtoY

iX to Z

YtoZ

)irec_ effect

ACBX Eff on Init Mlm

Combined Eff

_Direct effect

ACBY Fff on Init Mlm

Combined Fff

Direct effect

ACBZ Fff on Init Mira

Combined Fff

SFEX

SFEY

SFEZ

NCXX

NCYY

NCZZ

BDX

BDY

BDZ

RMS

Frror

0 ft

Oft

0 ft

0 ft/sec

0 ft/sec

0 ft/sec

3.6mr

O. 04 mr

O. 07 mr

0.1 mr

O.I mr

0.1 mr

0.2 cm/sec 2

0.2 cm/see 2

O, 2 cm/sec 2

87 PPM

87 PPM

87 PPM

10 _g/g2

I0 _g/g2

10 _g/g2

Direct effect 3.6 meru

Direct effect 3.6 meru

Direct effect 3.6 meru

ADIAX 15 meru/g

ADSRAY 10.5 meru /g

ADIAZ 15 meru /g

A 2 D(IA )(IA)X

A2 D(SRA)(SRA)Y

2

A A (IA)(IA) z

Root Sum Square Error (in ft and ft/ser)

Root Sum Square Error (in n. ml. and ft/sec)

Final Position Error

in Local Axes

(in feet)

Final Velocity Error

in Local Axes

(in ft/sec)

Alt. Track Range Alt. Track Range

0 0 0 0 0 0

-2,041 -17, 46

17 - 15 -0. 15 -0. 13

0.05

49 - 31 0,42 0.27

-180 115 -1.44 0.90

-116

89

-205

- 7

- 27

0

-210

IR

-191

-176 -9.93

79 -0.78

- 97 71

4 -6.06

0

- 41 -0.24

0 0

2 3 -0.02

107

82 - 73 0.77

17

290

1-237 210 -2.16

71

1 meru/g 2 23

1 meru/g 2 31 - 28 0.28

1 meru/g 2 7

376 2,074 276 3.25

O. 06 O. 34 O. 05 3.3

I. 64

0.15

1.49

-1.37

0.67

-0.70

0.03

0

-0.35

6

0

-0.02

-0.98

-0.66

0.14

-2.51

1.84

0.58

-0.20

-0.24

O. 05

17.74 2.32

17.7 2.3

Table 6-4 Effect of IMU Errors during SPS i st Burnat SPS i st Burn Cutoff

6-6

--_=_f%kl CIIt_CIklTI A L ,

e_

©cg

oL)<

2

!_]rror

Position

Velocity

Imtial

S.M.

AlignmentFrrors

Aecel. IA

Nonortho g -

onality

Bias

Error

Scale

Factor

Error

Accel. Sq,Sensitive

Indication

Error

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-

ation

SquaredSensitive

Drift

A

---..- R

TJff

E)XIo

(F)YIo

(F)ZIo

(I")VxI o

(F)Vyi o

(F)VzI o

RMS

A(SM)ZI

Error

0 ft

O ft

0 ft

0 ft/sec

O ft/sec

O ft/sec

A(SM)XI 3.6 mr

A(SM)YI 0.04 mr

0.07 mr

X to Y 0.1 mr

X to Z 0.1 mr

0.1mrYtoZ

Direct effect

ACBX Eff on Init Mlm

Combined Eff

Direct effect

ACBY Fff on Init Mira

Combinpd Fff

]i)irect effect

ACBZ_f_ o,1 Init Mlm

I Comb;n_- Fff

SFEX

SFEY

SFFZ

0.2 em/sec 2

0.2 cm/sec 2

0.2 cm/sec 2

87 PPM

87 PPM

87 PPM

in _, g/gZ

I0 /j g/g2

NCXX

NCYY

NCZZ 10 ug/g 2

BDX Direct effect 3.6 meru

B DY Direct effect 3.6 meru

BDZ Direct effect 3.6 meru

ADIAX 15 meru/g

ADSRAY 10.5 meru /g

ADIAZ

Final Position Wrror

in Local Axes

(in feet)

Aft. Track Range

Final Velocity Error

in Local Axes

(in ft/sec)

Alt. [Track [Range

0 0 0 0 0 0

32,300 72.86

- 2,715 3,804 - 4.62 2.01

573 - 0.67

58 -10,813 9.70

57 24,562 -22.12

1,945

- 1_669

276

4.36

- i.95

2.41

-25,226 71,29_ -75.49

_843 ..... --I 9, 39! _ 23_ 57 ....

-ll,3a3 51,3_Y -51.92 I

-' .........- _._a_.,............ 2_!"_3::'_c!2:VK....I 0

.... I i26a 1,aa._ - ,.,,_ :

- 1,508 4,29( - 4.54

714 2.57

- 5,996 4,75{ - 7.23

- 354 - 0.71

3,080 8.64

15,515 -I0, 67( 17.30

15meru /g - 1,085 - 2.51

1 meru/g 2 300 9.76iA2DIIA)(IA)X

A 2DISRA)(SRA )Y

A 2A(IA)(IA)Z

1 meru/g 2 - 2,010 1,361 - 2.23

] I. 26!i

- 2.70

l14.70

10.27

4.43

0.07

1 meru/g 2

Root Sum Square Error (in ft and ft/seri

Root Sum Square Error (in n. mi. and ft/sec)

_ _ ¥_r

I

[ 5.93

0.05

I O. 88

1.62

- 102 - O. 23

23,320 32,485 69,050 70.32 73.51 16.06

3.84 5.35 11.36 70.3 73.5 ] 16.1

I

I

Table 6-5 Total Indication Errors at Coast End

(SPS 2nd Burn Ignition)

6-7

M 0

M

<

0

Error

Position

Velocity

Initial

S.M.

Alignment_rrors

Accel. IA

Nonorthog-

onality

Bias

Error

Scale

Factor

Error

Accel. Sq.Sensitive

Indication

Error

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-

ation

SquaredSensitive

Drift

A

_---.--.----._ BTJ

ff

(E)XIo

i(E)Yio

(F)ZIo

(F)VxI o

i(E)Vyi o

(F)VzI o

RMS

l_rror

A(SM)ZI

oft

Oft

0 ft

0 ft/sec

0 ft/sec

0 ft/sec

A(SM)XI 3.6 mr

A(SM)YI 0.04 mr

0.07 mr

XtoY 0. i mr

Xto Z 0.1 mr

0.1 mrYtoZ

Direct effect

ACBX Eff on Init Mlm

Combined Fff

:irect effect

ACBY Fff on Init Mlm

Combined Fff

Direct effect

ACBZ Fff on Init Mlm

Combined Fff

SFEX

SFEY

SFEZ

NCXX

NCYY

NCZZ

BDX Direct effect

BDY Direct -effect

BDZ Direct effect

ADIAX

0.2 cm/sec 2

0.2 cm/sec 2

0.2 cm/sec 2

87 PPM

87 PPM

87 PPM

10 ,g/g2

I0 _g/g2

10 .g/g2

3.6 meru

3.6 meru

3.6 meru

15 meru/g

ADSRAY 10.5 meru /g

ADIAZ 15 meru /g

A 2 D(IA)(IA)X

A2D(sRA)(SRA)Y

2

A A (IA)(IA) z

1 meru/g 2

1 meru/g 2

1 meru/g 2

Root Sum Square Error (in ft and ft/see)

Root Sum Square Error (in n. mi. and ft/sec)

m 6_4 Awv ¥¢#"

Final Position Error Final Velocity Error

in Local Axes in Local Axes

(in feet) (in ft/sec)

Alt. Track Range Alt. Track Range

0 0 0 0 0 0

432 377 - 0.58 0.33

2 - 0.05

478 16.83

763 132 - 0.58 0.63

2,515 278 1.78 -2.09

64 I. 58

% - 0.14

58 i. 44

-7,204 18,366 -19.86 4.43

2,204 - 1,922 2.98 -1.68

-5,000 16,444 -16. 88 2.75

97 - 1i 0.07 -0,08

0 0

-1,841 4,679 - 5.06 1.13

1 C 0 0

- 129

0 0

328 - 0.36 0.08

20 0.94

-2,191 1,932 - 2.98 1.67

4 - 0.14

66 2.42

6, 142 - 5,39_ 8.34 -4.67

- 19 - 0.56

5 0.19

789 690 - 1.07 0.60

2 - 0.05

8,893 492 18, 180 19.96 17. i0 6.24

I. 46 O. 08 2.99 20.0 17. 1 6.2

Table 6-6 Effect of IMU Errors during SPS I st Burn at Coast End

(SPS 2nd Burn Ignition)

6-8

-"$: ".-I........ L '

*°ir

"D L.

