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MANNED ORBITING LABORATORY 

PROGRAM PLAN 

VOLUME 2 OF 2 

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VOLUME 2 

TABLE OF CONTENTS 

Page 

Number 

Preface 5 

Section I - Laboratorz. Vehicle, Support Module and 

System Integration 

Part 1. System Description and Performance 

Requirements 

A. General 7 

B. Technical Description 7 

Figure I-1 Lab Vehicle Bay Layout • • • 17 

Figure 172 Mission Simulator Floor 

Layout 18 

Part 2. Operating Concept 19 

Part 3. Schedule Consideratidns 

A. Critical Factors 23 

B. Laboratory Vehicle Development Schedule . . . 25 

Figure 1-3 LV Schedule 32 

Part 4. - Funding z,,....irements 

A. General 33 

B. Defer141 Status . • • • 35 

C. Identified Program Chix - . . • • • • • 36 

Figure 1-4 LV and Integration 

Fundir.g Requirements . . 39 

Section II - Miss. payload System Segment Integration 

Part 1. - System Description and Performance 

Requirements 

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A. General . 40 

B. Technical Description 41 

Figure II-1 Mission Payload System 

Segment Sketch • • • • 51 

Figure 11-2 Subsystem Critical 

Functions Matrix • . . 52 

Part 2. Operating Concept 53 

Part 3. Schedule Considerations 55 

Figure 11-3 MPSS Integration 

Schedule 58 

Figure 11-4 GE Integration 

Test Flow. •k• • • • • 59 

Part 4. - Funding Requirements 

A. General 60 

B. Deferral Status 63 

C. Identified Program Changes 6 

Figure 11-5 MPSS Funding 

Requirements 

Section III - Photographic System 

Part 1. - System Description and Performance 

Requirements 

A. General 67 

B. Technical Description 67 

Part 2. Operating Concept 70 

Part 3. Schedule Considerations 71 

Figure III-1 EKC Test Flow . . 74 

Figure 111-2 Photographic System 

Development Schedule. 75 

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Part 4. - Funding Requirements 

Page 

Number 

A. General 76 

B. Deferral Status 78 

C. Identified Program Changes 78 

Figure Photographic System 

Funding Requirements 82 

Section IV - Gemini B 

Part 1. - Segment Description and Performance 

Requirements 

A. General 83 

B. Technical Description 83 

Part 2. Operating Concept 87 

Part 3. Schedule Considerations 89 

FigureTV-1 Gemini B Development 

Schedule 92 • 

Part 4. - Funding Requirements 

A. General 93 

B. Deferral Status 94 

C. Identified Program Changes 95 

Figure IV-2 Gemini B Funding 

Requirements 101 

Section V - T-IIIM Launch Vehicle 

Part 1. Segment Description and Performance 

Requirements- - 

Part 2. Operating Concept 

Part 3. - Development and Schedule Considerations 

A. T-IIIM Airframe 

B. Liquid Rocket Engines 

102 

104 

105 

109 

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C. Solid Rocket Motors 113 

D. Inertial Guidance System 115 

Figure V-1 T-IIIM Development 

Schedule 119 

Part 4. - Funding Requirements 

A. General 120 

B. Deferral Status 121 

Figure V-2 T-IIIM Funding 

Requirements 123 

Section VI - Flight Crew Equipment and AMD Support 

Part 1. - Flight Crew Equipment 

A.. Pressure Suit Assembly 124 

B. Flight Crew Diet 126 

Figure VI-1 Pressure Suit. Development 

Schedule 129 

Figure VI-2 Feeding System 

Development Schedule . 130 

Part 2. AMD Development and Operational Support . . 131 

Part 3. Flight Crew Equipment and AMD Support 

Figure V/-3 Funding Requirements . . . 136 

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June 15, 1967 

VOLUME .11 

PREFACE 

MOL SYSTEMS SEGMENT REQUIREMENTS 

This volume describes in detail, the requirements and technical 

aspects of the MOL development program. Based on a general technical 

description and performance requirements for each MOL system segment, . 

the development, test and manufacturing schedules are discussed in 

the light of critical design and system problems, and in terms of 

critical program milestones. The funding requirements reflect the 

consequence of the critical schedule constraints, technological 

feasibility and testing philosophy. 

For the purpose of a clearer understanding of contractor responsibilities and the tasks that need to be performed in developing the 

MOL System, this volume is divided into system segments according to 

contractor involvement. The tasks performed by the integrating 

contractor and the other associate contractors in areas of teat operations, flight operations, recovery, crew training, and other support 

are included in themajor system segment discussions. The support 

activities treated separately are those which are accomplished by 

groups other than the major associate contractor. 

The sections discussed herein are as follows: 

I Laboratory Vehicle, Support Moduliand System Integration 

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II Mission Payload System Segment Integration 

III Photographic System 

IV Gemini B 

V Titan IIIM Launch Vehicle 

VI Flight Crew Equipment and AMB Support 

Each section is treated in terms of general description, operating, 

concept, critical factors, schedule considerations, and funding 

requirements. Only the basic MOL program is discussed. Al]..advanced 

studies and technology activities are treated separately in Volume I, 

Section V. 

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SECTION I 

LABORATORY VEHICLE, SUPPORT MODULE AND SYSTEM INIAGRATION 

(Douglas Aircraft Company) 

Part 1. Segment Description and Performance Requirements 

A. General 

The Laboratory Vehicle is composed of two modules: 

1.— A Laboratory Module, which is approximately 10 feet 

in diameter and 19.0 feet long, has an unpressurized section and a 

1,000-cubic foot pressurized compartment designed to provide a shirtsleeve environment for a two-man crew to accomplish a 30-day mission, 

in a nominal elliptical orbit of 80/180 N.M. at 90 degreea inclination. 

The pressurized compartment is protected by a meteoroid shield which 

also houses the space radiator. 

2. An unpressurized Mission Module structure, 10 feet 

in diameter and 37.0 feet long, located aft of the Laboratory Module. 

This module contains the Mission Payload, which is controlled and 

operated by the crew from the laboratory. 

B. Technical Description 

The MOL program involves two separate and distinct Laboratory 

Vehicle configurations which provide .a capability to operate in two 

modes -- manned or unmanned. In the unmanned configuration, about 

1000 pounds of major components and subsystems required for manned 

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operation are removed, some components unique to unmanned operation 

are added, and provisions are made to enable the Laboratory.Vehicle 

to interface with the Support Module which replaces the Gemini B. 

A feasibility and conceptual design study has been completed for a 

1 Support Module capable of supporting up to 60 days of orbital operation, although the baseline program contains only a 42-day unmanned 

capability. /A 42-day operation is possible with the baseline expendables leading. The Subsystems will be qualified for a 30-day operation only, and the remainder will be flown on an open ended basis. 

In order to preserve the option for growth up to 60 days' capability, 

the Support Module will be designed to accommodate this growth and 

certain critical wearout components in the Laboratory Module will be 

validated or modified for 60 days operation. Definitive roles and 

responsibilities for the Support Module have been assigned as follows: 

DAC will design and fabricate the Support Module structure, in- 

) 

eluding provision for additional expendables required for the extended 

mission. Support Module integration is also a DAC responsibility. 

1 

GE will design and fabricate the Data Return Vehicles (DRV). 

EKC will design and fabricate the fibs chute used to load the 

DRV's. 

Since most of the other system segments interface with the 

Laboratory Vehicle, MOL System Integration has been included in this 

segment and has been assigned to the Laboratory Vehicle contractor. 

The Laboratory Module contains subsystems oi major elements to 

enable the crew to do the necessary housekeeping and vehicle management 

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and to operate the Mission Payload. An environmental control and life 

support subsystem provides a habitable atmosphere in the Laboratory 

Module, thermal control for the orbiting vehicle, environmental control 

for space suit operations, and management of water and waste. A 

two-gas atmosphere consisting of 3.5 psia partial pressure of oxygen 

and a helium diluent, is maintained at 5.0,'10.2 psia. In a backup 

mode the compartment can be pressurized with 100% oxygen at 5.0 psia. 

A gaseous.oxygen accumulator is provided to maintain nominal cabin 

pressure in case of Accidental decompression through a hole as large 

as 1 inch in diameter,for a period of 5 minutes. 

The cabin atmosphere can be controlled to any temperature 

between 68 degrees and 78 degrees F primarily through use of watercooled instrument panels. Dew point temperature is controlled between 

36 degrees and 60 degrees F. Thermal control is provided by two 

coolant loops. The internal loop, containing water as the coolant, 

removes heat from electronic and electrical equipment, Mallon Payload 

equipment located in the Laboratory Module, and the instrument panels 

and transfers this heat to the external loop through an interface heater-, 

to the Gemini B for on-orbit storage, to the waste management system, 

for waste processing, and to the cryogenic heat exchanger. The 

external loop, containing Freon 21 as the coolant, radiates heat to 

space through a radiator. Carbon dioxide is controlled at a nominal 

5mm to Hg partial-pressure by redundant molecular sieve beds. Trace 

contaminant control is also provided. For water management, potable 

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water produced by the fuel cella is collected and thermally conditioned 

prior to delivery for use by the crew. Excess fuel cell water and 

atmospheric condensate are dumped. overboard. Crew waste products are 

collected using positive air flow control, are processed and the 

residue is stored in sealed containers. 

This subsystem also includes the atmosphere and.reactants supply 

subsystem group which provides storage and pressure control of liquid 

hydrogen and liquid.oxygen for the fuel cells and liquid oxygen and 

gaseous helium for the atmosphere. 

The electrical power subsystem employs.three modified Apollo 

design fuel cells as the primary power source to supply an average 

net power output of 1.83 kw for 30 days at a nominal 28v dc level with 

a peak power capability of 4.5 kw. Normal operation consists of two 

fuel cells operating with the third cell in a "hot" standby condition. 

The standby fuel cell replaces one of the operating cells through a 

time sequenced switching operation. However if any single fuel cell 

fails, the two remaining units can supply the required power including 

peak loads. This subsystem also has an emergency mode which supplies 

28v dc to critical loads from an auxiliary bus. 

The data management subsystem consists of the data acquisition, 

data computation, command, and timing subsystem groups. The data 

acquisition subsystem group contains the vehicle telemetry which 

processes all Laboratory Vehicle data. This subsystem group interfaces 

with the computer and all Orbiting Vehicle subsystems. It provides 

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for the acquisition of data required for crew safety, status indication, fault isolation and for ground evaluation of system performance, 

both during and post flight. 

The data computation subsystem group consists of two identical 

airborne digital computers, one auxiliary memory unit, one keyboard 

and display and control unit, two printers, two Laboratory Vehicle 

data adapter units and one computer subsystem controller. It interfaces, communicates with and performs data processing, computing and 

control functions for the Laboratory Vehicle and Mission Payload 

system segmentt. Each of the airborne digital computers has a memory 

.capability of approximately 16,000 words (word length -- 32 bits) with 

a growth capability to 24,000 words.' The computer memory has a 

- storage capacity of two million data bits and can transfer a minimum 

of 600,000 bits of information per minute between itself and the 

computer. Either computer can support all currently planned on-board 

computation functions. The command subsystem group accepts digital 

encoded uplink command messages which are executed when received 

(real-time commands) or are stored in the memory with a time label and 

executed at the time indicated by the label. Both clear and 

encrypted data and Commands are accepted, however the clear mode can 

only be initiated by a secure command. The data transmission rate 

is 1000 bits per second, which is compatible with the primary SGLS 

command uplink. A backup command capability is provided for a limited 

number of critical commands required for control of the vehicle in the 

event the computer subsystem is disabled. The timing subsystem group 

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provides all Laboratory Vehicle and Mission Payload timing requirements. This includes generation of synchronization signals and 

supplying binary vehicle time for on-board displays, data correlation 

and vehicle elapsed time indicators. 

The instrumentation subsystem provides the instrumentation and 

signal conditioning required for operation of .the Orbiting Vehicle 

and Mission. Payload through the displays and controls. A transducer 

subsystem group provides 173 transducers (not including redundancy) 

for monitoring all life-critical and mission-critical parameters and 

performance characteristics of the Laboratory Module subsystems. 

The signals from approximately 240 measurements in the Laboratory 

Module require conditioning prior to being presented to the displays 

and controls subsystem group, the monitor and alarm subsystem group 

and/or the data management subsystem. The monitor and alarm 

subsystem group continuously monitors 180 measured parameters to detect 

out-of-tolerance conditions. It provides visual and aural signals, 

and develops control signals where required, for out-of-limit conditions in all monitored functions. Twenty life safety parameters 

which are redundantly monitored are allocated to "warning" channels. 

The remaining 160 monitored functions are allocated to "caution" 

channels to alert the crew to mission critical or noncritical out-oftolerance conditions, including those which could become critical 

to crew safety if corrective action were not taken. The displays and 

controls subsystem group provides the man/machine interface for 

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operation of the Laboratory Vehicle and Mission Payload. It also 

provides data for status indication and telemetry. Approximately 

515 system parameters are displayed and approximately 663 are 

controlled from the display and control panels. 

The communications subsystem services both the Laboratory Vehicle 

and the Mission Payload. The tracking, telemetry, command and 

voice functions are provided by redundant SGLS,compatible equipment. 

A voice control subsystem group provides and controls all voice communi• 

cations between crew members. To provide secure voice communication 

between the crew and ground stations, analog-to-digital and digitalto-analog voice converters are being developed to work with modified 

existing airborne cryptographic equipment. .This development and 

modification program is being accomplished by the NSA. The entwine 

subsystem group includes' arth-oriented directional antennas for 

downlink transmission and .one omnidirectional antenna whigh, although 

normally used for Uplink signals, can also be used for downlink 

transmissions. 

An attitude control and translation subsystem (ACTS) will provide, 

without dependence upon the on-board computer: (1) automatic and 

manual attitude control; (2) orbit maintenance and adjust; 

(3) separation from launch vehicle; (4) vernierinsertion capability; 

and (5) Laboratory Vehicle disposal. The ACTS provides for both 

manual and automatic operation gad can be controlled from the Gemini B 

to separate the Orbiting-Vehicle from the Launch Vehicle when initial 

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orbit is attained and for certain abort modes. This subsystem consists 

of two subSystem groups. The ACTS propulsion subsystem group (ACTS/ 

Prop) consists of four removable thruster assemblies located 

90 degrees apart in the unpressurized service compartment of the 

Laboratory Vehicle. Each thruster assembly contains four 25-pound 

thrusters for attitude stabilization and control and one 100-pound 

thruster for longitudinal impulse. This thruster arrangement provides 

sufficient redundancy-to permit non-degraded performance in the event 

of any single thruster or sector failure. The system employs:a 

propellant combination of nitrogen tetroxide and monomethylhydrazine 

(2200 pounds, 2000 pounds usable). The ACTS stabilization control 

electronics subsystem group (ACTS/SCE) provides manual and automatic 

guidance signals to the ACTS/Prop necessary to control the Orbiting 

Vehicle attitude and to change or maintain the Orbiting Vehicle orbit. 

On-board attitude data are automatically generated in the system by 

means of signals from rate gyros, horizon sensors and a velocity 

vector sensor. Correction may also be initiated by ground station or 

manually by the crew. 

Provision is made in the Laboratory Module for an option to 

incorporate a wide-band readout system and/or a Data Return Vehicle 

in Flights 3, 4 and/or 5. Space weight and power provisions ere being 

retained in these flight,velacle -denislon7on-whethear-to-exereise 

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The habitable volume of the Laboratory Module's pressurized 

compartment is concentrated in the central core area. This allows 

the crew members to operate the Laboratory- Vehicle and Misdion 

Payload from the displays and controls with minimum movement. Eight 

compartment bays, which contain the displays and controls, are 

arranged in the following manner: 

BAY FUNCTION 

1 Photographic equipment and vehicle control 

2 - Mission Payload - acquisition, command, 

and instrumentation, electrical power 

and signal distribution 

3 :ECLS (Environmental Control & Life Sup. Sys.) 

