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TOR7469(5510-41)-1,VOL II 

(U) Study of Utilizing Apollo for the MOS, Mission 

Volume II: Subsystem Studies-Applied Mehanies Division 

1.1 JANUARY 1965 

Pliepared by 

. . . 

THE APPLIED MECHANICSDIVISION 

Prepared for COMMANDER SPACE SYSTEMS DIVISION 

AIR FORCE SYSTEMS COMMAND 

LOS ANGELES AIR FORCE STATION 

• Los Angeles, California • 

EL SEGUNDO TECHNICALOPERATIONS•. 

AEROSPACE CORPORATION 

CONTRACT NO. 04(695)-469 

FV- OF - 

RET • , ORATION u LicikARy • 

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Report No. 

TOR-469(5510-41)-1, 

Volume II 

(U) STUDY OF UTILIZING APOLLO FOR THE MOL MISSION, 

VOLUME H: SUBSYSTEM STUDIES - APPLIED MECHANICS DIVISION 

Prepared by 

Applied Mechanics Division 

EL SEGUNDO TECHNICAL OPERATIONS 

AEROSPACE CORPORATION 

El Segundo, California 

Contract No. AF 04(695)-469 

Prepared for 

COMMANDER SPACE SYSTEMS. DIVISION 

AIR_ FORCE SYSTEMS,COMMAND 

LOS ANGELES AIR FORCE STATION 

Los Angeles, California 

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Report No. 

TOR-269(5510-41)-1, 

Volume II 

(U) STUDY OF UTILIZING APOLLO FOR THE MOL MISSION 

VOLUME H: SUBSYSTEM STUDIES - APPLIED MECHANICS DIVISION 

Prepared by Applied Mechanics Division 

Approved 

F.C. Strible 

Assistant Director 

System Design 

D. Willens, Director 

Systems Design Subdivision 

The information in a Technical Operating Report is developed for a particular 

program and is therefore not necessarily of broader technical applicability. 

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CONTENTS1 

SECTION 1. Design 

N. G. Ivanoff 

SECTION 2. Weights 

C. Stewart 

Experiments Integration 

J. J. Fastiggi. 

Crew Activity Time Allocations 

E. F.' Schmidt and F.E. Cook 

Power Systems2 

H. J. Killian 

Life Support 

C. C. Wright 

Performance 

K. N. Easley and J. M. Dreyfus 

Test Operations 

L. Cooper 

Reliability 

F. P. Klein 

Fluid Mechanics 

N. R. O'Brien and M. J. Adams 

Propulsion 

M. J. Russi 

Solid Mechanics 

J. Hook, D. E. Hargis, G. B. Fox and A.J. Victor 

SECTION 3. 

SECTION 4. 

SECTION 5. 

SECTION 6. 

SECTION 7. 

SECTION 8. 

SECTION 9• 

SECTION 10: 

SECTION 11. 

SECTION 12. 

1

See individual sections for specific tables of contents and lists. 

2See TOR-469(5510-41)-1, Volume III, Section 5 for additional Power 

Subsystems (power requirements) material. 

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FOREWORD 

This report is one of three volumes, which jointly document the 

results of studies performed by the Aerospace Corporation, with cooperation of the Air Force/SST, to evaluate the feasibility of utilizing 

the Apollo spacecraft to perform the Air Force MOL mission. The 

volumes comprising the results of the Apollo /MOL study are: 

TOR-469(5510-41)-1, Vol I Summary 

TOR-469(5510-41)-1, Vol II Subsystem Studies - Applied 

Mechanics Division 

TOR-469(5510-41)-1, Vol III Subsystem Studies - 

Electronics Division 

Volume I briefly describes the Apollo spacecraft and the Apollo 

elements considered for the Mc11, program, summarizes the results of 

Apollo/MOL configuration .Ladies utilizing the Titan III-C and the Saturn 

IB launch vehicles, and presents the overall conclusions reached. A 

brief summary of,

fhe subsystem studies and the potential growth capability 

of Apollo/MOL Saturn IB is presented in the second part of the volume. 