CC

©

C_

f

!i

i

2

t<rror

°csition

Velocity

Initial

S. IVl.

Aligmnent

l"rrors

A cc,'l. IA

Nonorthog -

onality

Bias

Error

Scale

Factor

Error

A ccel. Sq.

Sensitive

Indication

Error

_! ias

Dr'ft

Acc _qer-

ation

Sensitive

Dri?tAcceier-

atlonSquared

SensitiveDrift

A

( K)Xio

J(F)YIo

(F)ZIo

(F)VxI o

(F)Vyi o

(F)Vz[ o

A(SM)XI

A(SM)YI

A(SM)ZI

R his

Frror

0 ft

O ft

0 ft

0 ft/sec

8 ft/sec

0 f_/sec

3.6 mr

0.04 mr

0.07 mr

X to Y 0. 1 mr

X to Z 0.1 mr

Y to Z 0. I mr

',

Direct effect !

iACBX Fff on [nit Mlm

Combined Nff

Direct effect

ACBY __iVff on Init Mlm

Combined Fff

IDirect effect

ACBZ Fff on Init Mlm

Combined Fff

SFEX

SFEY i

+SFEZ I

NCXX [

NCYY

INCZZ

BDX Direct effect

BDY Direct effect

BDZ Direct effect

0.2 cm/sec 2

0.2 cm/sec 2

0.2 era/sec 2

87 PPM

87 PPM

87 PFM

1O_g/g 2

1O ug/g 2

iO ug/g 2 ]3.6 meru

3.6 meru

3.6 meru

ADIAX 15 meru/g

ADSRAY 10.5 mere2 /gI

ADIAZ

!A2D ........

(SRA){SflA)Y

I 2]A A (IA)(IA) Z

15 meru /g

1 meru/g 2

! meru/g2

I meru/g 2

Root Sum Square Error (in ft and ftlser)

Root Sum Square Error (in n. ml. and ft/sec)

Final Position Error

in Local Axes

Alt.

- 2,720

- 323

574

-24,424

13,866

-10,558

- I, 268

- 9 fl55

260

(in feet)

Track

i-39,336

- 626

2 322

t- 1,824

4981

i, 022

- i. 460

- 6)019

304

3,926

15,864] -i3, 70_

1

]- i, 222

373

.... j

22,982 39,576 71, 72_

3.78 i 6.51 11.30

Final Velocity Error

in Local Axes

(in ft/sec)

Range Alt., ] Track I Range

' ii 1

o o o I o

II

!73.40

4,30_ - 4.96 2. i2

1- 0.43!

-10,662 9.92 1.40

iI24,23[ -21.87 I ° 3.52

1-----

3.47

- 1.241

2.23 i

75, 515 -79.58 I i5.03

-21,97] 25.28 I -i0.79

I53, 54( -54.30 4.24

I0,506 -I0.34 - O. 15

O 0

3n ._6_ -32. 23 5. 75

0 i,725 - I. 74 0.030

4,541 - 4.80 0.87

J .......

3.88

5,964 - 5.58 6.31

1.87

8.98

18.29 -13. 86

- 0.271-

O. 78

- % --.

- 2.35 1.80

- 0.03

73. 17 74. II 17.70

73.2 74. I 17.7

Table 6-7 Total ..................... at ...... Burn _ULOliII l_li C _:t. t lO i I _ , pc.r',L'L'OL'_ _D L_ ,D /.II(l

6-9

C n" ........ TIC,'

x o

>.O

Error

Position

Velocity

Initial

S.M.

Alignment|-'trots

Accel. IA

Nonorthog -

onality

Bias

Error

Scale

Factor

Error

A ceel. Sq.Sensitive

Indication

Error

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-

ation

SquaredSensitive

Drift

[loot Sum Square

Root Sum Square

A

t

jf

(E)XIo

(F)YIo

(F)ZIo(V)VxI o

RMS

Frror

0 ft

0 ft

0 ft

0 ft/sec

0 ft/sec

0 ft/sec

A(SM)X I 3.6 mr

A(SM)Y [ v. 04 mr

O. 07 mrA(SM)ZI

X to Y 0. I mr

XtoZ 0.1mr

Y to Z 0. I mr

Direct effect

ACBX Fff on Init Mira

Combined Fff

Direct effect

ACBY Fff on Init M1m

iCombined Fff

)irect effect

ACBZ Eff on Init Mlm

Combined Fff

SFEX

SFEY

SFEZ

NCXX

NCYY

NCZZ

BDX Direct eff, ct

BDY Direct effect

BDZ Direct effect

A DIA X

ADIAZ

2O. 2 cm/sec

Final Position Error

in Local Axes

(in feet)

Alt. t Track

0 0

-2,318

440

I

Final Velocity Error

in Local Axes

(in ft/sec)

Range A1t. Track Range

0 0 0 0

21.56

46] - 0.55 0.40

0.12

796 279 - 0.59 0.60

2,670 794 2.69 -2.59

185 0.93

0.2 cm/sec 2 ..3 O. 35

188 128

-__7__0.22 19, aI_ -2n 76 4 _5

0.2 cm/sec 2 2,245 I - 2,35_ 2.82 -2.03

-4,777 ] 17,26_ -17.94 2.92

87 PPM 110 [ 3_ 0.26 -0.20

87 PPM 0 0

87 PPM -1,802 4,988 - 5.45 1.03

10 pg/g2 2 - 1 0_ 02 -0.01

ug/g2_ 0 010

ug/g 2 - 126 350 - 0.37 0.0810

3.6 meru 175 I 2.35

3.6 meru -2,136 2,407 - 0.75 3.09

3.6 meru 97 2.39

15 meru/g 332 3.17

ADSRAY 10.5 meru (g__ 6,_2_49 ........ - .6f59__ __7. 81 .... -5.71

15 m_,ru /g 24 1.55

2 2 .....

A D_ ......... 1__mer_u/g _ -t 26 O. 25

SRA, ....lZo u/ : i........ 84:-100 0732 1 meru/g 2 2 0. 14

A A (IA)(IA) z

IJ 8_,870 2,3_9 19,472 ...........20.66 22.14 7.73Error (in ft and ft/_rl #-Error (in n. mi. anti ft/sec) _1.46 0.39 ; 3.20 20.7 22. 1 7.7

Table 6-8 Effect of IMU Errors during SPS 1st and 2nd Burnat SPS 2nd Burn Cutoff

6-10

Effect of IMU Errors during SP5 2nd }lurn at SI'5 2nd }{urn Cutoff

[d

r.1

OO

<

Error

Position

Velocity

Initial

S.M.

Alignment

l.'rrors

Aecel. IA

Nonorthog -

onality

Bias

Error

St

ff

i(E)Xlo

r(F)Ylo

(F)ZIo

(F)VxI o

(F_)VyI o

(F)VzI o

A(SM)XI

A(SM)YI

A(SM)ZI

XtoY

XtoZ

YtoZ

Direct effect

ACBXIEff on Init Mlm

Combined Fff

i.Direct effect

ACBY Fff on Init Mlm

Combined Fff

Direct effect

ACBZ Eff on Init Mlm

Combined Fff

Scale

Factor

Error

A ccel. Sq.

Sensitive

Indication

Error

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-

ation

Squared

Sensitive

Drift

RMS

Frror

Final Position Error

in Local Axes

0 ft

O ft

0 ft

0 ft/sec

0 ft/sec

O ft/sec

3.6mr

0.04 mr

0.07 mr

0.1 mr

0.1 mr

0.1 mr

O. 2 cm/see 2

5 F" K X

SFEY

SFEZ

NCXX

NCYY

NCZZ

BDX Direct effect

BDY Direct effect

BDZ Direct effect

A DIA X

A DSRAY

A DIA Z

2 2

2

A A(EA)(LA) Z I m_'ru/_ 2

Root Sum Square Error ,in ......ft .... ,ft/ ..... I liO- _- 73

"ootSum quareError,inn.mt..... 10.05100,

Final Velocity El for

in Local Axes

(in feet) (in ft/sec)

Alt. Track [ Range Alt. Track Range

I

0 0 0 0 0 0

227 4.88

5 2 0.10 0.05

8 0.17

5 - 3 0.11 -0.07

25 - 17 0.53 -0.35

30 -O 63

0.2 cm/sec 2 23 0.49

7 -0.14

17 25 0.35 0.53

0.2 cm/sec 2 23 12 -0.51 -0.26

6 13 -0.16 0.27

87 PPM 8 5 0. 18 -0. 12

87 PPM0 0

87 PPM 3 5 -0.07 -0. 10

9

10pg/g_ 1 0 0.02 -0.01

10 ug/g 2 0 0

I0 pg/g2 0 0 0 0

3.6 meru 65 1.42

3.6 meru 118 60 2,58 1.33

3.6 meru 115 2.52

15 meru/g 35 0.77

10.5 meru /g 70 35 -1.51 -0.77

15 meru /g 97 2.10

1 meru/g 2 3 .............. 0.06 .....