4 ECLS, film processor control 

5 Mission Payload control. 

6 Electrical power 

7 Data management 

8 Mission Payload- duplicate of Bay 2 

This arrangement is illustrated in Figure 1-1. vlewport is located 

in Bay 7 for visual horizon orientation.) 

DAC will also provide various equipment and accommodations for 

crew members such as crew conditioning equipment (engineering prototypes being developed and evaluated by the Aero Medical Division), 

suit storage and donning station, restraint and locomotion equipment, 

sleeping equipment a food management installation for OPE food, a 

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medical kit storage compartment for GFE medical components, and a 

personal hygiene/waste management compartment. A radiation 

monitoring system, consisting of both passive and active devices, 

will be provided to measure the amount of incident radiation energy 

impinging upon the crew members. 

The maximum weight allocated by DAC contract specification to 

the Laboratory Vehicle segment (the Laboratory Module and the Mission 

Module structures) 'is 14,449 pounds. 

A mission simulator is required to develop and validate flight 

test procedures and flight plans, develop operational procedures and 

alternate flight plans for contingency situations, validate ground 

command software, and train flight crews and ground controllers in 

combined network operations. This mission simulator is an integrated 

equipment complex made up of a Gemini B Procedures Simulator, Mission 

Payload simulation equipment and Laboratory Module simulation 

equipment. These three elements can operate simultaneously as well as 

independently. The unit will be located at VAFB where the flight crew 

completes its training and where Mission Control Center (MCC) personnel 

are trained. A floor layout of the simulator is shown in Figure 1.2. 

The mission simulator simulates the mission profile from launch 

through re-entry for development, training or validation in separate 

phases or tasks or in the complete mission profile. MCC flight 

directors are trained by a direct interface between the mission simulator 

and the MCC, while remote tracking station operators are trained with 

data tapes generated by the simulator. 

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Part 2 • Operating Concept 

The point of departure for defining the Laboratory Vehicle 

operating concept is the factory-to-pad flow, discussed in Book I, 

Section I, for the segments housed in the Laboratory Vehicle. The 

Laboratory Module will be acceptance tested in a vacuum chamber at 

the Laboratory Vehicle contractor's plant, where a simulated mission 

profile of several days duration will be run. Following acceptance, 

the Laboratory Module is mated with the Mission Module at the 

Laboratory Vehicle contractor's plant, to make up the Laboratory 

Vehicle. Following checkout and acceptance testing there, the 

Laboratory Vehicle is transported to the launch site where it is mated 

with the Titan III M and the Gemini.B. Following the completioi of 

integrated system tests at the launch site and a successful countdown, 

the integrated Orbiting Vehicle is launched into orbit. During launch 

and ascent phases of the mission, the Laboratory Vehicle is 

essentially dormant; it plays a major role in subsequent Mission 

phases, as follows: 

The early orbit phase, which begins when Mission Control 

determines that an acceptable orbital insertion has been accomplished, 

may encompass up to three orbits. This phase is primarily concerned 

with Laboratory Vehicle checkout, flight crew transfer, and orbit 

ephemeris. updating. Upon attainment of an acceptable orbit, the 

flight crew will activate and,check out Laboratory Vehicle subsystems 

to ensure an-acceptable environment. One crewman will then-enter the 

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Laboratory via the crew transfer tunnel that connects the Laboratory 

with the Gemini B. The crewman in the Laboratory will then activate 

and check out equipment while the crewman in Gemini prepares it for 

the 30-day standby mode. When both crewmen have accomplished their 

tasks, the second crewman transfers to the Laboratory through the 

same tunnel. 

The on-orbit phase begins upon completion of initial crew transfer. During the nominal 30-day period the flight crew will operate 

and maintain the Laboratory Vehicle equipment and monitor the status 

of the Gemini B to accomplish the objectives of the mission. Except 

for extra-vehicular activity or emergency cabin depressurization 

periods, the crew will not normally wear pressure suits. Operations 

will be conducted in accordance with a detailed sequence of events. 

This sequence depends on mission objectives, equipment constraints, 

crew eat/rest/sleep cycles, time over stations, day/night conditions, 

etc. If this sequence of events must be modified as the mission 

progresses, the ground system will react to reschedule and replan the 

timeline sequence and transmit this information to the vehicle and 

flight crew. All commands to the crew and vehicle will be from the 

Mission Control Center. Telemetry and voice data from the Orbiting 

Vehicle will be available in real-time and stored modes. Selected 

data transmitted from the Orbiting Vehicle will be displayed in real 

time at the Mission Control Center. Although all data will be 

recorded at the remote tracking stations, none will be displayed there. 

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When the decision to terminate the mission has been made, the crew 

will activate the Gemini B and transfer from the Laboratory to the 

Gemini B in essentially the reverse of the procedure followed at the 

start Of the early orbit phase. When all crew functions have been 

performed, the crew separates the Gemini B from the Laboratory 

Vehicle, re-enters and is recovered. Mission Payload data will be 

returned with the crew in data recovery containers carried in the 

Gemini B. 

The Mission will be terminated by a Mission Control Center 

command to de-orbit and dispose of the Laboratory Vehicle in a preselected open water area. 

The Laboratory Vehicle for Flights 6 and 7 will be assembled 

in the unmanned configuration. The Support Module, Laboratory 

Module and Mission Module will be mated, checked out and accepted 

at DAC prior to being transported to the launch site. The Flight 

Vehicle is then launched and the Orbiting Vehicle, after attain-- 

ment of an acceptable orbit, automatically performs the mission as 

directed by commands stored in the on-board computer and/or 

transmitted from Mission Control. Mission Payload data will be 

returned in Data Recovery Vehicles ejected from the Orbiting 

Vehicle. As in the manned mode, the mission is terminated by 

de-orbit and disposal of the Orbiting Vehicle in a pre-selected 

area. 

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In summary, Douglas as the Laboratory Vehicle System Segment 

Contractor is also the integrating contractor and is responsible 

for the overall integration of the MOL system. In this role, it 

will provide the design engineering required to assure total system 

compatibility. It is also responsible for Orbiting Vehicle System 

wt Engineering and Integration. This task includes the detailed design 

engineering and integration of Orbiting Vehicle interfaces, AVE, AGE, 

special tooling, and special test equipment and the Support Module. 

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Part 3. Schedule Considerations 

A. Critical Factors 

Establishing the development plan for the Laboratory Vehicle 

involves consideration of numerous interrelated activities and 

constraints. The development schedule is aimed at meeting the 

program objective of the first all-up manned/automatic flight in 

December 1970. The Laboratory Vehicle schedule is affected by 

several critical factors which are unique to the program. 

1. Critical Path 

The principal factor is concerned with the critical path 

for the MOL Program, which hinges on'the'Eastman Kodak Company's (EKC) 

development of the Photographic System. The EKC schedule must be 

supported by General Electric's (GE) development of other Mission 

Payload equipment and the GE schedule in turn directly influences 

the Douglas Aircraft Company's (DAC) schedule for the Laboratory 

Module, which interfaces.lith both the GE and EKC segments. The 

following are examples of Laboratory interface subsystems'for which DAC 

must in FY 1968 complete the design, and in some cases fabricate and 

deliver test items to support GE and EKC development schedules: 

a. The first Mission Module Forward Section test 

tests in August 1967. The other structures follow in December'1967 

Istructure, non-prime #1, must be shipped to GE for payload component c) 

(1F), February 1968 (2F), April 1968 (2FX)*, June 1968 (3F), 

October 1968 (4F) and January 1969 (5F). 

* 2FX is a dynamic test article 

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b. Your Mission Module. Aft Section test structures must 

be shipped to EKC.in July and. November 1968 and May and September 1969. 

c. A required AVE •digital computer functional equivalent prototype was provided to GE in.October 1966 for software 

development. 

2. Long Lead Subsystems 

The second .critical factor involves a..number of the 

Laboratory Vehicle flight subsystems, which have been identified as 

long lead items. Although MOL design criteria stress maximum use of 

available hardware components and established technology, design and 

development of these long lead subsystems was begun early in the 

program and is being vigorously pursued.in. order to' meet •requirements 

of minimum weight and/or 30-day life qualification. These long lead 

subsystems include the attitude control and translation subsystem, 

Laboratory Module structure, atmospheric and reactants supply 

subsystem group, electrical power subsystem, and Os data:management 

subsystem. 

3. Mission Simulator 

Another•critical element for which the Laboratory Vehicle 

contractor is responsible is the design and development of the 

Laboratory Module segment of the Mission Simulator and the integration 

of the total mission simulator. This work was initiated immediately 

after ATP, primarily because of its early need date. The simulator, 

to be located at VAFB, must be operational nine months prior to the 

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all-up manned flight in December 1970. It it needed for nine months 

prior to this flight to develop and validate flight test procedures 

and test plans, develop operational procedures and alternate flight 

plans for contingency situations, validate ground command software 

and train ground crews and ground controllers in combined network 

operations. The schedule of development, fabrication and checkout 

of the Laboratory Vehicle segment of the simulator and its integration 

by DAC with other simulator equipment cannot be completely defined 

1 until subsystem and segment designs are well underway. Therefore, 

preliminary or provisional designs and computer substitutes will have 

to be used in the early developMenttests of the mission simulator 

in order to meet the compressed schedule. To meet.the need date of 

March 1970, the design will be frozen by July 1968 to allow release' 

of detail hardware drawings in November 1968. Fabrication and 

assembly of this segment of the simulator will be completed by 

August 1969 followed by 7 months of installation, checkout and System 

1 development testing. The simulator will become operational in 

March 1970 and be available for mission planning, software and pro. 

cedures validation and crew training. Its configuration is to be 

updated by a block change in June 1970. 

B. Laboratory Vehicle Development Schedule 

In addition to the critical factors discussed above, which 

are unique to the program, all of the other interrelated activities 

associated with the Laboratory Vehicle development must be properly 

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time phased to support the first all-up manned flight. The 

Laboratory Vehicle development involves, in addition to the Aerospace 

Vehicle Equipment (AVE), a number of ground test vehicles and AGE 

sets. To generate the overall Laboratory Vehicle schedule, key 

elements are identified and the time required to complete each of them 

is determined. Then, working backward from the date for the first 

all-up manned launch, December 1970, the interrelated .schedules for 

the ground test vehicles, AGE and AVE are aligned and need dates for 

the key elements are determined. The schedule derived in-this 

fashion identifies the critical mileitonas for engineering, test and 

manufacturing activities required in developing the Laboratory • 

Vehicle. These milestones are discussed in the following paragraphs 

and shown in Figure 1-3. 

Laboratory Vehicle No. 3 for the first all-up manned flight 

must be shipped by end-September 1970 to allow two and a half months 

for integration, checkout and launch countdown at VAFB. Sixteen 

months are required for fabrication, assembly and checkout of the 

Laboratory Module, followed by a two and a half month period *lid' 

includes mating with the Mission Module, electromagnetic compatibility 

testing and final checkout of the Laboratory Vehicle. This establishes 

March 1969 as the start date for fabrication of Laboratory Module 

No. 3. 

The Mission Module structure is designed and manufactured by. 

DAC in two sections. .The aft section is shipped to EICC.for installation 

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of the Camera-Optical Assembly and associated AVE. The forward 

section is shipped to.GE for installation of the tracking mirror 

drive, environmental doors, and other GE AVE. It is •then shipped-to 

EKC for installations of the-tracking mirror and mating with the aft 

section. The complete Mission. Module is then returned to DAC .for 

mating with the Laboratory Module and final acceptance testing before 

shipment to VAFE for launch. 

The forward section of Mission Module No. 3.must be shipped to 

GE in April 1969 to allow time for installation of GE •equipment, 

shipment to EKC for installation of the tracking mirror and assembly 

with the aft section and return to DAC for mating with the Laboratory 

Module and final testing.- To meet this.need date,. fabrication and 

assembly of the forward section, which required .8 months, must begin 

in September 1968. Fabrication and assembly of the aft section of 

Mission Module No. 3 must begin in April 1969 to meet a shipping date 

to EKC .of December 1969. 

1 

An Electronic DeVelopment Compatibility Test Unit (EDCTU) 

(priviouely identified as "hot mock-up") •is required for several 

purposes. It-is used inponjunction with the development AGE-for early 

development integration of interfaces between the various subsystems 

assembled together for the first time, between the Laboratory Module 

and its supporting AGE and between the Laboratory- Module subsystems 

and other system segments. This .unit is also indispensable to the 

development and validation of Orbiting Vehicle and ACE ..computer 

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programs and procedures required for on-orbit operations. The need 

date for the Electronic Development Compatibility Test Unit with all 

subsystems installed is November 1968. To meet this requirement, 

prototype hardware of major subsystems is required prior to 

August 1968 for integration into the unit and validation of its 

functional capability. 

One of the major ground test vehicles is the Laboratory Module 

Qualification Test Vehicle (LMQTV). This is a production Laboratory 

Module which is required to demonstrate in advance of the first all-up 

launch that it can be successfully operated through a simulated 30-day 

on-orbit mission while exposed to the on-orbit thermal and vacuum 

environment. This qualification teat will be conducted in a thermalvacuum chamber and will utilize programs and procedures developed with 

the Electronic Development Compatibility Test Unit. The qualification 

test must be completed by the end of May 1970 to free the thermalvacuum chamber for acceptance testing of Laboratory Module No. 3. 

Fabrication of the vehicle must begin in July 1968 in order to complete 

the qualification by that date. 

Following qualification in the thermal vacuum chamber, the LMQTV 

will undergo qualification acoustic testing. Upon completion of this 

test, the LMQTV will be refurbished and.vonverted to the unmanned mods 

for use in Flight 7. Acoustic testing will be restricted to expected 

flight levels since testing at higher levels would preclude use of 

the refurbished structure for flight. 

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The LMQTV schedule described. above establishes an 'operational 

need date of December 1969 for AGE.needed_to support LMQTV checkout 

activities in the Production System Integration Area (PSIA). The 

test flow associated with compact 12.schedule requires additional 

AGE-PSIA to support checkout of Laboratory Vehicles. The schedule 

for Laboratory Vehicle No. 3 establishes an operational need date -

of May 1970 for this additional AGE. In addition to the AGE-PSIA, 

launch checkout AGE is required at VAFB. This is manufactured 

essentially in parallel with the PSIA sets. Cabling is installed at 

VAFB before Flight No. Land the reMainder of the AGE is installed and 

checked out between Flights 2 and 3. 

Another significant requirement in. the Laboratory Vehicle 

development schedule is completion of qualification of all subsystems 

individually, in addition to qualification as part of the Laboratory 

Module Qualification Test Vehicle. In order to complete this prior. 

to environmental acceptance of the first all-up vehicle, it must be 

- accomplished by May 1970. 

Structural test vehicles are required to insure by adequate ground 

testing prior to the first structure flight in April 1970 that the 

Laboratory Vehicle can. accommodate the boost and en-orbit structural 

. environment. A. Laboratory Vehicle primary (load bearing) structure 

will be subjected to 12 months of structural testing (model survey,-

shock, acoustic and burst). In addition, an LMINM primary structure 

must be subjected. to 5 months ofstatic testing. All of this testing 

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must be completed in sufficient time so that any required changes can 

be incorporated into the design of the flight Laboratory Vehicle. 

This establishes a need date of January 1969 for these vehicles. , 

Fabrication and assembly, which require* 12.1/2 months, must start 

in December 1967. 

The above discussion defines the activities required to support 

the first all-up manned launch in December 1970. Laboratory Vehicles 

subsequent to No. 3 will require equivalent periods of time for 

fabrication and checkout and will be scheduled accordingly to meet 

required launch dates. Since the Laboratory Module Qualification Test. 

Vehicle will have fulfilled its purpose by qualification in the 

thermal-vacuum chamber and acoustic faCility in May 1970, it will be 

refurbished and used as the flight article for Launch No. 7. Mission 

payload equipment that had previously undergone qualification testing 

at Grand BRCvill also be refurbished and used for Launch No. 7. 

In addition to the development/test/fabrication activities 

described above, the program is . taking advantage of the NASA S-IVB 

Orbital Workshop for an-on-orbit test of sane specific MOL equipment/ 

crew relationships. The flight date of the S-IVB Orbital Workshop 

permits the performance of these tests in sufficient time so that 

results can be incorporated into the MOL development if appropriate. 