_Pie recommendations in Volume I are based upon analysis of overall 

system considerations and in a few cases may not completely reflect the 

recommendations associated with the detail subsystem studies reported. in 

Volumes II and III. 

Volume II is a compilation of documents which details the results of 

subsystem studies performed in the applied mechanics area to evaluate the 

Apollo/MOL requirements and to develop the tradeoffs required to make 

subsystem recommendations. The studies included in Volume II are in the 

areas of: 

Design Performance 

Weights Test Operations 

Experiments Integration Reliability 

Crew Time Allocation Fluid Mechanics 

Power Systems Propulsion 

Life Support Solid Mechanics 
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Volume III is a compilation of documents which details the results 

of subsystem studies performed in electronics and related areas in support 

of the Apollo/MOL study. The areas of study included in Volume III are: 

Attitude Control Subsystems 

Communication Subsystems 

Guidance and NavigationSubsystems 

Photo-Optical Subsystems 

Power Subsystems (Power Requirements) 

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INTRODUCTION 

This volume includes studies pertaining to the feasibility of the 

Apollo/MOL concept, comparisons of. Apollo/MOL with Gemini/MOL, and 

investigations of increased Apollo/MOL mission duration. In general these 

studies encompass the technical disciplines within the responsibility of the 

Applied Mechanics Division. The possibility of using Titan IIIC as a launch 

vehicle is investigated, and a number of variations .in the basic Apollo 

command/ service module package are studied in connection with the 

Saturn IB launth vehicle. The experiments and system requirements are 

those used for .the Gemini/MOL at the time of these studies, which is 

September. 1964.. 

The principal purposes of these studies then, are to determine 

the significant problems in adapting the Apollo spacecraft and supporting 

equipment to the MOL mission, the number of flights required_to perform 

all of the MOL experiments using 30-day missions, and whether missions 

of 60 or 120 days are feasible. If there are no significant problems in 

adapting Apollo/MOL, and if fewer flights are required than with Gemini/ 

MOL, then the systems are to be considered roughly competitive, subject 

to a cost comparison. 

The studies include the formulation of prelithinary designs of 

several configurations and the analysis of those configurations which appear 

promising. Estimates of required power and life support equipment, and the 

necessary modifications to the Apollo -system to provide that equipment, are 

made to fit the various design concepts. Power and life support systems are 

chosen for both two and three-man crews for mission durations of 30, 60, 

and 120 days. 

On the basis of the preliminary designs, equipment selections and 

other modifications to the basic Apollo system, estimates of system 

weights are made. Performance analyses are also carried out to 

determine the payload that can be placed into orbit. The payload capability 

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and overall system weight are then compared to determine the weight 

available for experiments in each configuration. Also, crew time allocations are estimated on the basis of previous studies for MORL, XMAS, and 

Gemini/MOL. Space availability is estimated directly from the preliminary 

design drawings. Using all of these. estimated data, and following the same 

constraints that are used in the Gemini/MOL studies, the experiments are 

allocated among as many flights as are needed to perform them all. This 

number of flights provides a direct comparison with the Gemini/MOL 

concept. To make this comparison more meaningful, the over-all system 

reliability of Apollo/MOL is estimated and compared with that of Gemini/ 

MOL. 

In order to determine whether significant problems exist, a 

number of specialized areas were investigated, including structures, 

propulsion, aerodynamics, and operational considerations. These tasks 

were addressed to specific problems such as ascent vibration and wind 

loading effects, the suitability of propellants, abort feasibility, ascent and 

re-entry capability, the availability of ground stations, and the security 

of communications. 

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SECTION 1 

DESIGN 

N. G. Ivanoff 

SUMMARY 

Several preliminary layouts and descriptions of Apollo /MOL configurations for each launch vehicle are presented. One drawing shows the Apollo/ 

MOL as launched by the Titan IIIC and three other drawings. show progressively 

improved configurations as launched by the Saturn (1B booster. 

The Apollo/MOL/Titan IIIC configuration, due to weight limitations, 

has no pressurized laboratory and can carry only a portion of the experiments 

carried by the Gemini/MOL/Titan IIIC configuration; however, experinients' 

would be conducted from within the Apollo capsule, which is not possible in 

the Gemini. 