4 0.19 0.10

0.19

3.06 6.11 1.61

3.1 6.1 1.6

Table 6-9 Effect of IMU Errors during SPS 2nd Burn at SPS2rid Burn Cutoff

6-11

X O

ee

O

Id

_9_9<

Error

Position

Velocity

InitialS.M.

AlignmentFFFOFS

A'ccel. IA

Nonorthog -

onality

Bias

Error

Scale

Factor

Error

Accel, Sq,

Sensitive

Indication

Error

Bias

Drift

A cceler-

ation

Sensitive

Drift

Acceler-

ation

SquaredSensitive

Drift

A

f

(E)XIo

I(F)YIo

(F)VxI o

(F)VyI o

(F)VzI o

A(SM)XI

A(SM)YI

IA(sM)ZI

XtoY

YtoZ

Direct effect

ACBX Eff on Init Mlm

Combined Eff

Direct effect

ACBY Fff on Init Mira

Combined Fff

Direct effect

ACBZ Eff on Init Mlm

Combined Fff

SFEX

!!

i

Xto Z

0.1 mr

87 PPM

SFEY 87 PPM

SFEZ 87 PPM

NCXX

NCYY 10 ug/g2

NCZZ I0 ug/g 2

BDX Direct effect 3.6 meru

BDY Direct effect 3.6 meru

BDZ Dii'ect effect 3.6 meru

iA DIAX 15 meru/g

ADSRAY 10.5 meru /g

A DIA Z

_X

A 2 D(SRA)(SRA)Y

2

A A (IA)(IA) z

RMS

Error

0 ft

9 ft

0 ft

0 ft/sec

0 ft/sec

0 ft/sec

3.6 mr

0.04 mr

0.07 mr

0.1 mr

0.1mr

0.2 cm/sec 2

O. 2 cm/sec 2

0.2 cm/sec 2

10 _g/g2

15 meru /g

1 meru/g 2

i meru/g 2

I meru/g 2

Root Sum Square Error .(in ft and ft/s_'_'_

Root Sum Square Error (in n. mi. and ft/sec)

Final Position Error Final Velocity Error

in Local Axes in Local Axes

(in feet) (in ft/aec)

Alt. Track Range Alt. Track Range

0 0 0 0 0 0

58,534 52.86

2,400 5,968 - 6.26 2. 17

- 714 - 0.15

- 1,351 - 9,853 9.84 1.80

3,063 22, 286 -21.93 - 4.62

-18, 174

12,230

- 5, 944

348

- 7,444

117

- 1,092

i- 4,917

15,079

3,201 2.31

- 0.43- 2.079

Ij122 1.88

2,102

262

6,33O

- 1,227

88,387 -91.67

-30, 427 31.94

57, 956-59.73

10,7451-11. 12

35,501 -37.25

1,803 - 1.89

5,304 - 5.53

9,94f - 7.84

-24, 114 25.69

14.40

-11.06

3.34

- 0.37

0

5.63

0.01

0

0.84

3.23

5.61

1.87

6.83

-14.57

0.25

...... 2781....- 1,955!

117

19,231 58,944

3.17 9.70

0.58

3,104 - 3.31

0.02

1.89

!

80,193 82.52 [ 53.48 18.03

13.20 82.5 I 53.5 18.0

Table 6-i0 Total Indication Errors at Entry Start

6-12

• a __

E i

F:r 1"oI"

Position

A

Imp.. R

//

(E)XIo

Ii

RMS

! Vrror

0 ft

(F)YIo 0 ft

(F)Zlo 0 ft

m_

r_

[4

O

J

UU<

?

Velocity

Initial

S.M.

Alignmentl' rrors

Accel. IA

Nonbrthog-

onality

Bias

Error

Scale

Factor

Hrror

A ccel. Sq.Sensitive

Indication

Error

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-ation

Squared

Sensitive

Drift

(V)VxI o ') f,/sec

(F)Vy[ ° 0 ft/sec

V)VzI o

A(SM)XI

A(SM)YI

A(SM)ZI

XtoY

0 ft/sec

3.6 mr

O. 04 mr

0.07 mr

0. I mr

X to Z 0.1 mr

0.1 mrYtoZ

Direct effect

ACRX Fffon Init Mira

Combined Fff

Direct effect

ACBY Vff on Init Mlm

Combined Fff

Direct effect

ACBZ_on Init Mlm

C_m-ffi.--__KSFEX

SFEY

SFFZ

I .......t_lL, AA

NCYY

,3.2 cm/sec 2

0.2 cm/sec 2

0.2 cm/sec 2

67 PPM

87 PPM

87 PPM

,n _1_2_u_g/g

10 ug/g 2

10 /._g/g_NCZZ

BDX Direct effect 3.6 meru

BDY Direct effect 3.6 meru

BDZ I Direct effect 3.6 meru

i

_DIAX i5 meru/g

_DSRAY 10.5 meru /g

DIAZ

_A)X

2

D(SRA)(SRA) Y

2

A(IA)(IA) Z

15 meru /g

1 meru/g 2

1 meru/g 2

i meru/g 2

Root Sum Square Error (in ft and ft/soc}

Root Sum Squ "e Error (in n. mi. and ft/sec)

Table 6- i i

Final P()sition Frror

ie l,ocal Axi:s

(in feet)

Air. [Track Range

l l

t !0 t7 ! 0

8,514

4OO

37 I

827

Final Velocity Frror

in Local Axes

(in ft/see}

Alt. Track _('

I

i

II

0 C e

it

19.27

746 _ 0._5 0,40

0.11

793 - 1.01 0.69

2,971 I- 2,778 3.94

448 O. 80

...... I07 0.33

555 I. 13

-5,334 23,528 -24.05

2,037 - 3,816 3.82

-3,297 19_712 -20.23

l154 - 16C 0.28

0 _ __.... ----.-r T[ ....

98 42{

855 [ __ 2. 13

-1,179 3,867 1.12

793 2.20

1,243 2.83

5,633 -10, 67{ 10.56

477 !. 44

97 0.22

..................... J ____

725 1,35{ - 1.36

42 ........ 0.. 13- "

I-

7,513 8,715 23,88( 24.25 19.81

1.24 1.43 3.93 24.3 19.8

-3. 15

4.73

-2.05

2.68

-0.27

1.07

-0.02 i

0.08

2.14

-5.70

0.73

7.53

7.5

Effect of IMU Errors during SPS 1st and 2nd Burns

at Entry Start

6 -13

o

._

i

w •[. I

0 '

_J

0L)<

o0

Error

Position

V,.!oeity

Initial

S.M.

Aligrm_ent

l"rrors

A eccl. IA

Norlorthog "

onality

A

f

(E)XIo

(F)YIo

(I')Zio

(I')Vxi o

I(V)VyI o

.C )vzi o

__5sM_J_x!...................

A(SM)YI

A(SM)ZI

XtoY

iX to Z

L

Y to Z

I JDirect effect

ACBX Eff on Init Mlm

Combined Fff

Direct effectBias

ACBY T'ff on Init MlmError

Combined Vff

)irect effe(t

',CBZ l"ff on Init Mira

Combined I f[

Scale _,FEX

RMS

Error

0 ft

0 ft

0 ft

0 ft/sec

0 ft/sec

0 ft/sec

3.6mr

0. 04 mr

0.07 mr

O. lmr

0.1mr

0. I mr

O. 2 cm/see 2

0, 2 era/see 2

F'.nal Position Frror

in Local Axes

(in fertl (in

All. I Track Range

1ItI

0 ] G 0

Ii

ii

11,648 [

40 3

57

I i29 i- 39

),34 ] .... ]-180

- 214

196

- '_8189 118 0.59

0.2 cm/see 2 205 - 13 -0.66

16 105

87 Fi'/q 44 - 80

Final Velocity Error

in Local Axes

ft/sec)

Aft. ]Track iRange

II

i o o

!I

II

i4.49

0.13 I 0.010.16

o.o_ I I-0.11

1 l--0.44 ]

i

q 0 _ 5O

I_"-_- d----

, 0,31

I I-0.03

' i-0.07 i 0.28

0. 14 'I -0. 17

EaCIOF

Error

A ccel. Sq.