The crew members will perform spec,ific maintenance tasks on 

selected MOL equipment using MOL baseline tools and crew restraints 

to evaluatevn-orbitmaintainability and to establish timelines and 

procedures for performing the tasks. 

• 

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It is planned that a molecular sieve CO2 removal system of the 

MOL environmental control and life support subsystem equipment be 

tested in this effort. Since no molecular sieve will have been 

flown in space prior to its use in MOL, information which will be 

obtained about on-orbit performance of this unit will be a valuable 

bonus. 

The MOL baseline suit-donning/sleep station will also be 

evaluated in the S-IVB Orbital Workshop. This will permit a more 

realistic evaluation of the donning station design and pressure suit 

entry method than can be accomplished through one "G", neutral 

buoyancy, or zero "G" simulations. Suit donning and doffing 

techniques will be evaluated; the envelope required for donning and 

doffing will be verified and the time required for donning in a 

zero "G" environment will be determined. The MOL pressure suit, which 

will be used in this evaluation, is currently planned to be a primary 

suit on the S-IVB flight. Thus a subjective evaluation of MOL unit 

performance as well as signficant operational experience and 

performance data will be obtained. 

A preliminary agreement has been reached with the NASA to conduct 

the majority of the MOL suit qualification testing at the Manned 

Spacecraft Center using NASA facilities and technical assistance. This 

effort will result in program fund savings and enhance the interchange 

of suit technology information. 

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SF •11: 

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Part 4. Funding Requirements 

A. General 

The estimate of Laboratory Vehicle costs is a result of MOL 

Systems Office estimates and Phase II contract negotiations, and 

include an estimate for the schedule adjustment costs. The original 

estimate was built up from detailed labor hour and material costs at 

the subsystem (or equivalent) level using cost data from similar 

space programs such as Gemini and. Apollo. The costs shown are based 

upon the MOL baseline program of three manned/automatic flight 

configuration Laboratory Vehicles, two unmanned/automatic flight 

configuration Laboratory Vehicles, two test flights for structural 

verification, and one Mission Simulator. Also included are funds 

for the Support Module and National Security Agency secure communiCa• 

Lions hardware. 

The major work breakdown structure tasks ate as follows: 

1. Orbiting Vehicle Hardware 

This task includes the material and fabrication costs 

of the Orbiting-Vehicle AVE, AGE, special tooling and test equipment 

and the Mission Simulator and the Support Module. 

2. Laboratory Vehicle Hardware 

This-task includes material and fabrication costs of the 

Laboratory Vehicle subsystems. Also included are Laboratory Vehicle 

AGE, special tooling, special test equipment and assembly and checkout 

of the Laboratory Vehicle at the contractors facility. 

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3. Systems Engineering and Integration 

Includes system engineering and detailed design engineering 

of the Laboratory Vehicle system segment and its subsystems, AGE, 

special tooling, special test equipment, and the associated engineering 

testing and test support. The major subsystems of the Laboratory 

Vehicle segment are the Laboratory Module pressurized and unpressurized 

compartment structures, Mission Module structure, attitude control and 

translation subsystem (ACTS), environmental control and life support 

subsystem (EC/LS), data management subsystem, instrumentation, 

communications, electrical power, and the crew stations and accommodations. Mission Simulator Systems Engineering and Integration includes 

the detailed design engineering of the Mission Simulator, integration 

of associate contractor equipment, and Mission Simulator test and • 

test support. 

4. Services Test Operations 

This task consists of the Laboratory Vehicle contractor 

participation in launch site preparation, acceptance, assembly and 

checkout of the flight vehicle at the launch site, launch, orbit and 

recovery operations, flight planning, data management, and flight 

analysis. Site activation and operation of the Mission Simulator at 

VAFB and in the in-plant support are also included. This task includes 

training and support of flight crew, support personnel, and mission 

control network personnel. 

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5. Management and Administration 

This task consists of Program Control Management, 

Configuration Management, System Effectiveness Management, Material 

Support Management and Procurement and Production Management. 

Funding requirements, by major work breakdown structure, are shown 

in Figure 1-4 and include the negotiated contract amount with an 

estimate of the schedule adjustment cost, deferral amounts, and 

identified program changes. The deferrals are described in 

paragraph B and program changes in paragraph C. 

B. Deferral Status 

The following were deferred from contract negotiations. 

(Costs shown for these items are included in Laboratory Vehicle 

segment cost charts). 

1. Field Test Operations at Vandenberg AFB - $71.9M 

This effort is near end of test flow cycle. Deferral 

eases FY 68 funding requirements and produces better definition of 

contractor activities at Vandenberg AFB. 

2. Simulator O&M - $2.0m 

This item was deferred as part of the Test. Operations 

package for Vandenberg AFB. 

3. Acquisition Tracking System Integration - $2.1M 

This item was deferred pending the outcome of a study 

effort. 

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4. Data Readout System - $1.0M 

The hardware for this item has been deferred. However, 

it is necessary to retain an effort aimed at reserving the capability. 

for re-inserting the hardware at a later date. 

1 5. Support Module - $32.411M 

Deferred so as to ease FY 68 funds requirements. 

6. Ground Station Conversion for Secure Voice Cammunication 

41.532K 

The decision to require MOL Program funding rather than 

SCF funding for the Ground Station conversion was not firm at the 

time of contract negotiation. Consequently, this item was-deferred. 

1 C. Identified Program Changes 

Revised.or new technical requirements and Management 

decisions since December 1966 have resulted in additional tasks. The 

estimated cost and.reasons for these change items are described 

below. The costs are included in the Laboratory Vehiclelegment cost 

charts. 

1. Additional AGE-PSIA - $10.0M 

Differences in test flow requirementa (particularly 

acoustic and vibration testing) now cause an overlap in test and 

production flow which requires additional AGE-PSIA. 

2. Vandenberg APE Cables - $.556M 

Since insufficient time is available for electrical AGE 

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installation between flights 2 and 3, a new set of electrical cables 

must be obtained and installed in advance. 

3. Refurbishment of LMQTV for Flight No. 7 - $8.0M 

Additional damage to LMQTV is anticipated because of 

high acoustic test levels. 

4. Additional Test - $4.5M 

Additional acoustic, thermal vas and other tests are 

now recognized as requirements. 

5. Tower.Mod - $1.2M 

Because of the test flow impacts, a modification is 

required to the service tower at DAC to accommodate a second test 

area. 

6. Mission Module Forward Section - $.222M 

An additional Mission Module Forward Section is needed 

to meet an earlier static testing schedule. 

7. Integration of Sliding Mask - $1.032M 

A sliding mask to provide protection to the optics from 

thermal gradient and contamination is now recognized as a requirement 

and will be provided by GE. The mask must be integrated into the 

Mission Module by DAC. 

8. Interface Hardware- $.293M 

A number of additional exchange interface hardware 

I

substitutes (electrical, thermal and structural) are needed. 

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9. Mission Module Forward Section Doors - $2.95M 

A set of MS'S doors is needed to test the interface with 

the sliding mask. 

10. Reconnaissance and Rendezvous Study - $.100M 

A study to explore certain aspects of reconnaissance and 

rendezvous has been recognized as a requirement. 

11. Water Simulation - $.300M 

An extension of zero "g" testing via a water environment 

is needed. 

12. Static Test Hardware - $1.0M 

Schedule revisions and exchange hardware due dates require 

the qualification of LM/MM primary, tructure prior to delivery. 

Hence, a requirement exists to provide an additional Vehicle primary 

structure for the static tests. 

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SECTION II - MISSION PAYLOAD SYSTEM SEGMENT (MPSS) INTEGRATION 

(General Electric) 

Part 1 - Segment Description and Performance Requirements 

A. General 

The Mission Payload System Segment (MPSS)Lfs composed onthe 

photographic system and the subsystems necessary for its control and 

7Aese 

dynamics. [the subsystemQare located in the Laboratory Module and/or 

the Mission Module depending on the specific function concerned. 

General Electric is responsible for the overall integration of the MPSS 

and development ofEtheihardware with the exception of the camera and 

de.s re, So-.we clw 

the main optical system. This section Cdefine3 General Electric's 

aesponsibilities with regard to tha MPSS., eI/24'S Z,/ $40.5- 

*ot 

TheiSznzunic elementgof the MPSS are theresponsibility of 

General Electric and must perform the following functioZ 

1. Establish vehicle/target location relationship. 

2. Establish vehicle attitude reference. 

3. Acquire targets. 

4. Provides 2-degree-of-freedom motion to tracking mirror. 

5. Track targets with high precision. 

6. Provide commands to primary camera. 

1 

T. Control tracking mirror thermal environment. 

8. Regulate and distribute electrical power to the Mission 

Payload 

9. Provide Mission Payload Command & Control data handling. 

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B. Technical Description 

1. Aerospace Vehicle Equipment (AVE) Subsystems 

The Mission Module structureasjprovided by^Douglas is 

10 feet in diameterild 31.0 feet in length)and is fabricated in two 

sections. The forward.section, approximately 14 feet long, contains 

much of the mission related AVE provided by General Electric, such 

as the gimbal, mount and thermal control system for the tracking mirrors 

I -he A E2114 associated equiptenil The tracking mirror is provided by Eastman 

el s r, Ma, 4, lc. (WI PS) 1C 

Kodak. The Raftsection houses Eastman Kodak optical equipment, as shown in 

Figure II-1. 

The specific performance and description of the subsystems 

11P55 

related to the Mission Payloagldynamics are best presented in terms of 

the support they provide to the critical functions as depicted in the 

Subsystems/Critical Functions Matrix, Figure 11-2. 

a. Navigation and Control Subsystem 

The navigation and control subsystem generates the commands 

and provides the drive for pointing the optical system, sieving the 

primary optics tracking mirror, sieving the acquisition and tracking 

subsystem (ATS), and tracking targets with high precision. It depends 

on the Laboratory Vehicle attitude control and translation subsystem 

(ACTS) for vehicle attitude and stability; however it establishes the 

precise attitude references needed to generate the pointing commands 

for the tracking mirror, and precision tracking refinements after 

oetv e,,viettd 1.044;:lf 

target acquisition. This subsystem includes two star trackers to 

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provide precise orientation of the vehicle with respect to the 

celestial reference frame. A low-G accelerometer measures drag 

effects, permitting precise correction of the orbit ephemeris. 

ve4-2:4_1. LI:VV 

In addition, an image[Motionjsensor detects relative motion of the 

target across the optical format and provides an error signal to 

correct the rates of the tracking mirror drive. Three drive units 

are provided to allow two-axes control of the main tracking mirror, 

single-axis servo control of the tracking mirror viewport aperture, 

and two-axes servo control of the acquisition and tracking subsystem 

optics. The mechanism which drives the main optics tracking mirror 

must provide line of sight coverage of +40 degrees in roll and +30 

degrees in pitch with extre precision , and smoothness. The successful 

m44:0t7 

fabrication offlCer-Vit mirrors will allow -40 in pitch. The smearcontribution of the tracking mirror drive mechanism must not exceed 

per second. To meet this specification the roll gimbal 

motions must be capable of providing smooth action at angUlar rates 

less than +0.5 degrees/sec and must be accurate to +0.001 degrees/sec 

when the tracking mirror is rotated at a rate of .001.: degrees/sec. 

b. Structures and Thermal Control Subsystem 

The structures and thermal control subsystem consists of 

mountings, bearings and other hardware which comprise the gimbal 

structure ~la support and. Qmovement. of the tracking mirror, and 

the environmental control necessary to assure proper functioning of 

the tracking system and to minimize distortion of the tracking mirror. 

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The gimbal structuregp4provideS: 

1. Support for a 1200 pound tracking mirror and assembly 

during launch. 

2. Zi4ous ing for a control system capable of simultaneous 

pitch and roll motion. 

3. alitiffness (13.5 CPS) sufficient to maintain a limited 

margin between the natural frequency of the control system and the 

mounting structure. 

4. A capability of operating in a one "g" field and 

yet maintain alignment on orbit. 

1.4 

The gimbal Cwill bgfabricated mainly from beryllium and 

CsIribe one of the largest. structures ever made from this material. 

Beryllium was chosen primarily due to its high modulus of elasticity, #4s 

low density and ability to support the loads. 

The thermal control equipment includes necessary thermal 

coatings, hatches, blankets, louvres and heaters to maintain the 

tracking mirror bay at a temperature of 70 degrees F.+ 5 degrees. 

The temperature differential (face to back) of the tracking mirror (TM) 

must be no more than 0.1 degree throughout the photographic pass. 

.1.4 e f 

When the TM is exposed tordirect im ingmera of external heat fluxes 

during photographic periods)the thermal designausiminimizeithis 

disturbance by limiting the irradiation of the TM from areas outside the 

•fr4 44t$k • 

optical field of view. Totavist in accomplishing this the TM bay is 

provided with a viewport door, and a servo controlled viewport aperture 

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to aid in maintaining an acceptable average linear gradient throughput 

the tracking mirror thickness. 

C. Acquisition Subsystem 

The acquisition and tracking subsystem (ATS) enables man 

to make his most important contribution to the system. The quality and 

amount of reconnaissance data which can be acquired by an unmanned, 

completely automatic reconnaissance satellite is degraded by such 

things as target location errors, ephemeris errors, computation or 

Pi4v1, 

servo errors and weather obscuration of targets. [Crew memBeril using 

the ATS) can minimize these effects by exercising judgment in appraising 

targets for intelligence value, rejecting pre-programmed targets which 

are obscured by weather or-considered to be less valuable than other 

observed targets and driving the tracking mirror to more valuable 

alternate targets. Phe crewnemberacan also refine. the automatic 

pointing and image motion compensation systems, thereby enhancing the 

photographic quality. LWith the 2 acquisition scopes of the ATS,'the 

crew members can view the earth with magnification of from 15X to 30X 

and 63.5X to 127X with corresponding fields of view of from 4 degrees 

to 2 degrees and 1 degree to 0.5 degrees. The subsystem provides a 

ce.e.ti24 

maximum ground resolution

A 

 of about 3.6 feet for detailed target study 

64,4, 

in a 0.5 degree field of view at 127K magnification. Ohe tw43acquisition 

LAS 

scopeDoperates independently andlavela computer driven servo system 

with an automatic pointing accuracy of 0.1 degree. TWo scopes are 

provided to allow the crew members to inspect different areas 

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simultaneously thereby increasing the systems flexibility considerably. 

The limits of the tracking angles are +70 degrees to -45 degrees in 

pitch and + 45 degrees in roll. 

d. Consoles and Displays Subsystem 

The consoles and displays associated with the mission 

payload are physically located in the pressurized compartment of the 

Laboratory Module. TheyassentianZprovide the visual and mechanical 

man-machine links to allow the crew to monitor and control, as well as 

to identify out-of-tolerance conditions, in the Mission Payload. In 

general, the crew functiOns supported by this subsystem are related 

to ground station contacts, payload status and calibration checks,. 

payload pass preparation, the payload pass postpayload pass analysis 

and reporting, preventative maintenance, and data processing and return 

activities as required. Figure II-1, View A-A gives the general 

location of these consoles and display units. There are crew stations 

located at bays 2 and 8 which each. contain display and operating panels 

for the visual optics, acquisition scopes, cue displays, magnification 

controls, and teleprinters. Each crew station provides access to 

one acquisition scope and to visual optics viewing through the primary 

telescope. There are also miscellaneous indications of time-to-target, 

target priority, and other operational visual displays. The crew 

member at bay 8 has access to the window in bay 7 to enable him to 

view the horizon. Both crewmen have access to the computer keyboard 

located in bay one. When operation of the film processor and readout 

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(v4 provided) is desired the crew members may turn to bay 4 

and utilize the provided controls. 

e.. Communications, Command and Instrumentation Subsystem 

The heart of the communications, command and instrumentation 

subsystem for the Mission Payload is the Mission Data Adapter Unit (MEW 

which provides the interface for command and reaction between the Mission 

Payload and the Laboratory Vehicle data management subsystem. There 

are two redundant mission data adapter units wired with the two redundant 

airborne digital. computers in the Laboratory Vehicle data management 

subsystem to provide high reliability. The commands generated on the 

ground are fed toa mission data adapter unit through a laboratory data 

adapter unit and the airborne computer.. Cammands generated on board 

are provided to the mission data adapter unit through the airborne 

computer. The mission data adapter unit translates these commands into 

mission payload actions and provides feed back to the computer which 

in turn provides the verification functions to the Laboratory Module 

telemetry equipment for transmission to the ground either in real time 

or by delayed recording. 

f. Electrical Power and Signal Distribution Subsystem 

The electrical power and signal distribution subsystem 

conditions and distributes power and signals to meet all requirements 

of the Mission Payload. This subsystem is supplied power by the 

Laboratory Module electrical power subsystem. 