The preferred configuration of the Apollo /MOL as launched by the 

Saturn IB consists of the Apollo command module and a combined laboratory/ 

service module. All of the MOL experiments, from a weight and volume 

standpoint, .,cari;:be integrated within the laboratory on any one flight. The docking 

and airlock system of crew transfer, as used for the NASA Apollo lunar mission, 

was accepted as feasible in this study. However, several iterations on methods 

of crew transfer for the Apollo/MOL/Saturn IB configurations were performed 

and are included. 

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CONTENTS 

Page 

1. DESIGN 1-5 

1.1 INTRODUCTION 1-5 

1.2 BASIC APOLLO/MOL VEHICLE GROUND RULES 1-5 

1.3 APOLLO/MOL CONFIGURATION CONCEPTS 1-6 

1.3.1 Apollo/MOL/Titan IIIC Configuration 1-6 

1.3.2 Apollo/MOL/Saturn IB Configuration Using the Geometry of 

NAA Extended Mission Apollo Concept II 1-8 

1.3.3 Apollo/MOL/Saturn IB Configuration Modified 

NAA Concept II 1-11 

1.3.4 Apollo/MOL/Saturn IB Designed Specifically to MOL 

Requirements 1-12 

1.4 PERTURBATION OF THE APOLLO/MOL CONFIGURATION . 1-13 

1.4.1 Configuration Arrangements and Laboratory Size 1-13 

1.4.2 Methods of Crew Transfer 1-14 

1.4.3 Extension of Mission Duration 1-1 5 

1.4.4 Addition of Artificial Gravity Capability 1-16 

1.5 CONCLUSIONS 1-17 

REFERENCES 1-31 

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ILLUSTRATIONS 

Page 

1-18 1-1 Minimum Modification Command Module, Service Module and 

Laboratory Module Using NAA Concept II Geometry 

1-2 Command Module, Service Module and Laboratory Module 

Modified NAA Concept II 1-18 

1-3 Command Module and CoMbined Laboratory/Service 

Module With the Propulsion System Aft 1-19 

1-4 Command Module and Combined Laboratory/Service 

Module Reduced in Volume With the Propulsion System Aft . . 1-19 

1-5 Crew Transfer Effected by a Tunnel Through Command 

Module Heat Shield 1-20 

1-6 Crew Transfer Effected by an Inflatable Pressurized 

Tunnel From the Command Module to the Laboratory 

Module 1-20 

S 

1-7 Crew Transfer Effected by the Repositioning of the Command 

Module Using a Mechanical Trunnion 1-21 

1-8 Crew Transfer Effected by Extravehicular Excursion. of Crew 

Members Through Air Locks 1-21 

1-9 Apollo/MOL NAA Concept II Configuration 1-23 

1-10 Apollo/MOL NAA Concept II Modified Configuration 1-25 

1-11 Apollo/MOL Configuration for Saturn IB Launch Vehicle . . . 1-27 

1-12 Apollo/MOL Configuration, for Titan IIIC Launch Vehicle . . 1-29 

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1. DESIGN 

1.1 INTRODUCTION 

The present Apollo/MOL program study is a brief investigation to determine 

the feasibility of the Apollo capsule as a replacement for the Gemini B capsule, 

both as the crew conveyance and re-entry vehicle because of its larger size 

and potential capability. The study considers the use of the Titan IIIC and the 

Saturn IB as launch vehicles using the present docking and air-lock system of 

the Apollo lunar vehicle (References 1-3 and 144). 

Several arrangements of the Apollo/MOL/Titan IIIC vehicle system were considered as possible candidate configurations to satisfy the MOL program requirements. Due to the limitation of the boost capability of the Titan IIIC launch 

vehicle, the resulting arrangement consists of only a modified Apollo command 

module and an unpressurized service module shown in Figure 1-12 (Drawing 

ES-0152-004). Several arrangements of the Apollo/MOL / Saturn 1B vehicle 

system were also considered as possible candidate configurations to satisfy 

the MOL program requirements; these are shown in Figures 1-1 through 1-4. 