Srnsitive

Intlicatioll

Error

_ I"EY

S["V Z

NCXX

NCYY

Direct effect

NCZZ

:_DX :3.6 meru

2 1 0.01

Bias

Drift

Acceler-

ation

Sensihve

Drift

Acceler-

atlon

Squared

Sensitive

Drift

BDY Direct effect 3.6 meru

BDZ Direct effect 3.6 meru

%DIAX 15 meru/g

A DSRAY I0.5

ADIAZ 15

A2I)_ ............ 1

A2 DISRA)(SRA)Y 1

2 1A A(IA)IlA) Z

Root Sum Square Error (in ft and ft/serl

Root Sum Square Error (in n. mi. and ft/sec)

_z4r_ 4_,v v_r

Table 6-12 Effect of IMU Errors during SPS 2nd Burn

at Re-entry Start

(; -14

m

.6_kca [.,.1

r. ¢jo[.-

re

o

Error

Position

Velocity

Initial

S. M.

AlignmentFrrors

Accel.. IA

Nonorthog -

onality

Bias

Error

Scale

Factor

Error

Accel. Sq.Sensitive

Indication

Error

A

T./_--R

f

(E)Xlo

(F)YIo

(F)ZIo

(F)VxI o

(F)VyI o

(F)VzI o

A(SM)XI

A(SM)YI

A(SM)ZI

XtoY

[t M S

t'rror

0 ft

0 ft

0 ft

0 ft/sec

0 ft/sec

0 ft/sec

Final Position Frror

in L(*:a} Axes

(t,_ f*.,.t)

3.6 mr

Alt.

0.04 mr - 3. 317

0.07 mr

0. i mr

X to Z 0.1 mr

0.1 mrYtoZ

Direct effect

ACBX Eff on Init Mlm

Combined Fff

Direct effecti

ACBY[Fff on init Mlm

Combined Fir

Direct effect

ACBZ Eff on Init Mlm

Combined Fff

SEEX

SFEY

SFEZ

0.2 cm/sec 2

903

1,757

1_631

9.2 cm/sec 2 - 2,778

- lm147

ITrack Range

!t 0 0

72,198

8,430

953

- 6.514

NCXX

NCYY

NCZZ

Final Velocity Error

in Local Axes

(in ft lsec)

Aft. [ Trac_ [Range

tii

II

0 0 ] 01

l

II

....

1.86 t

- 11.06 ] - 1.21

- 1.13 l

.39

12,345 - i7.64 -15.0C

- 6.11

- 3.30

- 9.41

.m

...... q

----M

-25.28

6.18

-19.10

-23,933 100_229 -111.25

0.2 cm/sec 2 16,909 -42,983 56.39

- 7_023 57,246 - 54.86

- 1,298 10,112 - 11.47] - 1.82 d

m

87 PPM

87 PPM 0 0

67 PPM -11,211 40,671 - 49.221 - 9.58

10ug,/g _ 220 1,570 - 2.18 ___28 ]___._t

lO ug/g 2 o

1Opg/g 2 - 1,599 6,039 - 7.05 -- _.47

BDX Direct effect 3.6 meru a qnv *n no

Bias BDY Direct effect 3.6 meru -_'" __ ""'_- .... -J o_Drlft -12 a 100 1_4 6,,_ -.,#. V._$ - ,. o,

..... BDZ Direct effect 3.6 meru 8_270 ........ | 2_00 ........-- ---- --............ 4 :_..o ........

Acceler- ADIAX 15 meru/g 8,174 -1-t- 2. B_

1

atlon A R 0 ....................DS AY I .5 meru /g 22,045,]- -41,88t) 85. t9 ] I 2.:6

Sensitive ........... _--........... j_ .... ] -- ]Drift ADIAZ 15 inert _ - 5 670 -1_ _ i................. .... I---"- ...... ....[-- t _!eel 2 / 2 / ]

Ac. er- A 13.......... 1 meru _ I 8_" / ..... Iati,m _IP)JIIPLIA ' t ' r i..r_ ,

Squared A2D(SHA}(S_.IA)y 1 morll/g z _ 2,8D61 . 5,40n 8.8i[ [ 0. Sa ]

S,-nsi,ive TK ......................... :-_ -_-_ -- I --- _ .... ] , .... 1

,Or', S ..... Square Error _ft an(_ ,t]._o._' I 2">'23_2_7,6951 ]!6.0::_73

_oot s.... S.,,s_r e_,'o_ ,'inn._. a.d ft/_._) [ .. v,; I ,_.06 I ,,* 43 II 116 0 t 38.7 I ... I

Table 6-13 Total Indication Error_ nt F:ntry Pnd (at _a non ÷'+ 51+)

(;-15

>1 o

©,w

(9O

oO

EFFoF

I'_,:,itiun

Vvlocity

[nitialS. M.

A 11gnmcntI l'l'OI'S

4' <'_+I. IA

(_<,tlc3t't hog -

_m:Jlity

_las

Error

Scale

Factor

Error

Accel. Sq.

Sensitive

Indication

Error

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-

ation

SquaredSensitive

Drift

-A

................ _. ,- ,

(i.)N ( ° u +t

............. ""(i,)Z[o 0 ft

(}.)Vx[ ° U lt/_ec

(1.,IVy[ ° 0 n ,' sec

+(i)Vz[ ° 0 ft / _,ec

RMS

I rror

A(SM)X I :_. 6 mr

A(SM)Y[ O.U_ mr

AISIM)Z[ I). !17 n,r

X to Y 0, 1 nlr.i

X to Z 19, [ nlF

Y to Z O. I mr

IJirect ct [e<-t

ACB_ I'J[,n lnit Mint 0.2 cm/sec 2

('ond]hlr!(I l'ff

IIirect c fl"_'t:l

ACI3Y Iff on hut Min, 0.2 cn_/scc 2

Cornbined Fff

Direct effect

ACBZ P:ff on Init Mira 0.2 cm/sec 2

Combined Fff

SFEX 87 PPM

SFEY 87 PPM

SFEZ 87 PPM

NCXX 10 ug/g 2

NCYY 10 ug/g 2

NCZZ 10 ug/g 2

BDX Direct effect 3.6 meru

BDY Direct effect 3.6 meru

BDZ Direct effect 3.6 meru

ADIAX 15 meru/g

ADSRAY 10.5 meru /g

ADIAZ 15 meru /g

A2 D(IA)(IA)X i meru/g 2

A2D(sRA)ISRA)Y

2

A A(IA)(IA) Z

Final Position Error Final Velocity Error

in Local Axes in Local Axes

(in feet) (in ft/sec)

Aft. Track Range Alt. Track Range

0 0 0 0 0 0

20,284 31.15

720 1,195 - 2.54 - 0.26

3:]6 - I. 72

- 1,148 1,781 - 2.58 - 0.62

3,777 - 8,881 a.05 - 5.38

- I, 003

,980

,- 1_9,2_8 :_ ......

3, fi 73

- 1,093

262

- 2,300 7,140

12 ,- 154

4

- 122

- 7,370

- -i. 30

- 5.00

3.30

- 6,092 i2.9_ 1.31

21,025 -11.5_ - _i.80

74 0,76 2.03

51 - 0.07

-9.43 ,- 1.37

- 0.08 - 0.67

- 0.02

493 - 0.47 - 0.12

2. 8441 I0.67

4+180 -39.52 - 7.59

- 5,740 -31.32

2,510 - 0.23

10,148 -16,954 35.33 32.58

- 3,870 -19.11

323 1.34

1 meru/g 2 - 1,352 2,169 - 4.94 - 0.57

1 meru/g 2

Root Sum Square Error (in ft and ft/se¢')

Root Sum Square Error (in n. ml. and ft/sec)

391 - 2.08

13,665 21,867 29,873 56.08 50.27 11.91

2.25 3.60 4.92 56.1 50.3 11.9

Table 6-14 Effect of IMU Errors during SPS ist and 2nd Burnsand Entry End (Alt. of 50,000 ft)

6-16

--_I_I plhpL, l"_ai • I

_- _ _ _ _- "--21F_- -" J----

Error

FinM Position l<rror

hi Local A×es

(in f,,et)

Alt, 'Iraek Hange

_ effect

AC]RX _ 0.2 era/see 2

m L IC_mbtned Ff_f _! IDirect _ffect

._ ]ACBY]Fff on Init Mlm 0 2 cm/sec*

,at,o, 1--

INCXX ! l0 /2 _,/gAccel, Sq, _ ...... {

Sensitive tNCY Y ] 10 tag/E 2

Indication _-_ ........ --4 ...... } - -

_ Err°r :? Zz __g 10ug/g

Bias [BDY ] Direct effectA_ - 3.6 meru _

,)rift .4BDZZJ_ Direct effectT-[ .3.6 _m_eeru _

atton ]ADSRAY __. 10. 5 meru /1_,Sensitive _ ..................... _i....