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-*software 

The Mission Payload subsystems software including the 

incorporation of the mission logic is the responsibility of GE. 

Before any final design can be accepted a great number of simulation 

studies using the crew membera must take place, to evaluate and 

measure the man-machine factors concerning command and control and 

on board software affecting decision. Software is also being 

developed for system checkout in conjunction with the Aerospace Ground 

Equipment (AGE) and for system simulation equipment. 

Simnlation Equipment 

A comprehensive simulation program is an essential 

element of the successful-development and operation of the Mission 

Payload. The overall objectives of the simulation program are: 

(1) to measure and optimize the man-machine-mission interface and 

(2) to provide mission training. The first Mission Payload simulator 

is called an Elemental Development Simulator and is Co bElused 

primarily for establishing measurements for proper systems design of 

both AVE and software.'ItErigprovidestest data in three basic 

categories: 

1.. Man-Machine (cueing philosophy, sensor controls 

and alignment, etc.); 

2. Machine-Mission .(targeting mix, sensor drive signal 

validation, etc); - and 

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3. Man-Mission (targeting mix, cloud and atmospheric 

effects, activity detection, target acquisition, timing, etc.). 

The Elemental Development Simulator is in operation at 

the General Electric plant, Valley Forge. 

Mission Payload slmlation equipment is[being]designed to 

perform four basic functions: 

1. facilitate training the flight crew in the complete 

mission; 

2. develop and practice in-flight maintenance procedures; 

3. facilitate MOL integration with the Satellite Control 

Facility (SCF); and 

4. develop operating and contingency plans. 

ItkIlDincludeshardware and/or software simulation of 

control/display interfaces for all major functions of the Mission 

Payload. 

2. Aerospace Ground Equipment (AGE) 

Aerospace Ground Equipment (AGE) Ein. be provided i.O] etc,07 cIeS 

support the development, test, and operational activity of the Mission 

Payload e mechanical AGE includes theEcomon type such s,Qhandling 

and support equipment Cut which is peculiar to the type, size and 

weight of AVE which it is required to support; mechanical checkout 

equipment to perform alignment of the optical elements to tolerances 

as close as 5 seconds of axe and transportation and environmental 

control equipmentlipiImust]limit shock and vibration induced to the 

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Mission Module during transportation and maintain the required 

temperature, pressure, humidity and cleanliness. e electrical 

AGE consists primarily of the Computer Integrated Test Equipment (CITE) 

which is a programmable checkout system. The CITE can operate in 

a manual or semi-automatic mode via a stored computer „program with 

t3 

a minimum of operator assistance. This equipment will bused &oJ 0, /"-4 

0:- ,04 

Stumal]electrically mated system compatibility tests, development 

tests, qualification tests and performance demonstration including 

mission profile test. CITE(illilperfonn3these tests on all the 

Mission Payload System Segment subsystems and on the Photographic 

System subsystems. This checkout system will be deployed at the 

P E e tot,*if 

General Electric tasetozw. the Douglas fee ery and VAFB. A modified 

d f c.ri‘ 

configuration will be deployed at Eastman Kodak Company to support 

Davt'vie 

the Photographic System performance tests. At tire Douglas gaatow 

1-1 Down 71,v5(0. 

and at VAFB, the CITE WOISinterfaceSwith the Betnglas-easeaseetweetu4p-! 

issukt All Systems Test Equipment Group (ASTEG) to support the integrated 

systems testing performed at these locations. The CITE vial provides 

for functionally identical tests at all locations and 1.411 provide3 

for direct correlation of data throughout all levels of MPSS testing. 

This equipment %Via permitSmalfunction detection and isolation within 

• t4_1b the AVE to a replaceable black box level and loft& be capable of selfverification and. fault localization. During all testing the CITE0 1-i 

-timmorttre AGE software-0 4,01 permittmodification of theAest procedure 

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to correct errors in the procedure or to retest a subsystem or 

component using a slightly modified procedure. 

5. Data Return Vehicle (DRV) 

1r;1, 417.e automatic configuration of the orbiting vehicle)f-4 

's 1,e e Inc eel 64; A -r-v1 fl, is pre de, 

reloimeems1:61.4.-B—with a Support Module. Erle—sleetibthlalse 

Data Return Vehicles {DRY‘14.-±n—serfei• ) 44 ejected from the Support 

/-1 

Module to return the exposed film. The film is transported and loaded 

into the DRV's bit equipment supplied by Eastman,Kodak. The DRV's 

qr.t 

wil.%.441e similar to the General Electric Mark V model, modified as 

necessary, to return approximately 60 potnds of film each. The 

number of DRVs per orbiting vehicle 4on the final configuration 111 4 4 

of the Support Module 

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Figu

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Part 2 - Operating Concept 

After injection of the Orbiting Vehicle into orbit, the 

protective hatch covering the tracking mirror viewport mItt-bei4 

ejected. The tracking mirror bay viewport thereafter -apt covered 

doer 

by sliding mask except during photography. This mask is slaved to 

the tracking mirror pitch drive to assure that the bay is.amipopen 

43 4 

the amount necessary for photography, thereb5% aidm, in thermal 

TM 

control of the bay. 

An on-board computer, periodicAlly updated with target 

coordinates and ephemeris data from the tracking stations, aillkSera.^4-4- ...j 

programspointing angles and tracking rates for the acquisition 

scopes and the primary tracking mirror.iadeperidesibilyft In the manned/ 

automatic mode the crew may override or refine this automatic 

tracking by use of console mounted controls. Visual target cues will 

be provided to aid in recognition of targets. Crew members may 

correct target centering, provide vernier control to tracking rates, 

determine target visibility

/

and select active or secondary targets.- 

tr; 14.1 6:te ; re / lei;e is4.A.:‘, llrs - e-- 

itimibise.. readout option ke-ilimerpokeefi-ikr-ftilghts°4•weadiletT-tee- 

\)Ri, 

..-.4“;Vgrew may select and edit photography to be transmitted to a ddimp,<,eyw....q7 a Vv 

readout station. During station contact the selected photographs441114,00/01" 

be automatically scanned and. transmitted by the readout Erten.. Nit tile .,47 1-4 

0.s re'S Nt-44 p e e, c40. 1 45' Qom DRV the first 

A 

returned with -the remainder of the photographic take in data 

recovery capsules in the Gemini. 

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+t) • 

week's primary recordadia.be returned by DRV. All other film is 

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/ In the automatic configuration the film wiliwir.loaded into 

the Data Recovery Vehicle by means of Eastman Kodak film chutes, 

cutters and sealers, the Data Recovery Vehicle re-entry sequence 

is 

iMilipmbe initiated by ground or stored Command at the appropriate times. 

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Part 3 - Schedule Considerations • 

, ;16 1:....,ey,e), a: 

General Electric ie-Abbe integrating contractor for the 

Mission Payload System Segment)am* requires close

‘, ,

coordination 

kfC- DAC. e 1cNe.: 

with both EKC and 4461. 5111e40\criticalareas of thermal protection 

for the Mission Module Forward Section and precision mounting 

A 

bearings and servo mechanism for the large tracking mirror miniaborr14/4/e 

y;re detailed design, breadboarding, analysis, and development effort. 

GE is also responsible for an extensive mission payload simulation 

and training program. 

An image velocity senso his. reqUired in the automatic and 

manned/automatic mode to provide image motion tracking to less than 

0.1% of the velocity/range-ratioin order to. provide the extremely 

high resolution photography desired. GE 

integration of the Imag-ualoi3Opmeoaser 

sv;c0.421,, ,, 

is responsible for the 

S e eifisAn 

„s 

-11`4e 

into the vehicle, and' sensor 

. . .; .. . vi • 

The total GE schedule for design, development, fabrication, 

and test is shown in Figure 11-3. The GE effort is a continuing 

interaction between the systems and design engineering requirements 

and the implementation process. Cost effectiveness trade-off studies, 

degraded mission mode analysis, and systems effectiveness determination 

is 

are to be pursued throughout the entire program. There 144.6.1-4e a continuous interchange Of information among the USAF, GE and other 

associate contractors to insure systems optimization. System 

engineering tasks during the design, engineering phase are: 

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Q. establishment of systems, subsystems and environmental 

design requirement specifications; 

hi. technical and mathematical analysis to evaluate and 

optimize selected configurations (stress analysis, thermal analysis, 

reliability analysis, etc.); 

development and evaluation of subsystems and systems 

test specifications and procedures to insure that simulated tests, 

determine capability of hardware to meet the design criteria; 

EL management of the interface 'to insure overall system 

continuity. 

Concurrent with the above effort is the development, 

installation and checkout of the software for command and control 

of the payload operation on orbit. Detailed software planning must 

also be generated so that subsequent data reduction and evaluation 

of flight data may be accomplished in a timely and efficient manner. 

The systems cross checking and feedback loop of the aforeMentioned 

effort is accomplished on the ground priorto flight) through simulation 

and training programs which establish optimum relationships between 

flight crew, hardware and software. 

The test flow at GE is shown in Figure II-4. The Mission 

Module Forward Section engineering model, (non-prime is used for 

payload component and system tests before assembly and test of the 

first flight. The mission Module Forward Section test article, 

1F, is used -for thermal testing at GE; upon completion of these tests 

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a • 

it is transferred to EKC to provide the fOrward interface for thermal 

testing of the Mission Module Aft Section, and is then returned to 

GE for the use as their engineering model. A second test article, 2F, 

is used for dynamic structure tests of the tracking mirror support 

structure at GE. It is then transferred to EKC and mated with an 

aft section for Flight Vehicle 1. A third test article, 3F, is 

used for development testing of the Forward Section AVE, including 

extended thermal/vacuum operating tests of the gimbal and environmental door mechanisms. A fourth test article, 5F, is used for 

qualification testing to verify all.manufacturing and test procedures, 

and to qualify the Mission Module Forward Section in ambient, vibration, 

and thermal/vacuum tests and to conduct electromagnetic interference 

tests. Upon completion of qualification tests this Forward Section 

is transferred to EKC for the Mission Module Qualification Tests. 

In addition, two other Forward Sections are assembled, checked out 

and shipped to EKC to mate with Aft Sections for EKC dynamic testing 

and development testing, - 2FX for dynamic testing and 4F for 

development testing. 

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Part 4 - Funding Requirements 

A. General 

Costs are based upon detailed Systems Office estimates and 

the Phase II contract negotiations, and include an estimate of the 

schedule adjustment costs and deferredA ameints. 4163eferrals are 

described in paragraph B, and identified changes in paragraph C. 

Funding requirements are shown by major work breakdown structure 

in Figure 11-5. The tasks within the major breakdown structure are 

as follows: 

1. AVE Hardware 

44, 

Thle-prevottlee-fer design and development of all airborne 

mission payloadAequipment.- ibis equipment consists of tracking mirror 

drive and control mechanisml target acquisition and tracking system,' 

navigation and control, data return vehicles, target cueing system 

and environmental control system for the tracking mirror bay. 

e .I,d die e 

flight hardware for three manned and two linmsmned flights. 

2. AGE 

Wies:rncludes design, development, testing and qualification of all AGE to support the Mission Mbdule, including the Mission 

Module transporter and the environmental control and monitor unit for 

use with the transporter. Also included is electrical AGE equipment 

Fv r- to-be-004-aterbing:And.cnackout Of the 'Mission Module:' The AGE 

1.1 

m41.1-be installed and operated at General Electric and at other 

associate contractor facilities. The set used at General Electric 

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will be refurbished and installed at VAFB. All electrical AGE 

s 

will of similar design and function. However„ . AGE will be , 4 

TI $ f erf e tev ,e;■•••.,d-,15 

adqpted as required 'each specific location. Electrical AGE 

will be provided to EKC for mission module level testing. 

3. Training 

40dm. Amides for the development, testing and operation 

of all training equipment associated with the mission payload. It 

includes the operation of mission payload, simulating and training 

of astronauts at General Electric Company and at VAFB. Mission 

Payload simulation equipment is integrated with equipment from other 

associate contractors to make up the complete MOL simulatormWrilfkr 

viiiimirs. installed at VAFB., 

4. System Test 

gbte-tesk covers contractor performance of General 

Electric AVE in-house deVelopment and testing at the system level. 

4,0mp.1; includes General Electric Company's participation in system 

tests at other associate contractor facilities and overall MOL system 

test at the Laboratory Vehicle contractor's facility and VAFB. This 

task includes providing the hardware and support required to accomplsih 

all system testing of General Electric furnished equipment. It also 

includes General Electric hardware and support required by other 

associate contractors for their system development and qualification 

tests. 

--T 

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5. System Engineering and. Management 

This task includes all engineering analysis and design 

integration at the mission payload system segment level. System 

engineering and integration is accomplished on both AVE and. AGE 

hardware and software. Management under this task includes project 

engineering and planning utilizing PERT and providing a Cost Planning 

and Control System. Also included is a system effectiveness program 

to insure that overall system effectiveness requirements are met. 

Travel, living and training expenses of contractor personnel are 

included as a direct charge under this task. 

6. Documentation 

This task includes preparation and publication of all 

data and reports required of GE. 

7. Launch and Flight Operations 

Launch operations cover the GE support and services 

required at VAFB to test launch GE flight hardware and to support 

other launch activities required at VAFB. Flight operations cover 

GE software development including computer programs for caaMand and 

control from the Mission Control Center, on-board computer program 

for operation of target acquisition, tracking and functioning of all 

mission payload equipment. This item includes the GE support and 

services at the Satellite Test Center to provide mission operation 

support and flight data analysis and reports. The task also includes 

providing a capability for an off-line data reduction facility at 

Sunnyvale. 

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B. Deferral Status 

Several tasks were deferred from the Phase II baseline 

contract pending further definitization. The estimated costs and 

a brief description, of these items follow: 

1. Field Support Program - $43.3M 

This effort is in support of GE-AVE/AGE and of Exchange 

Hardware at DAC-Huntington Beach, VAFB STC-Sunnyvale, and On-Orbit/ 

Post-Flight Support. 

2. Image Velocity Sensor - $9.0M 

The IVS detects relative motion of the primary optical 

image across the optical format and provides an error signal used to 

correct the rates of. the tracking mirror drive. GE.is revising the 

IVS specification in preparation for participation in an AF and potential 

subcontractors specification review. 

3. Acquisition Tracking System - $27.2M 

The ATS is the target selection and acquisition system. 

The subcontractor, ITEK, has been selected and has been working since 

January 1967. 

4. Data Recovery Vehicle System - $13.8M 

The DRVS will provide for the return of the photographic 

record to the recovery area. 

5. Exchange Hardware - $19.1M 

This deferred effort covers those items of development/ 

test exchange hardware which GE will provide to DAC and EK. 

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6. Logistic Spares - $9.5M 

Logistic spare components/hardware will be used to 

support the field site operations and the support of GE hardware. 

7. Support Module - $3.1M 

C. Identified Program Changes 

Estimated costs and brief descriptions of each program 

change considered essential to program accomplishment are as follows: 

1. Sliding Mask Door (SMD) - $6.4m 

The sliding mask door, which replaces the baseline thermal 

door, will open to expose the tracking mirror - viewport.,, during target 

acquisition, and remain closed at othertimei to control the tracking 

mirror bay thermal environment. The SMD concept is currently under 

study. 