An arrangement designed specifically for the Apollo/MOL program is shown 

in Figure 1-3 and in more detail in Figure 1-11 (Drawing ES-0152-003), and is 

discussed in Section 1.3.4. Various methods of crew transfer from the Apollo 

command module to the laboratory module were also considered and are shown 

in Figures 1-5. through 1-8. 

1.2 BASIC APOLLO /MOL VEHICLE GROUND RULES 

The ground rules applicable to this study are the same as the present MOL 

ground rules and are listed as the minimum vehicle requirements: 

a. Thirty-day orbit duration 

b. Two-man crew 

c. Integral launch 

d. Shirt-sleeve environment 

d. Test and experiment capacity commensurate with MOL package 

(Reference 161) 

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f. Rendezvous, docking and transfer provisions 

g. AMR launch 

h. Low orbit: 150 - 250 nautical miles 

i. Minimum change to the Apollo-system 

Several major Apollo subsystems will have to be modified and integrated into 

the over-all vehicle system to accomplish the Apollo/MOL mission. These 

are listed as follows: 

a. Life support system 

b. Environmental control system 

c. Power supply system 

d. Water and waste management system 

e, Attitude control system 

1. 3 APOLLO/MOL CONFIGURATION CONCEPTS 

1. 3. 1 Apollo/MOL/Titan IIIC Configuration 

The Apollo/MOL configuration, launched by the Titan 'MC booster and shown 

in Figure 1-12 was selected for study from several candidate configurations. 

The MOL experiments (Reference 1-1) are shown integrated into the Apollo 

command module and the service module which also serves as the adaptor 

to the Titan IIIC transtage. Due to the limited payload capability of the 

Titan IIIIC booster (Section 2. 1), only a portion of the required MOL experiments can be carried for each mission flight; also, because of weight limitations, no pressurized laboratory module can be carried into orbit for an 

integral launch, thus requiring the integration of the experiments, which must 

be handled directly by the crew, into the Apollo command module. The experiments and related equipments which can be operated remotely are integrated 

within the service module. 

The integration of the MOL experiments into the Apollo capsule would require 

more than the minimum Apollo modifications required by the ground rules. 

Although an effort has been made to locate the various. experimental equipments within the capsule, so as not to interfere with the existing Apollo systems, 

many structural changes would have to be made. In case of the optical equipment, for example to mount the pointing and tracking scope through the capsule 

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wall, a pressure seal would be required within a structural well to allow the 

P/T scope to be stored during ascent and extended in orbit for observation. 

Provision also must be made for sealing the external movable head which 

rotates through an angle of 360 degrees. An aerodynamic fairing .to cover the 

scope would also be required both to reduce drag and to protect the optical 

surfaces of the instrument. 

Similar provisions would be required for the periscope/star theodolite optical 

instrument. Also, the geodetic camera which observes both ground targets 

and .a celestial reference requires optical ports which must be sealed around 

the instrument peripheries in the basic Apollo structure. 

The air-lock hatch and crew transfer tube of the Apollo capsule would have 

to be modified by the addition of an internal telescoping extension tube having 

a. second air-lock hatch to allow a single crew member to exit and enter the 

capsule without complete depressurization of the capsule. The astronauts 

maneuvering unit (AMU) could be stored in the crew transfer tube and put on • after leaving the capsule. 

The remote maneuvering unit (RMU) is located in the open end of the service 

module and is ejected from the vehicle into space where it is operated remotely 

from within the command module. No consideration has been made for recovery 

of the RMU after use because of difficulty in its refurbishment since the RMU 

as presently conceived cannot be brought into the pressurized capsule for 

servicing. 

The various displays are located on the left hand side of the pointing and 

tracking scope and are linked electronically to the various experiments 

mounted within the Apollo capsule and the service module. The radiometer 

is modified to be remotely operated so that the filters for various wavelength 

radiations are changed or rotated by a selector mechanism operated from 

the command module. The radiometer and the laser ranging unit are slaved 

to the pointing and tracking scope so that all three are looking at the same 

target when observations. are made. 

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The various supporting optical or electronic systems and antennae are 

packaged within the service module and are deployed when needed to make 

observations or to pick up radiations. 