Drift ADIAZ I 15 mPru /g

T_Aceeler- ItIA)X / 1 meru/g

ation ) , 2 _'-"_ ....... _-........... 5-Squared [A_D(sRAIISRAIY n 1 meru/g _

Sensittw' kT_- _z'_'" ='-L_=-"" " _ ................. _.--

Drift [A_A(IA)(__ . ____I.A)Z ......... ___ .... lm. 17_re_a/L_

Root Sum Square I,:rror (in ft and ft/seel

Root Sum Square Error (in n. mi. anti ft/sec) --.

}S'inal VPloclty Frror

in l.ucM Axes

fin ft/see)

0 0

Table 6-15 Effect of IMU Errors during SPS 2nd Burn and Entryat Entry End (50,000 ft alt)

6-17

___L-__,.__,__:_ :"- _" _ __ . "a''"2_- -

o

ML-

0

M,4MUU<

o>,cD

_rror

Position

Velocity

InitialS.M.

AlignmentFrrors

A ccel.. [A

Nonorthog -

onality

Bias

Error

Scale

Factor

Error

Accel. Sq.

Sensitive

IndicationError

Bias

Drift

Acceler-

ation

Sensitive

Drift

Acceler-

ation

SquaredSensitive

Drift

A

I_/

(E)XIo

{E)Yio

{F)ZIo

(F)VxI o

_-R

RMS

Error

0 ft

0 ft

0 ft

0 ft/sec

(F)Vyi ° 0 ft/sec

(F)VzI ° 0 ft/sec

A(SM)X I 3.6 mr

A{SM)Y I 0.04 mr

A(SM)Z I 0.07 mr

XtoY 0.1mr

Xto Z 0.1 mr

Y to Z 0. I mr

Direct effect

ACBX Eff on Init Mlm 0.2 cm/sec 2

Combined Eff

Direct effect

ACBY Fff on Init Mlm 9.2 cm/sec 2

Combined Fff

)irect effect

ACBZ Eff on Init Mlm 0.2 cm/sec 2

Combined Fff

SFEX 87 PPM

SFEY 87 PPM

SFEZ 67 PPM

NCXX 10 _g/g2

NCYY 10 ug/g 2

NCZZ 1O _g/g2

BDX Direct effect 3.6 meru

BDY Direct effect 3.6 meru

BDZ Direct effect 3.6 meru

ADIAX 15 meru/g

ADSRAY 10.5 meru /g

ADIAZ 15 meru /g

:_IA)X 1 meru/g 2

A2D(sR A)(SRA)Y 1 meru/g 2

2 1 meru/g 2A A (IA)(IA) Z

Root Sum Square Error (in ft and ft/ser)

Root Sum Square Error (in n. ml0 and ft/sec)

,lu¢¢ 4av ¥#J-

Final Position Error

in Local Axes

(in feet)

Alt. ITrack Range

Final Velocity Error

in Local Axes

(in ft/see)

Alt. Track [Range _

0 O 0 0 0 0

2,493

284

- 433

11 - 1,25

i0 - 69 - 0. ii

- 212 - 1,795 - 0.63

- 1,797

- 1.261

- 3.058

2.201 31E 6.57

1_446 5[ 6.39

3,647 373 12.96

69 53[ 0.31

51

55 7 - 0.67

18 - 132 - 0.09

- 4

13 2 0. Ii

928

- 8,757 254 -39.66

- 7,662

I0(

4,107 201 17.34

- 5,103

120

572 1{ - 2.63

500

10,359 10,069 1,942 45.30

1.70 1.66 0.31 45.3

28.14

-0,, 16

-1.75 1

._-- -0p78

I -4.31

- 4.33

- 5, 09

- 9, 42

-0.77

0.83

0.06

2.15

- 0.07

0.09

-0.66

- 0.02

-0.01

i0_25

-5.50

-31.84

- 0.68

1.84

-19.49

1.30

-0.39

- 2.11

48.88 7.80

48.9 7.80

Table 6 - 1 6 Effect of IMU Errors during Entry only at Entry End(50,000 ft Aft)

6 -18

7...C"'"" -- ),:" 'A"

7. G&N CONFIGURA TION

System 017 will be the G&N system for Mission 202. It is a Block I series 50

system with one modification; the wiring of the 11 spare relays in the main DSKY to

the MCP to provide the AGC/MCP signal interface (refer ICD #MH01-01200-216)

described in Section 3.

Without giving a detailed analysis of each G&N Block configuration, a brief descrip-

tion of each and the reason for its evolution is useful in understanding G&N's capabilities

for Mission 202.

Block I is the original G&N design. It is composed of IMU, AGC, PSA, CDU's

(mechanical), Harnesses, and OPTICS (sextant and telescope). As the G&N flight

requirements became more clearly defined it was apparaent that Block I would need

modification to qualify for flight.

Block I, series i00 therefore evolved. It is the Block I system modified generally

as follows:

(a) IMU - Vibration dampers added; moisture insulation added.

(b) AGC - Cooling inierface modified; humidity proofing added.

(c) PSA - Cooling interface modified; humidity proofing added.

(d) CDU's - Minor electrical and mechanical changes.

(e) Harnesses - All wiring changed to teflon; connectors humidity proofed.

(f) OPTICS - Addition of automatic star tracker, photometer and minor servo

modifications.

When the full design and production schedule impact of the series i00 modifications

become clear the Block I series 50 configuration was originated, being a limited i00

series modification qualified for flight and available on an early schedule.

Block I series 50 is basically the Block I series i00 system less the automatic star

tracker and the photometer.

7-I

8. INSTRUMENTA TION

8. 1 G&N Instrumentation

The inflight information from G&N is available in three distinct forms: PCM

telemetry of the AGC DIGITAL DOWNLINK, (PCMD); PCM telemetry of low band-

width G&N measurements, (PCM +, PCM, PCME); and on-board recording of high

band-width G&N measurements (TR).

The PCM telemetry of the AGC DIGITAL DOWNLINK has been clearly defined

at the MIT/NAA interface as 50 words of 40 bits each per second. The particular

format of this DOWNLINK is AGC program variable and can remain under MiT's

control without having interface repercussion (see 8. i. i).

The PCM telemetry of the low band-width measurement and the on-board

recording of the high band-width measurements have been defined by NASA in "NASA

Program Apollo Working Paper 1141, Apollo SC Measurement Requirements, Apollo

MissionA-202, Spacecraft 001" dated November ii,1964 (see 8.1.2).

8. i. 1 AGC Digital Downlink

The AGC digital downlink consists of 50 words/sec on the high rate

and i0 words/sec on the low rate. Each "word" contains 40 bits (a 16 bit

register transmitted twice and an 8 bit "word order code"). Since the high

rate will be used exclusively for flight 202 all further discussion will use the

50 words/sec rate.

The digital downlink format is controlled by an AGC program which

loads the next word to be transmitted into register OUT4. The program has

an established priority (see Fig. 8-1). This program is entered on an inter-

rupt caused by an "endpulse" from the telemetry system. Relay words have

the highest priority and will be sent down on the next telemetry word. These

relay words contain the state of all latching DSKY relays and therefore indi-

cate displays (display word) and mode status of the G&N and MCP/SCS. Relay

words for flight 202 are listed below.

The maximum rate for relay words is 1 word/120 msec. If a relay

command has occurred, a relay word is loaded into OUT4 and the AGC returns

to whatever program it was in before the interrupt. In a similar way, if no

relay word is used, the AGC checks to see if an "input character word" has

been received (manual keyboard entry, mark, or uplink). The maximum rate

of input character words could occur due to uplink words; this rate is 1 key-

board character/ll0 msee (see section 3.1.2. I). If no relay or input word is

indicated, the AGC checks to see if an IDword is required (there is an ID word

for every block of 4 data words). If no relay, input, or ID word is senLthe AGC

will load OUT4 with the next data word to be sent. A list of the data words

While it is theoretically possible to get almost 18 relay and input character

words/sec leaving only 32 words/sec for both ID and data words, it is estimated

that the AGC will average at least 32 data words/sec and 8 ID words/sec.