2. Flights 1 and 2 Support - $.9M 

In support of Development Flights #1 and 42, GE will 

define and integrate the GE/EK flight objectives and sensor/data 

requirements. This will include: equipment substitutes to simulate 

payload flight conditions, data analysis for GE equipment, and post 

flight report. 

3. Alignment Equipment - $LAM 

Alignment equipment and alignment operation will be 

in support of GE equipment at the field sites. Alignment will include 

alignment of low "gffaccelerometer, gimbal to COA, deflection 

compensators, etc. 

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4. Refurbishment, Qualification Vehicle Consoles - $.3M . 

The qualification vehicle consoles will be refurbished 

and retested for FV-5 use. 

5. Engineering Model - $2.6M 

Refurbishment of Engineering Model 103 for use as an EKC 

development test vehicle. 

6. Mission Module Test Set Installation and Checkout, 

Qualification Vehicle Test - $.1M 

The revised test flow requires that the NM Test Set be 

installed and checked out with the qualification vehicle prior to 

its delivery to EK. This is required to assure compatibility of the 

test equipment with the vehicle prior to delivery since MS will be 

used at EK. 

7. Power Interrupt - $1.5M 

GE will evaluate and institute the power interrupt 

capability, which requires that each component and subsystem, as well 

as system, be provided with protection to withstand a power interrupt 

and proper initialization sequencing so as to be able to meet all 

system performance requirements within 30 seconds. This requirement 

is presently being evaluated for design methods to minimize schedule 

and cost impact. 

8. Command Software - $1.3M 

This deferred effort is the design, development and 

integration of the Executive Routines GE is currently working on the 

proposal. 

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SECTION III —PHOTOGRAPHIC SYSTEM 

(Eastman Kodak Company) 

Part 1 - System Description and. Performance Requirements 

A. General 

The Photographic System contains optical, mechanical, thermal, 

electrical, and structural components necessary to perform high quality, 

stereo, photographic reconnaissance. It will provide a static lensfilm resolution equal to or greater than 114 lines per millimeter at 

the film plane at a 2:1 apparent scene contrast and illumination of 

890-foot-Lamberts when using film of type 3404 (but with an exposure 

index of 6.0). The Photographic System will be capable of both northbound and southbound photography from orbital altitudes of 70 to 230 N.M. 

It will be capable of both manned/automatic and automatic operation and 

provides for visual observation through the primary optics at high 

magnifications in the manned mode. 

The baseline Photographic System for the MOL consists of 

a primary lens system with necessary support structure, provisions 

for camera recording of optical image or visual observation with 

variable magnification, control console components, film viewing and 

processing equipment, necessary electronic and thermal control equipment, 

and ground support equipment. 

B. Technical Description 

The primary lens system is a Ross telephoto design with a 

SECRET 

focal length 70-inch primary 

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field of view. A lens barrel provides optical alignment, controls 

thermal environment, and provides the mounting structure for the 

70-inch reflecting aspheric mirror, the Newtonian and Rosa (45-degree) 

folding mirrors, and the Ross corrector lens assembly. 

A 70-inch tracking mirror provides for nulling the ground 

image motion and for directing the optical line-ofsight obliquely up 

to 40 degrees to either side and approximately 30 degrees fore and aft. 

The camera, located in the Laboratory Module, is a frame type. 

The platen is equipped with across-the-format image motion compensation 

to decrease resolution degradation at the edge of the frame due to 

local relative motion during exposure. The image is circular, 9.4 

inches in diameter, with attitude; sequence, exposure time, tracking 

mirror point angle and camera settings recorded within the format of 

each frame. Variable exposure times are possible with the focal plane 

shutter at a recycle rate of one frame/second. In the manned/automatic 

configuration a capacity for 10,250 feet of type 3404 black and white 

primary film is to be incorporated with additional provisions for 

2900 feet of secondary film consisting of color, infrared or other 

special film types which are employed in a secondary camera. In the 

automatic configuration 17,400 feet of type 3404 black and white film 

is provided. The automatic version has necessary film transports 

(chutes), and cutters to load the re-;entry vehicles automatically. 

In-flight processing equipment is provided by EKC for the Laboratory 

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Module to allow the crew members to prepare film from the secondary 

camera platen for photo readout to the ground, when this option is 

provided. The an-board processor is capable of processing 50 pictures 

per day at a rate of less than .5 minutes per picture. 

The primary optical system can be used by the crew members 

for high magnification (127K visual viewing. An eyepiece 

is provided for this capability with an apparent field view of 40 

degrees and a full field angle on the ground of 0.4 to 0.04 degrees. 

Approximately five per cent of the total image illumination is utilized 

for this purpose. 

An environmental control subsystem utilizing passive techniques 

to the maximum extent is provided for proper temperature control of the 

Photographic System during ground conditioning, powered flight, and 

on-orbit. 

Special optical and mechanical alignment equipment for the 

Photographic System will be provided to validate optical and mechanical 

dimensions of the sensor, to align the navigation and control reference 

with the optical sensor, and to align automatic film handling equipment 

as necessary. Handling and transportation equipment will be provided 

to implement Photographic System vertical and horizontal displacement, 

erection to a vertical attitude, and in-plant transport in a horizontal 

attitude. 

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Part 2 - Operating Concept 

The Photographic System incorporates automatic focus control 

and automatic alignment. The exposure will be controlled by computer 

command as will the pointing and tracking of the tracking mirror. In 

the manned/automatic mode these functions can be overridden or refined 

by the crew. 

The automatic mode will provide film loading cutting and take-up 

with the film feeding through chutes into the Data Recovery Vehicles. 

In the manned/automatic mode the crew will perform these functions 

and load the film into the Data Return Vehicle (DEV) when provided, 

and into the Data Recovery Capsules (DRC), for return in the Gemini B. 

In addition, the crewwill,load the secondary camera platen and process 

film for system health checks. The crew may determine the quality • 

of the photographic take and recommend or initiate changes in the 

optical system (focus, exposure, and alignment); the photographs may 

be read out to the ground when this option is provided. 

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Part 3 - Schedule Considerations 

The fabrication of the large optical surfaces required for the 

MOL Photographic System is a most exacting and time-consuming process. 

The procurement/production cycle of the large DORIAN fused silica raw 

blank is approximately one year. The initial order for MOL, which 

includes masters, development models and flight models, was placed 

in October 1965. The first blanks have been delivered and are in the 

process of being ground and polished. The grinding and polishing of 

the mirror surfaces is a laborious, artful process requiring repeated 

machine operations alternated with optical testing using interferometer 

techniques in an evacuated test chamber. Each iteration of this process 

refines the surface definition, and the final DORIAN'requirements 

• (0.05 wave length for the tracking mirror and 0.1 wave length for the 

primary mirror), are achieved by selective electroplating techniques. 

The figure of the primary and tracking mirrors requires repeated 

comparisons with parabolic, spherical and piano masters during the 

grinding and polishing operation. Two sets of these masters have 

been ordered and must be ground and polished prior to flight model 

primary mirror tests. Frame mounting of the mirrors must eliminate all 

strains which could cause image distortions, and this process dictates 

extensive, time consuming testing. The primary and tracking mirrors 

will distort;their surface figure in most attitudes in a one-g field 

and, therefore, cannot be tested on edge in a horizontal facility. 

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They must therefore be provided with an air-bag type of support in 

a vertical facility to hold the on-orbit surface figure. 

Production facilities and equipment for development, fabrication 

and testing of the optical system of EKC were approved in April 1966. 

Facility construction and equipment procurement and installation will 

not be complete until approximately December 1967. Present EKC 

facilities are scheduled for use in early payload equipment design and 

fabrication but are not adequate to handle the extensive development 

and flight model fabrication, assembly, and test program required. 

Engineering design was started early in the contract definition 

phase and will progres$ through mid-CY 1967 to support the development 

and production of the various test and flight models required. A 

.brief description of the test models and their use follows: 

a. Formula Sample - Asphere Ross Corrector elements are 

ground to the optical 'mathematical formula and physically tested with 

the primary mirror to confirm the accuracy of the optical system 

tolerance calculations and to develop testing and adjustment techniques. 

b. Structural Development Model - Duplicates mass, inertia 

and structural rigidity of the actual flight model. Early tests of this 

structure uncover resonance problems, structural magnification factors 

and produce data for component design requirements. This model will 

be shipped to DAC for structural testing. 

c. Structural Model (Static) - Will duplicate the CameraOptical Assembly (COA).in strength and rigidity.and be used to evaluate 

effects of acceleration loading. 

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d. Thermal Model (2A) - Duplicate the mass, thermal time 

constants of the structure, insulation effectiveness and surface 

radiation characteristics of the flight hardware. The model will 

be integrated with the GE thermal test model for thermal testing. 

e. EngineeringlMod 1 (3A) - Used to develop flight-type 

hardware by functional test of components and systems before 

assembly and test of first flight model. It is also used to test 

design changes as the program progresses. 

f. Qualification Model (4A) - Demonstrates capability of 

the system to meet performance requirements in a simulated space 

environment. 

The EKC test plan is shown in Figure III-1 and the schedule 

for the attainment of major design, test, development and production 

functions is contained in Figure 111-2. 

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Part 4 - Funding Requirements 

A. General 

A CPFF contract has been signed with Eastman Kodak for the 

development and manufacture of five camera payloads for the MCL. The 

contract includes construction of two buildings, procurement of required 

industrial, test, and support equipment as well as the hardware for 

the sensors. A brief description of the effort contracted for follows: 

1. Special Industrial Requirements (SIR) 

SIR is the designation for non-fee bearing equipment EKC 

is authorized to procure under terms of the DORIAN contracts. The 

major facilities in this listing are the Camera-Optical Assembly Test, 

Primary Mirror Test, Tracking Mirror Test and VariouaThermal Chambers. 

2. Research and Engineering 

a. Aerospace Support Equipment 

This line item is for ground and test equipment required 

for development of the sensor. This is not a fixed list of equipment 

but based on need only. It includes such items as solid test mirror 

blanks, camera assembly, test equipment, test consoles, unevacuated 

test towers and chambers, handling equipment, etc. 

b. Aerospace Flight Equipment , 

This line item is the estimated cost for development 

and manufacture of five camera optical assemblies. 

c. Other 

Other costs include Vality control, reliability, 

manuals, specifications, reports, etc. 

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3. Construction 

The EKC contract authorizes construction of two buildings 

for the development, manufacture and test of the DORIAN sensor. The 

line item also includes installed equipment such as cranes. 

4. System Test Operations 

EKC has been directed to conduct static and dynamic 

testing which involves the Mission Module structure, the GE tracking 

mirror drive and other equipment installed in the Mission Module 

structure by GE and the Camera-Optical Assembly as a complete unit. 

Acoustic testing will also be accomplished by EKC. In addition, EKC 

will support field tests at other installations. These costs are 

additive to the basic contract. 

5. Special DORIAN Efforts 

Contractual arrangements are being made to grind and 

polish the two experimental 72-inch ceramic vitreous (Cer-Vit) mirrors. 

EKC will not do this work but, will be available to consult on the 

putlity of the mirrors and their incorporatiod into the DORIAN design. 

Other work includes simulation, image velocity sensor, and special 

advanced studies and technology. 

Some contract changes have been issued. Costs for 

the payload segment testing, acoustic testing, and support of field 

test are included in a deferred package yet to be negotiated with 

EKC. The deferrals are described in paragraph B and new work in 

paragraph C. photographic System costs are shown in Figure 111-3. 

DORIAN 

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B. Deferral Status 

Several tasks were deferred in the original contract 

negotiations. Deferrals are included in the Photographic System total 

costs shown in Figure 111-3. 

1. Acoustic and Mission Payload Segment and Field Testing - 

46.8m 

Deferred pending completion of test definition studies. 

2. Facilities and Equipment - $5.OM 

Deferred pending completion of acoustic test study and 

decision on nitrogen reliquification plant. economics. 

C. Identified Program Changes 

1. Mirror Mounts - $6.om 

Parallel design effort of tracking mirror assemblies for. 

10-inch thick Cer-Vit and baseline 12-inch thick fused silica mirrors 

and parallel effort to design, fabricate and test a mount for a 10-inch 

thick primary Cer-Vit mirror. 

2. Data Recovery Capsules (DRC) - $1.0M 

Effort to finalize the design of pritery and secondary 

DRC's compatible with the structural and environmental undefined interface constraints of the Gemini vehicle and the untested operational 

handling requirements. 

3. Support Module - $4.3M 

Pending definitization of program requirements, a minimum 

sustaining effort is continuing on film handling hardware for the six 

Data Return Vehicle configuration for unmanned flights numbers 6 and 7. 

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4. Thermal Slats - $2.0M 

Provide tracking mirror thermal slats for appropriate 

test models and all flight models which will optimize the thermal and 

optical effects produced when the sliding mask is operated during the 

photographic passes. 

5. Conformal Coating - $0.5M 

Insulate all conducting surfaces not at chassis potential 

to enhance electrical circuit reliability. 

6. Sealed Laboratory Module Systems - $5.2M 

Seal Laboratory Module components to the extent practical 

as a means of controlling- contamination, fire, and explosive hazards. 

7. Mass Substitute- for Flights I and 2 - $1.5M 

The plan to fly structural development hardware on flights 

1 and 2 resulted in a requirement for additional mass substitute components not previously in the program. 

8. Overrun - $19.1M 

Overruns have been identified in the facility and Itek 

contracts. 

9. Increase in Scope of Deferred Items - $7.6M 

Subsequent to the compilation of the deferred list the 

requirement for increased scope in areas such as security, configuration 

engineering, interface definition, safety, and subcontract management 

were determined. 

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10. Current CCN's -16.6M 

New requirements have been recognized for the following 

items and are reflected in current CCN's: 

a. Acoustic Testing Study 

b. Support Module Study 

c. DRC Mock-Ups 

d. Addition to Building 101 

e. Power Switching 

f. Mission Development Simulator (MDS), Simulators 

g. Increased Cassett Film Capacity 

11. Modal Survey == $1.4M 

This effort will determine natural modes, shapes, structural 

"damping characteristics and coefficients to verify the dynamic analysis 

predictions. 

12. Environmental Testing - $5.5M. 

Additional and redefined effort to test Mission Module 

segments in the Acoustic Test Facility. 

13. Other - $2.8M 

The following list includes the remainder of the identified 

- I 

deferrals, all of which require further definition prior to being negotiated: 

a. Mission Module Simulation Equipment (MMSE) hardware 

and software inputs for the Mission Module Simulator. 

b. MMBE support at GE, Valley Forge. 

c. Crew training and modified rehearsal schedule. 

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d. Mission simulator support at VAFB. 

e. Systems Effectiveness. 

f. In-house software. 

g. Schedule Interface Log (SIL). 

h. Training and operations of associate contractor test 

equipment. 

i. Acquire Mission Module Forward Section test capability. 

• 

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SECTION IV - GEMINI B 

(McDonnell Aircraft Company) 

Part 1 - Segment Description and Performance Requirements 

A. General 

The Gemini B is a modified NASA Gemini Vehicle. Its primary 

function is to transport the flight crew during launch and to return 

crew and all mission data to a water landing recovery. The MOL requirements that have the largest impact on the Gemini B design are: launch by 

a Titan IIIM into a polar orbit; dormant storage on orbit for 30 days; 

autonomous orbit operations (loiter) for 14 hours; re-entry from 80 to 

100 degrees inclination orbits; and computability with the ground tracking and control network. 

Certain, modifications to the NASA Gemini Vehicle will be made 

to meet the MOL requirements identified above. These modifications can 

be grouped into two basic categories: 

1. Modifications for ascent/abort crew safety requirements, 

and 

2. Modifications to permit use as a segment of the Orbiting 

Vehicle rather than as an autonomous spacecraft. 