The various major subsystems are located within the service module and are 

connected to the Apollo capsule through an umbilical connector. The life 

support system requires the recycling of the capsule air through lithium 

hydroxide cartridges for the removal of carbon dioxide. These cartridges 

are located within the service module because of both weight and volume 

limitations of the Apollo capsule. 

The Apollo heat shield is modified for an earth orbit re-entry and the 

attachment of a retro-package consisting of six XM-85 solid. propellant 

rocket motors, any five of which will provide a sufficient retrograde 

velocity to initiate re-entry. Among the major problems limiting the 

experimental capability of the Apollo command module are the recovery 

and abort systems weight restriction. Although the weight limitations 

permit only a restricted number of experiments to be carried per flight 

(Section 3), on the Apollo/MOL/Titan IIIC configuration, at least one 

advantage the Apollo capsule has over the Gemini B capsule, for example, 

is the capability of conducting some experiments without leaving the Apollo 

capsule. 

1.3.2 Apollo/MOL/Saturn IB Configuration Using the Geometry of 

NAA Extended Mission Apollo Concept II 

A preliminary layout, Figure 1-9 (Drawing ES-0152-001), was made of the 

Apollo/MOL on the Saturn IB launch vehicle with all the MOL experiments 

integrated into the laboratory module using the geometry of the North 

American Aviation Extended Mission Apollo Concept II (Reference 1-2). 

Reference 1-2 is a study which had been performed by NAA on the use of 

the Apollo vehicle as part of an integrally launched space laboratory with 

mission durations of 120 days or more, using a resupply technique and 

crew rotation. The resulting configuration, using the geometry of NAA 

Concept II, ,was based on the minimum modification of the Apollo command 

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module and of the service module. Therefore, the heat shield of the Apollo 

vehicle is left unmodified from that required for re-entry from a lunar orbit, 

but a retro-package of six XM-85 solid propellant motors was incorporated 

with the heat shield of the Apollo command module for effecting a re-entry 

from an earth orbit. The service module retained .the same basic structure 

but was modified to allow for the incorporation of the following: the retropackage, the LEM rocket engine of 10.5 K instead of the SM engine of 21. 9 K 

thrust, the propellant tanks of smaller volume, the life support system, the 

environmental control system, the fuel cell power system, and the reaction 

jet attitude control system for CM/SM repositioning maneuvers. 

The NAA Extended Mission Apollo Concept II laboratory module for six men 

was duplicated .in the arrangement of the floors, docking cones, umbilical 

connectors, etc. , and the MOL experiments were appropriately located to 

suit the mission requirements. Similar to the NAA Concept II, one compartment of the laboratory module nearest the command module, was designated 

for rest, feeding, personal hygiene, and recreation of the crew. The other 

compartment was designated for the location of all of the MOL experiments. 

An attempt was made to locate the experiments and the instruments in such a 

manner that related experiments could be observed from adjacent displays 

and instruments. Some thought also was given to the disposition of the weight 

of various pieces of equipment to result in a reasonable center of gravity 

location. 

The two compartments are connected by a tunnel which serves both as a 

structural support for .the large diameter floor panels and as a means of easy 

transfer from one compartment to another by the use of hand holds. Two sets 

of hand rails are provided in each compartment to facilitate crew motion in a 

weightless environment. A track for a restraining mechanism is provided in 

the "floor" of each compartment so that the astronaut can keep properly 

oriented. Seats with restraining straps are provided to maintain a suitable 

frame of reference for operation of various display consoles. The seats can 

move on tracks and are adjustable for the best seating location. 

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The incorporation of the MOL experiments requires some modification of the 

basic laboratory module of Concept II. The pointing and tracking scope 

requires the placement of the movable head portion externally to the laboratory 

wall and must allow for 360 degrees of rotation through a pressure seal. An 

aerodynamic fairing is required over the exposed scope during the injection 

into orbit after which it is jettisoned. A similar.provision is required for the 

periscope/star theodolite optical system. A new separate air-lock compartment of 40-inch diameter and 75-inch length is required to conduct the 

"extravehicular system" experiments and operate the remote maneuvering 

unit. This compartment reduces the repressurization gas requirements for 

performing the experiments. 