8 -1

TELEMETER

RELAY

WOR D

( MODE S &

DISPLAYS )

i

EXIT

TELEMETER t

INPUT

CHARACTER

WORD

EXIT

TELEMETER LID

WOR D ,F

EXIT

YES

YES

YES

PULSE

I D \.

I TELEMETERDATA

WORD

EXIT

Fig. 8-i AGC Downlink Transmission Logic

8-2

RELAY WORDS

A. Display Words

Item

Vg x Vgy Vg z

Tff

WX

B.

i•

2o

3o

Remark

three components of velocity-to-be-galned duringpowered flight

free-fall time to 300, 000 ft when calculated

W Wy z

Spacecraft Body Rates when calculated

Other Relay Commands

Item Remark

G/N ATT CONTROL SELECT

G/N AV MODE SELECT

G/N ENTRY MODE SELECT

CM/SM SEP COMMAND

+X TRANSLATION ON/OFF

G/N FAIL INDICATION• 05 G INDICATION

GIMBAL MOTOR POWER ON/OFFFDA I A LIGN

T/C ANTENNA SWITCH

MCP/SCS Modes

ZERO ENCODE

COA RSE A LIGN

LOCK CDU

FINE ALIGN

RE-ENTRY

ATT CONTR

ZERO OPT. CDUVs

G&N Modes

CDU ZERO LIGHT

CDU FAIL LIGHT

PIPA FAIL LIGHT

IMU FAIL LIGHT

OR OFALLALARMS

COND LAMP TEST

FAILURE & WARNING LIGHTS

8-3

DA TA WORDS

This list is comprised of three groups: Group I is transmitted throughout

the flight; group II is transmitted only in non-powered flight; and group III is

transmitted only in powered flight.

Point Measured Remarks

Group I

Time I AGC Timing Register

Time II

IN0

IN2

IN3

AGC Timing Register

Contains keyboard characters, mark, block uplink,inhibit upsinc

Four lowest order time bits, CDU, PIPA, and IMUFail and Parity Alarm, Lift Off, Guid Release,SIVB Separate,

Zero CDU encoders, lock CDU, fine align, re-entry,OPT modes 2 & 3, star present, zero OPT, Coarsealign, A TT SW and TRN SW, Sextant On, OR OFC1-C33

OUT1

Position & Velocity

Engine on; block end pulse; ID word; RUPT trap reset;T/M, program, and program check fail alarms, keyrelease, and computer activity

Six double precision words (12 words in all)

Group II

3 actual CDU counters Used to monitor platform alignment

Group III

PIPA Contents of the three PIP accumulation registers

CDU's (actual anddesired)

6 AGC registers which give actual and desired CDUangles

8.1.2 G&N PCM Telemetry (exclusive of DIGITAL DOWNLINK) and On-Board Recording for Mission #202

OPERA TIONA L

CG0001 V Computer Digital Data PCMD

CGll01 V -28 VDC Supply PCM+

CGlll0 V 2.5 VDC TM Bias PCM+

CG1503 X IMU +28 VDC Operate PCME

CG1513 X IMU +28 VDC Standby PCME

CG1523 X AGC +28 VDC PCME

50 s/sI

1

I0

i0

i0

(See 8.1.1)

8-4

CG1533

CG2110

CG2112

CG2113

CG2117

CG2140

CG2142

CG2143

CG2147

CG2167

CG2170

CG2172

CG2173

CG2177

CG2206

CG2236

CG2264

CG2266

CG2300

CG2301

CG2302

CG2303

CG3102

CG3112

CG3200

CG3209

CG3220

CG3229

OPERA TIONA L (Cont'd)

X OPTX +28 VDC PCME

V IGA Torque Motor Input PCM

V IGA IX Res Output, sine,

inphase PCM

V IGA IX Res Output, cos,

inphase PCM

V IGA Servo Error, inphase PCM

V MGA Torque Motor Input PCM

V MGA IX Resolver Output,

sine inphase PCM

V MGA IX Resolver Output,

cos, inphase PCM

V MGA Servo Error in Phase PCM

V OGA Servo Error in Phase PCM

V OGA Torque Motor Input PCM

V OGA IX Resolver Output,

sine inphase PCM

V OGA IX Resolver Output,

cos, inphase PCM

V OGA Servo Error, in Phase PCM

V IGA CDU IX Res Error,

in phase PCM

V MGA CDU IX Res Error,

in phase PCM

V OGA CDU 16X Res Error, PCM+

in phase

V OGA CDU IX Res Error, PCM

in phase

T PIPA Temp. PCM+

T IRIG Temp. PCM+

C IMU Heater Current PCM+

C IMU Blower Current PCM+

V SXT Trun Motor Drive

in phase PCM

V SXT Shaft Motor Drive,

in phase PCM

V Trun CDU Motor Drive

in phase PCM

V OPTX Direct Trunnion

Contlr, in phase PCM

V Shaft CDU Motor Drive

in phase PCM

V OPTX Direct Shaft Contlr

in phase PCM

10

10

10

10

100

10

10

10

100

1

10

lO

10

100

1

1

10

1

1

1

1

1

10

10

10

10

10

10

8-5

8.2

OPERA TIONA L (Cont'd)

CG4300 T AGC Temp. PCM I0

CG5000 X PIPA FAIL PCME i0

CG5001 X IMU FAIL PCME i0

CG5002 X CDU FAIL PCME i0

CG5003 X Gimbal Lock Warning PCME I0

CG5005 X Error Detect PCME i0

CG5006 X IMU Temp. Light PCME i0

CG5007 X Zero Encoder Light PCME I0

CG5008 X IMU Delay Light PCME i0

CG5020 X AGC Alarm #I (Program) PCME i0

CG5021 X AGCAlarm #2 (AGC Activity) PCME i0

CG5022 X AGC Alarm #3 (T/M) PCME i0

CG5023 X AGC Alarm #4 (PROG CHK

FAIL) PCME I0

CG5024 X AGC Alarm #5 (Scalar FAIL) PCME I0

CG5025 X AGC Alarm #6 (Parity FAIL) PCME i0

CG5026 X AGC Alarm #7 (Counter FAIL) PCME i0

CG5027 X AGC Alarm #8 (Key Release) PCME I0

CG5028 X AGC Alarm #9 (RUPT Lock) PCME I0

CG5029 X AGC Alarm #i0 (TC Trap) PCME i0

CG5030 X Computer Power Fail Light PCME i0

CG6000 P IMU Pressure PCM 1

FLIGHT QUA LIFICA TION

CG2010 V XPIPASG. Output, inphase TR 2000 eps.

CG2030 V Y PIPA SG. Output, inphase TR 2000 cps.

CG2050 V Z PIPA SG. Output, inphase TR 2000 cps.

CG6001 D NAV Base Roll Vibration TR 2000 eps.

CG6002 D NAV Base Pitch Vibration TR 2000 cps.

CG6003 D NAV Base Yaw Vibration TR 2000 cps.

External Data Requirements

G&N requirements for external data fall into three categories:

8.2. 1 Navigation Data via the Uplink

No requirement for this data is made at this time.

8.2.2 Radar Tracking Data for Post Flight Analysis

Tracking data requirements to a degree of accuracy and completeness

which would permit the most comprehensive determination of G&N flight per-

formance, are given in Table 8-1. Subsequent revisions of this plan will

reflect more realistic requirements.

8-6

8.2.3 Radar Tracking Data for Real-Time Monitor of G&N

This requirement is given by Table 8-2, which is derived from the

total indication error expected in the position and velocity data telmetered

to the ground via the AGC DOWNLINK.

8-7

TA BLE 8-1

EXTERNA L TRA CKING DA TA REQUIREM ENTS

TO SUPPORT POST FLIGHT ANA LYSIS OF G&N

Three orthogonal components of position and velocity are required in IMU

coordinates at one second intervals during each powered phase. The re-

quired accuracies are given in this table in local vertical coordinates.