B. Technical Description 

1. The Titan IIIM Launch Vehicle imposes a more severe ascent 

abort environment that the NASA Gemini Launch Vehicle, primarily due to 

the operation of solid rocket motors at liftoff. However, addition of 

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guidance and control redundancy and minor changes to the malfUnction 

detection system (MDS) in the Titan IIIM plus two more retrograde 

rocket motors (required for polar re-entry) permits use of the basic 

NASA Gemini escape system. Changes in the escape system are primarily 

in its operational utilization. For escape on-pad, through burn-out 

of the solid rocket motors, the six retrograde motors will be employed 

to separate the Gemini and gain altitude. For on-pad and just after 

liftoff, the ejection seats must be used. The ejection seats will be 

used for escape from the Gemini only if land landing is imminent, if 

insufficient altitude will be attained to permit the Gemini parachute 

recovery system to deploy, or if the parachute recovery system malfunctions. 

This permits simplication of the NASA Gemini ejection seats by 

'elimination of the sustainer rocket and reduction of the catapult impulse. 

For escape beyond solid rocket motor burn-out, use of the 

separation motors or partial use of the retrograde motors will be made 

to separate the Gemini, followed by turnaround and re-entry with normal 

water recovery. 

2. The most significant modifications result fram the use of 

Gemini B on orbit as part of the overall Orbiting Vehicle. During the 

relatively short ascent, loiter and re-entry phases, life support and 

active temperature control are provided by the Gemini environmental 

control system (ECS). 'During the longer period of Gemini, dormant storage 

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on orbit, electrical power and minimum pressure and heat are supplied 

from the Laboratory Module Subsystem. Batteries in the Gemini are 

available for necessary electrical power during loiter and re-entry. 

Thus, the fuel cells, orbit attitude and maneuvering propulsion system 

(CAMS) and radiator cooling system are eliminated from the NASA Gemini 

version, resulting in simplification of the Gemini adapter. 

3. A means of crew transfer between the Gemini B and the 

Laboratory Vehicle is provided. The primary mode is through an internal 

pressurized tunnel. Emergency transfer is by the same tunnel in an unpressurized state or by means of extra-vehicular travel through the 

Gemini exterior hatches and the DRV hatch in the LM. The tunnel 

requires a hatch in the Gemini B heat shield and main pressure bulkhead. 

Initial validation of this design was accomplished with the Heat Shield 

Test (HST) flight in November 1966. Final validation will be 

accomplished in the unmanned Gemini B Qualification flight on flight #1. 

4. Crew and ground monitoring via the Laboratory Vehicle 

telemetry of Gemini B status during the on-orbit period requires the 

addition of an instrumentation interface between the Laboratory Vehicle 

and Gemini B. There are other selected systems requiring additional 

sensors. 

1 5. The . SP/DR: weight for the Gemini B is 6,120 pounds which 

includes crew member weight and GFE. 

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6. The Aerospace Ground Equipment (AGE) philosophy and hardware used on the NASA Gemini has been adopted for use with Gemini B. 

This system is manumily operated and will interface with other MOL 

segment AGE principally by procedural and test conductor contacts. 

Two sets of AGE are required--one at the contractor's factory and one 

at the launch site. The AGE previously used for NASA Gemini at the 

contractor's facility will be utilized to the maximum extent; however, 

certain revisions must be made and simplifications incorporated as a 

result of subsystem changes outlined above. The present operational 

concept requires that the launch site van-mounted AGE items be removed 

and hard-mounted in the Launch Control Center at Vandenberg AFB. 

7. System trainers and a procedures simulator are required to 

'support MOL operations. These items of hardware are to be refurbished 

units from the NASA Gemini program. 

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Part 2 - Operating Concept 

The Gemini B Vehicle will house the crew during ascent and re-entry. 

On orbit, it will be stored in a dormant condition ready for return to 

earth of the crew and data upon completion of the scheduled mission or 

for immediate occupancy and subsequent return of the crew in the event 

of an emergency abort of the mission. 

The ascent and re-entry phases are being patterned around NASA's 

flight experience. The crewmembers are carried into orbit aboard the 

Gemini. After booster shut-down, attitude and translational adjustments 

will be made from the Gemini through interface control of the Laboratory 

Module Attitude Control and Translation Subsystem (ACTS). The necessary 

subsystems in the Laboratory Vehicle will be brought to operational 

.status from the Gemini cockpit, and normal operation verified. The 

crewmen then transfer into the Laboratory Module through the tunnel 

and hatch installed in the Gemini heat shield. Laboratory Module and 

Gemini B atmospheres are stabilized and the Laboratory Module hatch 

is then closed for the on-orbit mission duration or until there is .a 

need to return to the Gemini. 

Provisions are included for backup emergency return to the Gemini 

by exiting through the Laboratory Module DRV hatch and entering through 

the Gemini exterior hatches. During dormant storage on orbit the 

Gemini atmosphere will be maintained at 0.1 psis and Gemini subsystems 

will remain in a powered down state for the mission duration. The 

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Laboratory Module oxygen system is sized to provide three repressurizations 

of the Gemini B while on orbit without reducing mission orbital duration. 

Upon completion of the mission, the crewmen return to the Gemini B 

by way of the transfer tunnel and activate the Gemini B subsystem for 

re-entry. Two small solid rocket motors are used for separation from 

the Laboratory Vehicle.. Six retrograde solid rocket motors are fired 

for de-orbiting. 

The Gemini B uses the same re-entry techniques as were used with 

NASA Gemini including the drogue and descent parachute. Water recovery 

by surface units as used by NASA in their Gemini flights is also planned. 

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Part 3 - Schedule Considerations 

Principal Gemini B development testing is grouped into four 

categories: wind tunnel, structural, systems and thermal-vacuum. 

Wind tunnel testing extends through early FY 68 to validate 

analysis of Gemini B aerodynamic,.thermodynamic, and structural changes 

from the NASA Gemini. Testing will be performed in the McDonnell and 

AFDC facilities for thermodynamic data, and in the Ames transonic and 

supersonic facility for structural dynamics investigations. 

Structural tests required as a result of the new Gemini B adapter 

design and the crew transfer modifications were initiated in FY 67 with 

refurbished and newly fabricated test articles. Testing will extend 

into FY 70 on adapter and re-entry module test articles. 

The dynamic conditions induced by the Titan IIIM booster during 

ascent will require early qualification by vibration testing of the 

adapter section, and components installed therein must be made in preparation for qualification. 

Six solid rocket motors are required for retrofire prior to reentry. The higher total thrust as compared to that used on NASA Gemini 

(four motors) is required to satisfy the higher retrograde energy requirement for re-entry from polar orbit. Salvo firing of the six retro-grade 

motors will also be employed to effect pad and early launch abort in the 

event of catastropic failure of the boosters. Static loading tests will 

be conducted to verify structural integrity of mounting proVisions for 

• - 

the motors. 

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Design adequacy, functional parameters, and operational characteristics 

of various Gemini B subsystems will be determined and evaluated during 

the period through mid-FY 69, initially as a part of design verification 

development testing, and later in support of component and system 

qualification. Electronic systems test and instrumentation mock-up units 

will be fabricated for this purpose. 

The requirement to activate Gemini B subsystems prior to launch 

and then shut down during the quiescent 30-day period on orbit represents 

a significant departure from NASA's experience on the Gemini program. 

On-orbit storage and thermal vacuum tests will be conducted in an 

environmental chamber to verify adequacy of design under simulated 

atmospheric and thermal conditions, as part of the Gemini B qualification. 

These tests will utilize the refurbished NASA Gemini 3A spacecraft with 

the new adapter design. Test articles must be available by the end of 

FY 68 to allow completion of the thermal-vacuum test series in time for 

delivery of the GBQ spacecraft. 

The Gemini B development schedule is shown in Figure IV-1. The 

level of AVE engineering effort required for FY 68 is approximately 

double that needed' in FY 67 and is necessary to provide complete design 

release for the FY 69 principal engineering effort in support of 

manufacturing and test. In addition to the early engineering effort 

required to support AVE and AGE development, substantial engineering will 

be necessary during FY 68 to develop and/or modify mock-up vehicles, staticDORIAN Page 90 of / .34; pages 

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test articles, development test assemblies and fixtures, and Laboratory 

Vehicle interface specirens for Douglas Aircraft Company, all of which 

are required early LA Om davalopeont 

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CALENDAR YEAR] 

MONTHS 

PROGRAM FLIGHTSCHEDULE 

GEMINI B SPACECRAFT 

PrelimDesignReview(

AVE) 4 

PrelimDesignReview(

A

GE) 

Critical DesignReview 

Structure 

Equipment(

AVE). 

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Part 4 - Funding Requirements 

A. General 

Costs are based on Systems Office estimates and the Phase II 

contract negotiations and include an estimate of the schedule adjustment 

costs for the Gemini B System Segment. They reflect experience from 

the NASA Gemini Program modified by estimates of complexity factors, 

mission peculiar requirements, and number of vehicles. A brief resume 

of the effort, by work breakdown category follows: 

1. AVE Hardware. Includes the engineering and production costs 

of four Gemini B spacecraft, one aerodynamically similar flight article, 

and costs for non-recurring spacecraft development and qualification 

■■• 

testing. 

2. AGE. The hardware and integration of both factory and 

launch site AGE, and for modifications to existing NASA AGE for use 

by the MOL Program. 

3. Training. Provides for the conversion of the NASA Gemini 

simulator to the Gemini B procedures simulator, and for crew training 

and associated part-task trainers, including the modification of NASA 

egress and centrifuge trainers. 

4. Data and Support. Includes the effort in Technical Publication, 

Documentation, Repair Service, and Facilities Planning. Technical 

Publications consists of preparation, printing, and distribution of all 

AGE and AVE operation and service manuals and acceptance test procedures. 

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Documentation consists of preparation of System Engineering documents, 

and all graphs and art work. Repair Service consists of repair and 

maintenance effort for all Government Furnished Property. Facilities 

Planning is a MdDonnell effort in support of the Launch Site Planning 

Activities. 

5. Operations. Support of contractor personnel at the launch 

site, mission control, and other remote sites. The activities at 

Vandenberg AFB include support of site activation spacecraft receiving 

and inspection, buildup and support of integration, checkout, countdown, 

and launch. At the Mission Control Center, McDonnell engineers will 

support the mission director during the flight by. monitoring Gemini B 

status displays and by reviewing data. 

6. Management. Includes Configuration Management, Data Management and Schedule Control. 

The funding summary is shown by work breakdown structure on 

Chart IV-11. It includes deferred and identified program changes which 

are described in paragraphs B and C. 

B. Deferral Status 

Several tasks were deferred from the contract negotiations. 

The estimated costAnd reason for the deferrals is described below. 

Deferral costs included in the Gemini B segment total costs are shown 

in Figure IV-2. 

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1. Pad Abort Control System (PACS) - t2.500M. As a result of 

crew safety analysis it was determined that the probability of survival 

for aborts on or near the launch pad was extremely low. Consequently, 

the Gemini B baseline was revised to include the incorporation of a thrust 

vector control system to assure optimum spacecraft orientation during 

abort. The system consists of solid propellant thruster units and an 

attitude control electronics package. 

2. Spares - $14,000M. This item includes services and equipment which could not be firmly established until NASA Gemini residual 

spares and equipment inventories and their availability to support MOL 

were identified. 

3. Remote Site - $25.000M. This item includes the post FY 67 

'remote site operations effort and the associated McDonnell mission planning 

effort. ,These efforts required further definition prior to contractual 

action. 

4. Cost/Schedule Planning and Control Specification (C/SPCS)-$2.50M 

This item includes the effort to operate and maintain a Cost/Schedule 

Planning and Control System throughout Phase II and the associated reporting requirements which will satisfy the MOL Systems Office management 

needs. 

C. Identified Program Changes 

Revised or new technical requirements and management decisions 

since December 1966 have resulted in added tasks. The estimated cost 

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and reasons for these change items are described below. The costs 

are included in the Gemini B segment total costs shown in Chart IV-4. 

1. Diode Assembly - $.030M 

Interface discussions between McDonnell and Douglas 

disclosed that the Laboratory supplied electrical power could exceed 

Gemini B voltage limits. In order to preclude this possibility, McDonnell 

was directed to install a diode filter assembly on the Gemini B side of 

the interface since it appeared that this would be lees costly than 

restricting the laboratory power. 

2.. Non-Availability of 16. AGE*Items - $.580M 

This includes the effort required to replace 16 NASA, AGE 

items that were previouSly assumed to be available. 

3. Increased Tension Strap Loads - $.900M 

The baseline GeminiB re-entry module to adapter attachment 

straps were identical to those used on the NASA Gemini. During load 

cycle 2 it was determined that these straps would not withstand the 

tension loads imposed by Stage 1 shutdown. After a thorough review 

by the SPO and Aerospace, it was determined that these straps and the 

associated supporting structure should be increased in strength to 

avoid later and more costly design. 

4. Electrical Substitutes - $.200M 

This item includes the delivery of two additional Gemini B 

to Laboratory Vehicle electrical substitutes to support Douglas testing 

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at Huntington Beach. This requirement resulting from an Interface 

Control Meeting, requires Douglas to use substitutes for interface 

testing at Vandenberg AFB and at Huntington Beach. 

5. Recovery Documentation - $.100M 

Documentation support by McDonnell to aid in defining 

the Gemini B recovery requirements. Work has started and will 

continue through FY 70. 

6. Facility Design at Vandenberg AFB - $.100M 

Due to a change in responsibility after negotiations, 

McDonnell was required to define Gemini B peculiar support buildings 

at Vandenberg AFB. Additional effort in this area will be included 

in Remote Site effort. 

7. Adapter Shock Tests - $.100M 

The MOL Systems Office requested that the quantity of 

development test articlei being delivered to Douglas by McDonnell be 

increased in order that Douglas could complete the three required 

separation shock tests at station Z-49. 

8. Blast Shield - $4.000m 

After several extensive reviews by the MOL Systems Office 

and Aerospace, it was determined that the most effective method of 

providing protection to the laboratory unpressurized section was to 

change the baseline from local protection to an integral blast shield. 

McDonnell is presently engaged in design effort to redefine the 

McDonnell-Douglasinterface. 

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9. Ground Activated Abort Lights - $.100M 

This item will modify the baseline Gemini B by adding 

a separate Mission Control Center activated abort light and a 

separate guidance switch-over light also activated from the Mission 

Control Center. This change will also require a modification to the 

Laboratory Space-Ground Link. System (SGLS) to accommodate the 

necessary uplink and interface to Gemini B. 

7 

10. Positive Separation -- $.020M 

As a result of crew safety studies by McDonnell, Martin 

and the MOL Systems Offide it has been determined that a booster 

destruct inhibit will improve crew abort survivability. This change 

is madeto assure that the booster will not be destructed until the 

spacecraft has separated during abort. 

1 

11. Fire Hazard - $2.000M 

McDonnell is presently conducting a two month study as 

a result of the recent Apollo fire. The results of this two-month 

study are expected to lead to some changes in spacecraft equipment 

and procedures. 

12. Heat Exchanger - $.050M 

Based on a change of inlet temperature on the Laboratory 

side of the heat exchanger, McDonnell was requested to submit an 

Engineering Change Proposal (ECP) to add heaters to the warmant loop 

of the Gemini B.- These heaters will be powered from the laboratory during 

on-orbit dormant storage and are required to assure the adequate 

temTerature of the Gemini B. 

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13. Egress Trainer Operation and Maintenance - $.350M 

This item will provide flight crew members with experience 

in egress from the Gemini B during post landing recovery operations. 

The basic vehicle was used by NASA during the Gemini Program and 

it is intended to reconfigure the trainer to a MOL Gemini B 

configuration for training of flight crew members and recovery crews. 

14. Abort Trainer Operation and Maintenance - $.450M 

This item includes effort by Ling-Temco-Vaught Co. (LTV) 

at Dallas to conduct an engineering simulation:for fast reaction pad 

and ascent mode aborts. his simulation will be similar to others 

that LTV did for NASA. Gemini B program, Martin, and McDonnell are 

also required to support this program. 

15. Gemini B Procedures Simulator (GBPS) Instructor at 

Remote Site - $.250M 

This- item will cover six contractor simulation instructors 

at the remote site. flight crew training will start twelve months . 

prior to scheduled launch of Vehicle 3. GBPS will also be used to 

rehearse Mission Control Center controllers for both the Gemini B 

Qualification and Vehicle 3 flights. 

16. T-IIIM Electrical Substitutes - $.050M 

This design effort is needed to support Martin with 

equipment for use in a functional test of the Gemini B and Booster. 

inertial guidance systems. 

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17. Data Recovery Capsule (DRC) Flotation - $.100M 

This is currently not included on either the DRC or 

Gemini B contract baselines, but will become a McDonnell responsibility 

and design effort will start in the fourth quarter FY 67. 

18. Survival Kit - $.150M 

This item was assumed during negotiations to be contractor 

furnished equipment. A preliminary list of equipment has been identified 

and McDonnell should commence effort in FY 68. 

19. Support of Air Training Cammand.(ATC) - $.050M 

McDonnell support to Air Training Command (ATC) for 

the development and preparation of training courses for flight 

controller training. 

20. Gemini B Procedures Simulator (GBPS) Computer 

Complex - $.640M 

The Gemini B Procedures Simulator (GBPS) Computer 

Complex shipped from NASA for MOL does not have the capacity to 

perform the MOL GBPS functions. The primary reason is the MOL 

requirement to generate and transmit booster telemetry data in 

operational data format, the ability to operate closed loop with the 

Mission Control Center, dynamic simulation of operating conditions, 

and the requirement to verify Mission Control Center computer software 

programs. 

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SECTION V 

TITAN IIIM LAUNCH VEHICLE 

(Martin Marietta Corporation and Associates) 

Part 1. Segment Description and Performance Requirements 

The Titan IIIM Launch Vehicle, an adaptation of the Titan IIIC 

Vehicle, is required to be capable of inserting from Vandenberg AFB 

launch complex, without yaw steering, a payload of 31,325 pounds, 

minimum, into a 90 degree inclination orbit with perigee of 

80 nautical miles at 55° north latitude, and apogee of 180 nautical 

miles. 

The Titan IIIM Launch Vehicle for MOL consists of three propulsion stages. Stage Zero provides the initial thrust from two 

120-inch diameter, seven-segment solid rocket motors. Liquid rocket 

engines are used in Stages I and II which together comprise the vehicle 

core. The forward compartment of Stage II contains the inertial 

guidance system, flight controls, electrical systems, instrumentation 

systems, and. tracking and flight safety equipment. The interface 

between the Launch Vehicle and the MOL Orbiting Vehicle occurs at the 

forward end of Stage II, since the Titan III transtage will not be 

used for MOL. 

Performanceidesign requirements for the Launch Vehicle segment 

are based upon MOL-pecUliar requirementa for flight dynamics, 

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payload weight, orbital parameters, crew safety, system effectiveness 

and launch operations. Other system requirements remain as developed 

for the Titan IIIC system. Maximum use is being made of Titan IIIC 

design and components, with changes only where necessary to satisfy 

MOL-peculiar requirements. 

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Part 2 - Operating Concept 

The system provides for launching a vehicle within a one-hour 

window on any of three successive days specified two weeks in advance 

of the scheduled launch date. Provisions are made for a hold 

capability including programmed holds at appropriate points in the 

pre-launch countdown sequence. 

The Titan IIIM will be capable of launch within the launch 

window on the next day following'a launch abort which occurred during 

the terminal count and which was caused by * conditions external to the 

launch vehicle/Aerospace Ground Equipment (AGE), except when core 

propellants or solid rocket motor thrust.yectar injectant unloading 

is required. 

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( 

Part 3 - Development and Schedule Considerations 

The overall Titan IIIM system involves four contractors: 

Martin-Marietta Corporation (MMC); United Technology Center (UTC); 

Aerojet General Corporation (AGG); and AC Electronics Division of 

General Motors (AC). Martin has reponsibility for the Titan II1M 

airframe and launch vehicle systems integration, checkout and launch. 

Each of the other contractors is responsible for the on-pad checkout 

of his portion of the system. 

Design changes to the Titan IIIC and development requirements 

necessary to meet the MOL requirements, discussed by individual 

contractors, are as follows. The schedule is shown in Figure. V-1. 

A. Titan IIIM Airframe - Martin 

1. Development - Flight Hardware 

Modifications and redesigns associated with the airframe 

involve the structure, flight controls, hydraulic system components, 

flight sequencer, malfunction detectioU equipment, explosive bolt 

ordnance and fuel tank pressurization system. 

(a) Stage I Airframe 

The method of construction of Stage I remains 

virtually unchanged from that of Titan IIIC. Principal configuration 

and structural modifications will be: 

(1)' Relocation of the solid rocket motor attach 

fittings. 

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(2) An increase of 37 and 31 inches, respectively, 

in the lengths of the Stage I oxidizer and fuel tank barrels. 

(3) Redesign of the Stage I heat shield (boottail), 

primarily the floor, to satisfy the new requirements of the Stage I 

propulsion system. 

(4) Skin gauges, stiffners, raceways, longerons 

and internal conduits will be modified as dictated by the increased 

thrust of both the solid rocket motors and Stage I engines plus the 

increased length of Stage I for added fuel capacity. Various 

structural beef-ups are also involved for increased crew safety. 

(b) Stage II Airframe 

A redesigned Stage II forward skirt is needed to 

house the equipment normally located in the Transtage Control Module. 

With the exception of the forward oxidizer skirt, the basic configuration and method of construction of Stage II remains virtually 

unchanged from that of Titan IIIC. Principal configuration and 

structural modifications will be:,. 

(1) Replacement of the existing forward oxidizer 

skirt with a redesigned skirt which will be attached to the forward 

oxidizer tank juncture frame. 

(2) Replacement of the existing forward oxidizer 

tank juncture frame with a redesigned frame which will have attachment 

provisions to accept the new forward oxidizer skirt. 

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(3) Replacement of the equipment and instrumentation trusses in the forward oxidizer skirt with trusses of a new 

design. 

(4) Addition of solid rocket motor attach 

provisions on the new forward oxidizer skirt. 

(5) Deletion of retro-rockets including supports 

and attachment provisions. 

(6) Skin gauges, stiffeners,.z.raaewgys,Aongerans, 

and internal conduits will be modified for compatibility with the 

vehicle loads and propulsion requireMents. 

Deaign engineering of these changes was begun in 

April 1966 to meet the original launch schedule. With the current Mt 

schedule, this activity will be held to a minimum in FY 67 and 68 

commensurate with meeting the critical milestones of a Controls 

Mock-UP (CMU) in the third and fourth quarter of FY 69 plus the 

Structural Teat in the fourth quarter of Ft 69 and first quarter of 

FY 70. 

(c) Flight Controls. 

A triple redundant analog autopilot with-majority 

voting circuitry is required for adequate crew safety. Existing 

designs must be modified to meet the guidance requirements during 

Stage Zero flight and to also accept either the Booster Inertial 

Guidance or the Gemini Inertial Guidance inputs. The design activity 

will be conducted at a rate to meet the Controls Mock-Up milestone 

in the third quarter FY 69. 

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(d) Hydraulic System Components 

In order to mate with the redundant flight control 

system, a three torque motor majority voting actuator must be 

designed. In addition, the Stage I hydraulic system requires a 

redundant hydraulic power supply. This design effort must also meet 

the Controls Mock-Up milestone. 

(e) Flight Sequencer and Malfunction Detection Equipment 

Crew Safety and mission success dictates redundancy 

of the flight sequencer. Due to excessive weight of the Titan IIIC 

flight sequencer relay matrix, when expanded into a redundant system, 

a new light-weight sequencer is mandatory. 

The Malfunction Detection system requirements for 

adequate crew safety require the major component redesign for the 

Titan IIIM. New design logic is needed plus crew displays in the 

Gemini. The addition of new electrical equipment plus relocation of 

some existing equipment requires almost complete redesign of the 

Titan IIIC vehicle wiring. 

(f) Explosive Bolt Ordnance and Fuel Tank Pressurization 

The heavier payload and increased thrust requires 

redesign of the Titan IIIC explosive bolts. Increased fuel tank 

capacity of Stage I necessitates some redesign and resizing of the 

fuel tank autogenous pressurization system. 

2. Production - Flight Hardware 

In all the above areas of airframe modification and 

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redesign, the requirement exists to meet the Controls Mock-Up and 

Structural Test milestones with production articles in late FY 69. 

Component development will be completed during the third quarter of 

FY 69. Design insurance and structural tests will be complete during 

the first quarter FY 70, with First Article final assembly also 

complete during first quarter FY 70. 

3. Development and Production - Aerospace Ground Equipment 

(AGE) 

The numerous changes to the Titan IIIC vehicle plus the 

conversion from an Integtate Transfei and Launch (ITL) to the MOL 

Space Launch Complex at Vandenberg AFB necessitate resdeSign or 

resizing of many items such as the Air Conditioning Liquid Nitrogen 

converters and many electrical components. Delivery of AGE to 

Vandenberg AFB, initial installation and systems test must begin 

during the third quarter FY 69. Final installation and checkout for 

booster AGE should be _complete during the third quarter FY 70 in 

order to permit facility acceptance just prior to the first launch in 

April 1970. 

B. Liquid Rocket Engines - Aerojet 

The increased thrust requirements and higher altitude 

performance levied on the Stage I and II liquid. rocket engines due, 

respectively, to the MOL payload weight and the increased Stage Zero 

performance demanfUtedesign of combustion chambers, injectors and 

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frames for both stages, plus development of a 15:1 expansion ratio 

nozzle, an ablative skirt and higher performance turbopump for 

Stage I. 

1. Development -Stage I Engine 

(a) The first new combustion chamber was delivered for 

development testing in the fourth quarter FY 67. Testing has been 

accomplished to determine thermodynamic, flow and durability 

characteristics prior to the development of the injector, and the 

subsequent chamber-injector compatibility tests. A Preliminary 

Design Review (PDR) has been completed and a Critical Design Review 

(CDR) is scheduled in the first quarter FY 68. 

(b) Dynamically Stable,Injector 

Fabrication of various baffle configurations, using 

developed injectors, has been accomplished. Pulse tests have 

established the baffle configuration and confirmed the dynamic 

stability. 

(c) Ablative Skirt 

A conservative thickness basic design skirt was 

delivered in the fourth quarter FY 66. Component tests have been 

run on several skirts with the new combustion chambers and baffled 

injector to acquire data such as char rates and back wall temperatures. 

Final designs will be formulated from these tests. The Critical 

Design Review was conducted in March 1967. 

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(d) Turbopump 

Increased Stage I thrust performance demands 

higher fuel flow rates and therefore higher rotational speeds than 

the Titan IIIC turbopump can reliably provide. Therefore, redesign 

of the pump high-speed shaft was required plus analysis of the 

compatibility of such redesign with other components of the gearbox 

assembly. The first prototype pump assembly was available for testing 

in the second quarter FY 67. Critical Design Review will be 

conducted in the first quarter FY 68. 

(e) Engine Frame 

Dual hydraulic actuators for MOL crew safety plus 

increased engine thrust require a new engine frame. The first unit 

was delivered in the third quarter FY 67 for a static load test 

followed by engine firing. The Critical Design Review is scheduled 

for first quarter FY 68. 

2. Demonstration - Stage I Engine 

Procurement and Assembly of prototype components for 

overall engine evaluation began in the second quarter of FY 67 and 

will be followed by a final engine qualification teat demonstration 

on two complete engines during the second and third quarters FY 68. 

3. Production - Stage I Engine 

The lead time for engine production is.17 months. This 

requires initiation of production 4 months prior to completion of 

development/demonstration testing in order to deliver the first engine 
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to the Vertical Test Facility (VTS) at Martin in September 1969. This 

overlap could be reduced by installing the engines subsequent to the 

Vertical Test but would not be desirable from an overall system viewpoint. The first engine will be completed in July 1969 with 

subsequent engines at 2-month intervals. 

4. pezelmessielIkilne 

Necessary changes to the Stage II engine are not as 

extensive as for Stage I and therefore result in much less effort and 

cost. 

(a) Combustion Chamber' 

Increased burn time and a basic chamber-injector 

incompatibility resulted in redesign of.the,chamber tubes to more 

effectively withstand this environment. Design of the new chamber has 

been completed, and it is available for development/demonstration 

testing in the second and third quarters FY 68. 

(b) In ector 

The present Gemini Launch Vehicle baffled injector 

is being redesigned to provide better film coolidg and uniform mixture 

ratio across the face. The resultant design should produce improved 

performance needed for the MOL mission. 

4 

(c) Engine Frame 

Dual hydraulic actuators resulting from the MOL 

flight control redundancy requirement demand redesign of the engine 

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frame. The first unit is scheduled for testing in the fourth quarter 

FY 67. 

(d) Miscellaneous 

Several small items, such as redesigned fittings,. 

plus the gearbox develnped for the Gemini Launch Vehicle are 

scheduled for incorporation. These are low-cost product improvements 

and have minor effects on development/demonstration schedules. 

5. Demonstration - Stage 11 Engine 

Because the changes to the Stage II engine are not as 

extensive as those to the Stage I, coat and schedule reductions will 

be effected by eliminating the evaluation tests and conducting only 

the formal demonstration tests. These tests, will be identical to 

those of Stage I. They will be conducted during the second and third 

quarters FY 69 to be consistent with FY fund ceiling limitations. 

6. Production - Stage II Engine 

The production lead time situation will necessitate 

initiation of production 9 months prior to completion of development/ 

demonstration testing. The delivery schedule will match that of the 

Stage I engine. 

C. Solid Rocket Motors - UTC 

1. Development 

The MOL payload requirement places maximum design demand 

on the solid rocket. motors. To achieve the required Stage Zero thrust, 
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two segments will be added to the existing 5-segment 120-inch 

diameter motors, a longer forward closure will beincorporited, the 

thrust vector control (TVC) fluid tankage and injectant valves will be 

redesigned and the nozzle and propellant modified. 

Verification of the design parameters of the nozzle 

materials, internal motor ballistics and TVC system, by means of 

sub-scale testing, was completed in the fourth quarter FY 67. 

Design of the forward closure, nozzle and TVC system 

continued into the last quarter of FY 67 as follow-up to the sub-scale 

testing. 

2. Demonstration 

The first development motor test is scheduled for the 

first quarter FY 69. Three motors will be tested through the third 

quarter FY 69. Five additional motor tests will constitute the 

Preliminary. Flight Rating (PFRT) demonstration during FY 69 and 70. 

Fabrication of these motors begins in the first quarter FY 69 to allow 

sufficient lead time for the critical forgings of the forward closure . 

and the nozzle shell forgings. 

3. Production 

Due to the long lead time for parts processing and 

fabrication of the solid motor cases and nozzles, considerable overlap 

of flight motor production with development and PFRT motor fabrication 

is necessary. Production is scheduled to begin during the fourth 

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quarter FY 69 to meet a first flight motor acceptance in the first . 

quarter FY 70 followed by delivery in the second quarter FY 70. 

Additional flight motor pairs will be delivered initially on fourmonth centers. 

D. Inertial Guidance System - AC Electronics 

A redundant guidance and control system in the Launch Vehicle 

is required to meet the MOL crew safety and mission success requirements from launch to orbital injection. Since crew safety is 

particularly critical during Stage Zero flight, a triple redundant 

flight control system will be employed in that stage with dual 

redundancy of guidance system functions in Stage I and Stage II. 

During the last 120 seconds of Stage II. flight, the vehicle will be 

beyond the range of the ground radar tracking stations used as part 

of the ground-based slow malfunction detection system. Under this 

condition, a third redundant configuration will be employed to provide 

an airborne autonomous slow malfunction detection system. 

The basic components of the guidance system for the Titan IIIM 

consist of an Inertial Measurement Unit (IMU), Missile Guidance 

Computer (MGC) and a Signal Conditioner (SC). No design effort is 

required for the Inertial Measurement Unit since it is basically the 

same as presently used in the Titan IIIC. The Missile Guidance 

Computer is basically a Univac 1824 computer, but requires redesign 

for adaptation to the 'MOL requirements. The Signal Conditioner is a 

new design which handles the instrumentation requirements, consists of 

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micro-electronic circuits, is within the present state-of-the-art and 

is available for long lead procurement. 