Phase

one sigma

Position Error (ft)

one sigma

Velocity Error (fps)

Alt. Track Range Aft. Track Range

S-IB Boost 200 1900 i00 0.9 7.2 0.4

ist SPS Burn 40 210 30 0. 3 i. 8 0.2

2nd, 3rd, 4th SPS Burns I0 30 i0 0.3 0.6 0.2

Entry 1100 1000 200 4.6 4.9 0.8

8-8

TABLE 8-2

EXTERNAL TRACKINGDATA REQUIREMENTSTO PROVIDEREAL-TIME MONITOROF G&N

Three orthogonalcomponentsof position andvelocity are required in IMUcoordinatesat one secondintervals during eachpoweredphase. The re-quired accuraciesare givenin this table in local vertical coordinates.

onesigmaPosition Error (ft)

one sigmaVelocity Error (fps)

Aft. Track Rang_ Alt. Track Range

S-IB Boost 200 1900 I00 0.9 7.2 0.4

ist SPSBurn 400 3900 300 i. 4 8.0 0.8

2nd, 3rd, 4thSPSBurns 2300 4000 7200 7.3 7.4 1.8

Entry 2900 7300 8800 ii. 6 3.9 2.9

8-9

b

9. G&N Performance Analysis

This section presents brief summaries of the performance of those phases of Ii_.

202 mission that are under G&N control. The data (in Fi;_s. 9-i, 9-2, 0-3, 9-i, ?-7,,

and 9-6) has been derived from point mass studies using the Saturn Boost phas_ _ o_ the

trajectory referenced in Section 5.

The data in the Tables 9-I through 9-8 present performance data derived by

perturbing the nominal mission with the dispersions listed on the following page.

The affects of these dispersions are demonstrated in the tables as follows:

Table 9-1 Time, latitude, longitude, altitude, velocity, flight path angle and

range (central angle from SIVB eut-off point) at the start of the firstSPS burn.

Table 9-2 Same as Table 9-1 at the end of the first SPS burn, plus fuel remain-

ing and burn time.

Table 9-3 Time latitude, longitude, altitude, velocity flight path angle, R, A,and E from Carnarvon at the start of the second SPS burn.

Table 9-4 Same as Table 9-3 at the end of the second SPS burn.

Table 9-5 Same as Table 9-8 at the final cut off.

Table 9-6 Time latitude, longitude, altitude, velocity flight path angle at entry

after fourth burn or fuel depletion.

Table 9-7 Velocity and flight path angle at entry without the two short burns.

Table 9-8 Same as Table 9-6 after the first burn only.

The radar at Carnarvon was taken to be at 24. 867S latitude and liB. 63E longitude

at a radius of 20, 913,669 feet.

The latitude and longitude at entry in Table 9-7 above will be practically the same

as Table 9-6 above.

Fig. 9-6 shows the track during the nominal second SPS burn and the two short

burns. The ignition point and final cut off points of extrem_ cases are also shown. It

should be observed that

a) The maximum westerly dispersion at ignition is about 0.5 ° longitude.

b) The dispersion in track (213036) cannot be rectified by modification of

the second ignition logic.

Any downrange dispersion at SIVB cut-off will move the entire trajectory downrange

by the amount of dispersion.

D

9-1

Mac Run

210066

210067

210068

210069

210070

210071

210072

210073

210074

210075

210076

210077

210078

210890

210891

211683

213036

List of Dispersions

617.

617.

617.

617.

617.

617.

617.

617.

617.

617.

617.

617.

617.

60O

4 see

4 sec

4 sec

4 sec

4 sec

4 sec

4 sec

4 sec

4 sec

4 sec

4 sec,

Dispersions

+ 200']sec inertial velocity

+ 40']sec inertial velocity

-40'/see inertial velocity

+ 3000 ft altitude

- 3000 ft altitude

+ 0.5 ° flight path angle

- 0. 5 ° flight path angle

+ 3 sec Isp

-3 sec Isp

+ 660 Ibs thrust

-660 ibs thrust

4 sec, + 500 ibs weight

4 sec, - 500 Ibs weight

sec, + 30, 000 ft. altitude

-2 ° flight path angle

-3 sec Isp

- 660 ibs thrust

+ 500 Ibs weight

600 sec, negative of above

617.4 see, nominal

617.4 sec, +i ° azimuth

-i. 63 southern latitude

NOTE: i. Nominal I was increased by 3 seconds over the Novembersp

figure.

2o AI] cases have an ii second coast between SIVB time indi-

cated and SPS 1 ignition.

3. Altitude is in feet

Velocity is in ft/sec

All ang]es are in de_rees

Time is in seconds" _ot_l time is measured from lift-off

Range [_1om Carnarvon is slant range in n.m.

Radius of earth used in 20, 925,738 feet.

The coast time used is 3041 seconds

Precision integration was used during coast

9-2

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'_ co _I4 c,o o_ (:_ ',_0 00 0 0

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_ 1.0 _ u'3 LO LO L('b

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LO LO LO

CO _ 0') 0 0 0 0 0

_0 _ _ _ CO _ _ ",/3

6 c_ (5 6 c_ _5 c_ c_

t_- L'.- 0 0

C.O c..o _0

• ° ° °0 ¢) ¢_ 0

C0

,-.lo,1 ¢',,Io,.I o,I _ o,1

c',l _ 00 D--

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0 0 0 0 0 0 0 00,1 ¢'_ o,] O,1 0_ o,1 0,1 o,1

0,] 0_ 0,1 ¢_ _ 0,] o,1

o3 CO 0 COI/3 LO 0,1 tO

• ° ° °

0,1 0,1 0,1 0,1

0

0,1 0,1 0,1 0,] 0,1 _ 0,1 0,1 ,_

• . ° o

¢.0 ¢.0 _0 ¢.0

C..O _ CO _ ['-.. L"- L'-'- L"- 'L"W" Cb (_) CO 0'_0 C' 0 0 0 0 ¢_, 0 0 C) CO CO _C) 0¢_' ¢) 0 ¢) C, 0 ¢_' 0 +" (_' ¢) C) _ O'3

9-9

!

0

0

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[

ct_ _ oO _ _ _ CO _ L_- CO _ O_ 0"_ CO CO _ 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 0 0 0 0 0 _

9 -10

0

I

0

c_

_ O_ _ CXI C_ CXI C_ 0,1 CXl 0,1 O_ _ 0,1 0,1 _'_ CXl

I I I I I I I I I I l I I I I I I

I I I I I I I I I I I I I I I

_0

0

0

_ _ _ 0 _ _ _ _ _ _ _ _ _ _ _ 0

0 0 0 0 0 0 0 0 0 0 0 0 0 oo oo co 00 0 0 0 0 0 0 0 0 0 0 0 0 0 0 _ co

9-11

I

,.Q

cO _ _ c_- 0 oo q_ ,--4 Go _1_ _ b.- _ _ b '= _ _'_o'_ _ 0 _ C-- _ _ _ _ _ _ L_- L_- _ _ _ CO

_ ¢0 _._ _ _ _ _ _ L_- CO _ _ _ _ CO

c_ _ C_- C-.= _ _ _ L_- L_- _ _ _0 _ _0 _ _

I I I I I I I I I I I I I I I I I

L_- _ b- L_- _ _ L_= _'- L_" _ _ I_- _ i.(3 rE) _ _.

O _'_ _ 0 _ _ _ _ L_" _ _ _ 0_ _ _q_ O3

3_0 o_ _ _ _0 _o L_ ,_ ¢_I 0 _0 L_- ¢_ 00 0_ L_ O0

CO L_- _ "_ b" _ 0 _ ¢_ O_ _ b- LO ¢_ "_ _0

0 0 ¢_ _ 0 0 _ O 0

00 O_I

_0 L'_ _0 _ 0 _'_ C_I ¢0 _ _ _0 E_- O0 0 ,'_ _0_0 _0 _0 ¢.0 L_" L_" _ L" E" _ _ L_" L"- O_ ¢_ O0 O_

0 O 0 O 0 0 _ 0 ¢_ _ 0 _ O 0 _ _ ¢0

9-12

0

0

r,.)t_

LO 0 _ CO O_ _ C_ _ _._ ,._ O_ _ _ CO

0 0"_ _ O_ 0 _ QO CO L_" _ 0'_ _ CO

CO 0 _ CO _ _ 0 _ C_ CO 0'_ CO _

I I I I I I I I I I I I I I

oO _ C'_ L_ ,--_ _.- _ _ _ CO O0 _ CO0 _ _ CO _ _ _ _ _ _ ,-_ _ 0 _"

CO CO CO CO 0'_ _ CO CO O0 _ CO _ CO

L_- _ _ _- 12"- _ _ L_. 12"- _ _ _ _ L_-

_ _ _ _ _ '_ CO _ LO _ _ L_- '_

o=

VVVVVTVT VVV VVV

O0 _- _ "_ L_. _ _ ,-_ _0 O_ _ _ h_ _0

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

o o0 _- _- _- oo ._ _o _ _ o _ _- _-

_ _ _ _ _ _ _ _ L_- L_- [_- 12.- CO,0 0 0 0 0 0 0 0 0 0 0 0 0 _._;I

0 0 0 0 0 0 0 0 0 0 0 0 0 ,--_

9 -13

O

o_

O'_ O O O O O _ O O O O ¢Xl O'_ L_ 00 O O

I I I I I i I l I I I I I I I I I

>CO L_ ¢N1 O L"- _ L_ _ O'_ _ Ca CO CO O _ L_

O O_ O'_ 00 CO O'_ O_ O'_ 00 C_ O'_ O_ O _ _ COb- ¢.,.C; _ e,.C. _ ¢..O e..C; ¢_ ¢4D _ e_ e,O t"- _ ¢_ _ ¢D

CO CO CO CO CO CO CO CO O0 CO CO CO O0 _ b" CO CO

C',1 ¢"xl ¢N1 _ C_1 C',1 C'd C_1 ¢",1 ¢X1 C',1 _ _ C_1 _ C'd ¢N1

00 O_ C_ _ _ _ _0 _ ¢._ _ _ O O'_ _ O O0

"_ _ _ _ L_ _ _ _ _ L_ L_ L_ L_ _ C'_ L_

I I I I I I I I I I I I I I I I I

I

o

'-o

o=

> O O _ co O_ '_ O O'_ O'_ _ ¢'_ _-_ _'_ _ co Ot _. b- _ ¢_ ¢.O ¢._ _ _ _ C.C) cad ¢._ _ _ _ ¢._

CO CO CO CO oO CO CO CO CO CO O0 o0 CO _._ L-_ cO 00

__O_ ______O_OO_OO_

b_ t'-- C_ O O_ _ _ _ O C_ _ O'_ O CO oo _'_ O O

o "_ 4 _ ,_ ,_ ,# _ ,_ _ ,_ _ _ ,_ _ ,_ _ _

"_ _ _ O _ _ _ _ O _ O COT_ C-,1 C_

I I I I I I I I I I I I I I I I I

C',3 CN1 CN1 CN1 O ¢'_ _ C-,.1 _ _ ON1 _ CO erj _

O O O O O O O O O O O O O CO oo _ O

O O O O O O O O O O O O O O O _ ¢'_

9-14

Table 9-7

Entry Conditions(400,000ft) After SecondBurn(noshort burns)

MacRun V 7

210O66210067210068210069210070210071210072210073210074210075210076210077210078211683213036

28497 -3. 57-3. 57-3. 57-3. 57-3. 57-3. 57-3. 56-3. 57-3. 57

-3. 57

-3. 57

-3. 57

-3.56

-3. 57

-3. 57

28518

28514

28502

28501

28518

28515

28516

28516

28508

28520

28506

28526

28501

28514

9-15

cO m 01 _ 0

O_ 0

_ O_

N

I I I I I I I I I ! I I I I I

> 0",1 C_ 0,1 0,,1 0,.1 _ 04 _ 0,1 0.,1 C',,1 _ Ol _'1

C_ _'_ @,1 C'4 @,1 _ C'_ C_1 _ 0'.1 0,1 _ 0',1

0"_ L"- _t _ C'_ CO CO CO L_ _,0 CO I_ _ 0"_ LO 0

o ,.6 _6 ;H _6 _ _ _ _ u3 _ _ _ _6 _ _

I I I I I I I I I I I I I I I

_.D _ _ C.O _ _ _ _ _ _ _ _ _ CO CO tO O O O O O O _ O O O O O _,O OiO O O O O O O O O O O O O _ c'_

O

O

O

O

OOO

c_O

.o

O

O)

9-16

R-477

DISTRIBUTION LIST

Internal

R. Alonso

R. Arrufo

R. Baker

R. Battin (5)

P. Bowditch

D. Bowler

R. Boyd

E. Copps

R. Crisp

J. Dahlen (5)

E. Duggan

K. Dunipace (MIT/AMR)

J. B. Feldman

S. Felix (MIT/S&ID)

J. Flanders

J. Fleming

G. Fujimoto

F. Grant

Eldon Hall

Edward Hall

E. Hickey

D. Hoag

A. Hopkins

F. Houston

L. B. Johnson

M. Johnston

A. Kosmala (3)

A. Koso

M. Kramer

A. Laats

L. Larson

J. Lawrence (MIT/GAEC)

T. J. Lawton (2)

T. M:. Lawton (MIT/MSC)

D. Lickly

H. Little

G. Mayo

J. McNeil

H. McOuat

R. Morth

James Miller (2}

John Miller

J. Nevins

J. Nugent

E. Olsson

J. Rhode

M. Richter

M. Sanders

M. Sapuppo

R. Scholten

E. Schwarm

J. Shillingford (3)

W. Shotwell (MIT/ACSP)

J. Sitomer

B. Sokappa

M. Sullivan

J. Suomala

R. Therrien

W. Toth

M. Trageser

R. Weatherbee

L. Wilk

R. Woodbury

W. Wrigley

Apollo Library (2)

MIT/IL Library (6)

External

(ref PPl-64; April 8, 1964)P. Ebersole (NASA/MSC)

W. Rhine (NASA/RASPO)

L. Holdridge (NAA SAID/MIT)T. Heueremann (GAEC/MIT)

AC Spark PlugKollsman

RaytheonMajor W. Delaney (AFSC/MIT)

NAA RASPO: National Aeronautics and Space AdministrationResident Apollo Spacecraft Program OfficeNorth American Aviation, Inc.Space and Information Systems Division12214 Lakewood BoulevardDowney, California

FO: National Aeronautics and Space Administration,Florida Operations, Box MSCocoa Beach, Florida 32931Attn: Mr. B. P. Brown

HDQ: NASA Headquarters600 Independence Ave., SWWashington 25, D.C. 20546Attn: MAP-2

AMES:

LEWIS:

FRC:

LRC :

National Aeronautics and Space AdministrationAmes Research CenterMoffett Field, CaliforniaAttn: Library

National Aeronautics and Space AdministrationLewis Research CenterCleveland, OhioAttn: Library

National Aeronautics and Space AdministrationFlight Research CenterEdwards AFB, CaliforniaAttn: Research Library

National Aeronautics and Space AdministrationLangley Research CenterLangley AFB, VirginiaAttn: Mr. A. T. Mattson

GSFC : National Aeronautics and Space Administration

Goddard Space Flight Center

Greenbelt, Maryland

Attn: Manned Flight Support Office Code 512

MSC

(2)

(1)

(1)

(1)

(10)

(10)

(10)

(I)

(1)

(3)

(6)

(2)

(2)

(1)

(2)

(2)

MSFC:

ERC:

GAEC :

NAA :

GAEC RASPO:

ACSP RASPO:

WSMR:

MSC:

National Aeronautics and Space Administration

George C. Marshall Space Flight CenterHuntsville, AlabamaAttn: R-SA (10)

L. Richards (10)

National Aeronautics and Space AdministrationElectronics Research Center

575 Technology SquareCambridge, MassachusettsAttn: R. Hayes/A. Colella

Grumman Aircraft Engineering CorporationBethpage, Long Island, New YorkAttn: Mr. A. Whitaker

North American Aviation, Inc.

Space and Information Systems Division12214 Lakewood Boulevard

Downey, CaliforniaAttn: Mr. R. Berry (40)

Mr. L. Hogan (10)

National Aeronautics and Space AdministrationResident Apollo Spacecraft Program OfficerGrumman Aircraft Engineering CorporationBethpage, L. I., New York

National Aeronautics and Space AdministrationResident Apollo Spacecraft Program OfficerDept. 32-31AC Spark Plug Division of General MotorsMilwaukee I, Wisconsin

Attn: Mr. w. Swingle

National Aeronautics and Space AdministrationPost Office Drawer MM

Las Cruces, New MexicoAttn: BW44

National Aeronautics and Space Administration

Manned Spacecraft CenterApollo Document Control GroupHouston 1, Texas 77058

Mr. H. Peterson

Bureau of Naval Weapons

c/o Raytheon Company

Foundry Avenue

Waltham, Massachusetts

Queens Material Quality Sectionc/o Kollsman Instrument Corp.

[Building A 80-08 45th AvenucElmhurst, New York 11373Attn: Mr. S. Schwartz

Mr. H. AnschuetzUSAF Contract Management District

AC Spark Plug Division of General MotorsMilwaukee, Wisconsin 53201

(20)

(i)

(i0)

(50)

(i)

(I)

(2)

(lOO)

(i)

(i)

(i)