1. Development and Test-Flight Hardware 

In order to. obtain maximum flight reliability confidence 

for the MOL mission, the inertial guidance system components are 

programmed for use on two Titan IIIC flights between the first and 

second quarter FY 69. Therefore, these flight dates establish the 

first major milestone for the AC development schedule. The basic 

design activities for the computer and signal conditioner began in 

May 1966 and will continue through the development and qualification 

tests during the first quarter FY 68. These qualification tests,. 

prior to the Titan IIIC flights, consist primarily of the dynamic 

tests, such as acoustic, shock, acceleration and vibration. Additional 

flight proof testing will occur in the second quarter FY 69 using 

1 early MOL production articles. 

2. Production - Flight Hardware 

The initial production, started in September 1966, will 

provide computers for the Titan IIIC flights plus one for the 

0 qualification testing and two others to support the systems integration 

testing at AC and the controls mock-up at Martin. Subsequent to the 

Titan IIIC flights, computer production for the MOL flight articles 

will be slowed down to a minimum rate of 1 per quarter. The normal 

production capacity is 4 per month. Production will, therefore, be 

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completed by the end of FY 69. Production of the signal conditioners, 

started in September 1966, will provide for the Titan IIIC flights 

plus a spare and 1 for the qualification tests. No signal conditioners 

are required to support the systems integration testing at AC or the 

controls mock-up at Martin. Production of the Titan IIIM flight 

articles will - continue at a minimum rate of one per quarter, to 

completion by the first quarter of FY 69. Sufficient Inertial 

Measurement Units are presently available to support the Titan IIIC 

flights and none is needed for qualification testing. Three LMU's 

are available from the Titan III resources and will be run through a 

modification line to incorporate some improved components which will 

update them to new production quality. This activity will begin in 

the third quarter FY 68 and result in three IMU units for about 

one-tenth the cost of each new production unit. New unit production 

will also begin in the third quarter FY 68 and terminate with the final 

unit in the second quarter FY 71. 

3. Development and Production - Aerospace Ground Equipment 

During the last quarter FY 66 basic design effort began 

on the Telemetry Data Monitoring Set plus modification effort on the 

Missile Guidance Alignment and Checkout Group and Missile Guidance 

System Test Set. CEI specification Parts I and II were provided in 

FY 67 on the Telemetry Data Monitoring Set and Missile Guidance System 

Test Set, both of which will receive major modifications. Minor 

modifications to the Missile Guidance Alignment and. Checkout Group 
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will be covered by changes to Titan IIIC specifications. Production 

to meet the installation checkout and launch support of the three 

Titan IIIC flights and subsequent MOL flights will include three 

Telemetry Data Monitoring Sets, three sets of computer ground support 

equipment and modification kits for five Missile Guidance- Alignment

and Checkout Group and Missile Guidance System Test Set units. 

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Part 4 - Funding Requirements 

A. General 

The costs shown for Titan II1M reflect the negotiated 

Phase II contracts, and include an estimate of the schedule adjustment 

cost. They provide for development and fabrication of seven vehicles. 

The primary development involves 7-segment solid rocket motors; 

15:1 ratio liquid engines and redundancy required for crew safety. 

The total costs, including deferrals described in paragraph B are 

shown in Figure V-2. A resume of the tasks included in the cost 

summary follows: 

1. AVE Hardware 

Includes airborne design and non-recurring tooling, which 

involves basic design, preproduction test hardware, airborne equipment 

evaluation and confirmation testing; airborne hardware production 

which involves fabrication, delivery and acceptance testing of airborne'' 

equipment, recurring engineering and recurring tooling; spares; and 

production propellants. 

2. AGE 

Includes the AGE design, production and tooling. 

3. System Test and Evaluation 

Includes system engineering associated with design 

requirements, analytical effort, system test planning, control 

requirements, system effectiveness and system integration. 

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4. Systems Engineering and Management 

Includes. program.control (costs and schedules management), 

configuration management and maintenance of Phase II documents.. 

5. Data 

Includes all reports generated for the program. 

6. Operations and Services 

Includes the receipt, installation, and checkout of Aav 

launch planning vehicle assembly, launch; pre..launch and post-flight •• 

evaluation. 

7. SE and TD 

Aerospace Corporation system engineering and technical, 

direction. 

B. Deferralatatus 

Several tasks were deferred from the contracts negotiated 

with Martin Company and AC Electronics Division. The estimated costs: 

and reasons for deferring these tasks are listed below. Deferred 

costs are included in the Titan IIIM Launch Vehicle total costs shown 

in Figures 

1. Martin Company 

(a) Several studies are currently underway to determine 

precise requirements for additional structural beef-up of Stage It, 

qualification •tests incorporation of POGO hardWare, failure analysis/ 

traceability effort, and'development of a 1 Amp-hour battery, The 

estimate for these efforts is $4.2m. 

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(b) Support to other MOL contractors is not fully 

defined at this time. The costs estimated for this effort are $2.5M. 

(c) The incorporation of GERTS has been directed by 

the MOL SPO. Estimated costs are $1.5M. 

2. AC Electronics Division 

Several tasks, such as validation of software for 

flights past the first MOL flight, rehabilitation of operational 

computers from qualification units and the controls mock-up (CMU) 

computer program were deferred during the negotiations. Since the 

original deferment approximately $0.2 million of FY 67 money has been 

applied.towards the CMU computer program. Estimated Cost for the 

remaining deferred items for FY 68 are $0.4 million. 

3. Aeroiet General Corporation 

Included in the deferrals are items such as POGO hardware, 

ablative skirt, stage II chamber effort, exit cover test, certain 

qualification testing, and a contamination task. Estimated.cost of 

this work is $3.7. million. 

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SECTION VI - FLIGHT CREW EQUIPMENT AND AMD SUPPORT 

This section describes the Flight Crew Pressure Suit Assembly, 

Flight Crew Diet and AMD Developmental and Operational Support. 

Discussion of each element_follows with consolidated funding requirements 

covering all the elements indicated in tabular form at the end of this 

section. 

Part 1 - Flight Crew Equipment 

A. Pressure Suit Assembly 

1. General Performance Requirements and Program Approach 

The Pressure Suit Assembly (PSA) provides for a breathable 

atmosphere under sufficient pressure to gustain life, provide thermal 

protection, micro-meteoroid protection, feeding and drinking, body 

waste storage, and communications. Mobility provisions will permit the 

crewman, pressurized if necessary, to accomplish MOL mission requirements. The MOL PSA development will be managed by the MOL Systems 

Office. Technical guidance, testing and operational support will be 

provided by a technical team formed by the Aerospace Medical Division 

of AFSC. The Phase I effort completed in August 1966 resulted in the 

identification of PSA technical requirements, interface requirements, 

the solution of specific MOL technical problems, and the publication of 

a Performance and Requirements Specification. ,A request for proposal 

for Phase II suit development was subsequently. released. The RFP was 

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unique in that it required the submission of a sample suit to demonstrate 

the contractor's technical competence and approach. The source 

selection evaluation of the proposals and the issuance of a contract to 

the winner, Hamilton-Standard, has since proceeded on a schedule which 

will permit the timely delivery of suit assemblies for systems engineering 

and crew training. Pressure suits are required during the 3rd quarter 

of FY 68 for Air Force developmental testing and for Douglas and McDonnell 

engineering and design verification tests. The numbers and delivery 

dates for the suits for Douglas and McDonnell are called out as part of 

the vehicle contracts. The PSA's required to support the MCL Program 

will be provided as government furnished equipment and will be delivered 

in accordance with the schedule shown in Figure VI-2. 

2. Operating Concept 

The PRA will be designed to permit its use with the• 

environmental systems of the Gemini and Laboratory Vehicle and high 

pressure gas supplies from the emergency oxygen system or accumulators 

on the vehicle. In the extra-vehicular modes, it will be able to operate 

either on its emergency integral supply carried by the crewmen, or 

through umbilical system from the Orbiting Vehicle. The suit assembly . 

will include an integral communications sgbsystem that will permit 

exchange of information between crewmen during suited phases, recording 

of data on the Orbital :Vehicle system, and communications through the'- 

vehicle to the ground. 
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B. Flight Crew Diet (Feeding System Assembly (FSA)) 

General Performance Requirements and Program Approach 

The diet will contain adequate proteins, fats, carbohydrates, 

minerals, vitamins, and energy to maintain the crewman's health during 

the 30-day MOL flight and recovery period. The Gemini and Apollo 

freeze-dehydrated diet has been selected as the most appropriate starting 

point for diet development to meet MOL requirements. This present diet 

'provides 2650+250 K Calories, occupies 143 cubic inches of space, and 

weighs 1.7 pounds, including packages, per man per day. 

It is planned to manage the MOL diet development prograt 

within the MOL Systems Office. Technical aspects have been assigned 

to the Aerospace Medical Division of the AF Systems Command. Included 

within the technical team will be members of the U.S. Army Natick 

Laboratories, for food preparation and packaging, and the NASA. Manned 

Spacecraft Center, for operational experience with space flight feeding 

systems. Although the 'Apollo Block I diet serves as a starting point, 

it does not meet MOL requirements with respect to preparation time, 

adaptability to MOL mission timeline variation, or effective nutrition 

for long periods of time on orbit. The general experience indicates 

that the Apollo diet: 

will require some 1.5 to 2 hr per man per day for preparation 

and eating; 

does not have a nutritionally adequate liquid component; 

bite sized food items do not exploit the current known 

capability of the food industry. 

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Therefore, reconfiguration of the Apollo diet will be accomplished: 

appropriate nutrient liquid components will be developed 

and added; 

bite-sized food items will be improved; 

both will be combined with freeze-dehydrated items for 

a complete feeding system. 

The flexible menu will thus consist of one freeze-dehydrated meal per 

day requiring about 1/2 hour for preparation combined with supplementary 

feeding based on bite-sized or bar sized foods requiring no preparation, 

and the nutrient liquids prepared from a dehydrated powder. 

This new MOL menu must have a testing cycle which AMID will 

undertake in their manned space chambers as part of their MOL simulation 

studies. The prospective FSA contractor will provide the developmental' 

test rations for these AMID chamber simulation studies. The feeding 

program is based on a continuing development/production effort aimed at 

improving palatability, reducing preparation time and increasing the 

variety of available foods with a view to enhancing crew acceptance. 

Food production specifications, menu, and quantities for testing, training. 

and flight use will be based on recurring testing requirements so as 

to maintain a continuous high quality production capability for producing 

space-qualified food. The U.S. Army Natick Laboratories will assist in 

developing improved foods and packaging, and will maintain a capability 

for solving production and quality control problems as they develop. 

Procurement has been initiated and cOntract proposals have been received 

DORIAN 

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• from three prospective contractors. Contract award will be on a schedule 

'Which is consistent with the need to assure provision of test diets for 

the AND chamber simulations. The milestones schedule for the FSA is 

slibwn on Figure VI-$. 

-DORIAN 

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Part 2 - Aerospace Medical Division Development and Operational Support 

Analysis of the degrees of crew activity required to meet the 

MOL program objectives and the ideal operation of the vehicle indicates 

a need to maintain alert and efficient crewmen throughout the 30 day 

mission. The bioastronautics program has been constructed to meet this 

requirement. 

A review of the manned experience in space flight up through 

14 days of flight, has confirmed that predicted physiological and 

psychological adaptive changes in man as he adjusts to living and working 

the space environment are occurring. The adaptive changes were predicted 

through the use of ground based analogue simulation. The simulation 

events, however, had been limited to predicting accurately the areas of 

adaptation and the direction the adaptive processes would take. After 

14 days weightless experience, the prime areas of concern are the 

maintenance of cardiovascular, and musculo-skeletal tone and effectiveness. 

Although these changes have been seen in flight, the 14 day exposure is 

still of insufficient duration to cause mission limitation. The lack of 

fidelity of early simulation had prevented prediction a mission 

significance of the changes as they would relate to the crewman's ability 

DORIAN Page 3 /of/3 Gpages 

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to continue reliable activity through postlanding recovery. As the space 

flight programs continued, the fidelity of the ground simulations improved 

and the accuracy in mimicking the actual space events has become much 

better. 

The 30 day MOL mission will provide the additional length of 

flight duration sufficient to permit the crew adaptation to reach levels that 

could effect the crewman's efficiency on orbit and, as a result, cause 

special problems for survival immediately after returning to the 

earth gravity. The extension of the mission from the present 14 day 

experience to 30 days also permits some new body systems, which have 

slower. physiological cycles, to reach mission significant levels of 

adjustment for the crewman. The candidate longer duration cyclic areas of 

concern include the neurological and endocrine systems, the cellular 

development process and the psychomotor response pattern. 

The MOL bioastronautics program is designed to identify and 

quantify the adaptive adjustment of the crewmen to 30 days of exposure in 

the space environment, and in those cases where unchecked adaptive changes 

are suspected to have an crew operational impact, to develop sufficient 

control measures to maintain the adaptive processes within acceptable 

limits for performing the MOL mission. The MOL program Office has levied 

the Aerospace Medical Division (AMD) of Air Force Systems Command to undertake an extensive simulation program to accomplish these goals prior to. 

DORIAN 

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the first manned flight. At the time of flight AMD has been requested 

to support the operations with medical personnel and inflight data 

collection to insure that the ground simulation predictions, and the 

planned control regimens are adequate. AMD is meeting this levy by making 

the MOL 30 day study and development requirements an important part 

of their in-house exploratory development program. Resources for these 

efforts are being individually negotiated to establish the best Air Force 

approach to meeting the 'MU needs. An example of such efforts is the 

approach to developing special equipments for crew conditioning and the 

measurement of crew status during flight. In the case of the crew 

conditioning' equipment, the AMD is providing the physiological criteria, 

developing prototype equipment and demonstrating the adequacy of design. 

After completion of this effort AMD will provide technical assistance to 

the Laboratory Vehicle contractor for final design and integration into 

the vehicle. AMD will develop the operational procedures and the inflight 

measurement techniques to insure that the devicee and procedures do 

maintain the crew. Another specific example of AMD's unique activity is 

the development of the inflight mass determination devices. AMD is 

following a similar invention, development and test of the equipment 

as noted for crew conditioning. At the time of commitment to MOL use, 

the same arrangement for technical assistance to the laboratory contractor 

will be used. .other equipments will be provided by AMD as Government 

Furnished Equipment (GFE). 

DORIAN 

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An important part of the AND support includes the validation 

of the ground simulation data under flight conditions. In this regard 

AMD, with the encouragement of MOL, has proposed to develop and flight 

test models of the crew conditioning exerciser and the mass determination 

devices on flights of the early NASA Saturn IVB Workshops. This 

preliminary verification of the data on crew adaptation to the space 

environment and demonstration of the adequacy of the control measures and 

procedures should permit a high level of confidence for MOL to undertake 

a 30 day weightless mission safely before the first MOL manned flight. 

The funds noted under this item reflect the matching or unique 

MOL funds identified as required to carryout the AND assigned tasks. 

A summary of present tasks and functions assigned to AMD include the 

following functions: 

1. Blood pressure system (GFE). 

2. Body temperature measurement device (GFE). 

3. Crew conditioning device development and standardization. 

4. Medical kit development. 

5. Crew clothing and hygiene equipment evaluation. 

6. Mass determination device development and standardization. 

AMD will also: 

1. Acquire and validate equipment for MOL preflight and 

post-flight examination and evaluation of the flight crew, (i.e., crew 

physiological adaptive changes due to weightlessness and operational 

control procedures). 

DORIAN Page/34f/3 6 pages 

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2. Accomplish MOL specific laboratory testing of human 

subjects in order to validate the MOL baseline flight environment as 

reflected in the Laboratory Vehicle and personal equipment design and 

operating specifications (i.e., testing of the- atmosphere, temperature, 

pressure and gaseous mixture, diet adequacy and other life support 

provisions). 

3. Provide bioastronautics technical assistance to the 

MOL program in the-physiological, medical and human performance areas. 

The funding schedule in Figure VI-3 will assure timely inputs to system 

design, integration, and operational concepts. 

-1 

.• 

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