There are numerous optical ports and small sealed compartments required 

in the laboratory wall for the experiments. Each of the experiments and 

related equipment will have to be integrated structurally, mechanically, and 

electronically to meet the environmental effects imposed by the mission. 

Some items such as the parabolic and flat antennae are placed within the 

unpressurized section of the laboratory module and are deployed for operation. 

The Apollo/MOL configuration based on the NAA Concept II geometry was laid 

out without regard to the payload weight limitation of the Saturn IB booster 

(Section 2. 2). The resulting laboratory module exceeds the payload. weight 

limitation of the Saturn and is considerably in excess of the volume requirements for the MOL mission for two crew members. It is representative, 

however, of the type of laboratory that could be used for a larger crew -when 

both the mission requirements .and payload capabilities are increased. 

The NAA Concept II was designed for docking of the combined command and 

service modules on the laboratory module to effect crew transfer after a 

repositioning maneuver and requires making two umbilical connections 

adjacent to the crew transfer tunnel. This means that the supply lines ,must 

be carried from the service module through the Apollo command module and 

to the umbilicals. The making of umbilical connections in addition to the 

docking procedure unduly complicates the system design and adds certain 

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weight penalties. The .NAA Concept II requires the docking of a second 

Apollo command and service module before the first command/service 

modules can separate for return of crew members from orbit. If this is 

not done, the laboratory module would stop functioning because the environmental control and life support systems supplied by the service module would 

not be available. 

It is clear that the Apollo/MOL configuration based on the NAA Concept II 

is best suited for extended missions requiring resupply and crew rotation. 

• 

1.3.3 .Apollo/MOL/Saturn II3 Configuration Modified 

NAA Concept II 

The Apollo/MOL configuration, using an arrangement similar to the NAA 

Extended Mission Apollo Concept. II, but modified for two crew members 

instead of six by reducing the laboratory volume, is shown in Figure 1-10 

(Drawing ES-0152-002). 

The modified configuration has an Apollo command module modified for an 

earth orbit and re-entry and a modified service module attached to a single 

compartment laboratory module. The heat shield .of the Apollo vehicle is 

modified for an earth orbit re-entry instead of a lunar orbit re-entry and 

consequently is nearly 600 pounds lighter. A retro-package of six XM-85 

solid propellant motors, as in the previous configuration, is used for effecting 

re-entry. The service module was shortened by 62 inches and was modified 

to incorporate the same subsystems as before. 

The single laboratory compartment is arranged to combine areas for rest, 

feeding, personal hygiene, recreation, and the MOL experiments. The 

experiments and instruments are located so that related experiment displays 

could be observed from one position. The aft docking cone and air-lock was 

replaced by an air-lock compartment 40 inches in diameter and 75 inches long 

for conducting the "extravehicular" experiments and for operating the remote 

maneuvering unit. 

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A single hand rail is provided around the interior of the compartment to 

facilitate crew motion in the weightless environment. Again, a track for a 

restraining mechanism is provided in the "floor" of the compartment and 

adjustable seats on tracks are provided with restraining straps. 

The pointing and tracking scope, the periscope/star theodolite, the deployable 

antennae and various other experiments were integrated into the laboratory 

similarly to the larger laboratory shown in Figure 1-9. 

The modified configuration is designed for a single docked command/service 

module which supplies the environmental control and life support. The 

laboratory again stops functioning after separation because of its dependence 

on the command/service module for support. There is some capability for 

growth and mission extension but it requires enlarging the service module to 

carry more supplies or the use of a resupply technique and some modification 

such as adding the second docking cone and air-lock. 

In the case of both Concept II configurations, to achieve the final circularization of the orbit, the following sequence is performed during the 45-minute 

coast period after the Saturn IVB stage burnout: 

a. Separation of the command and service modules as a unit from the 

laboratory module 

b. Repositioning maneuver (180-degree turn around) of the command/ 

service module unit 

c. Docking of command/service module unit on the laboratory module and 

making two umbilical connections 

d. Separation of the Saturn IVB stage from the laboratory module 

e. Complete 180-degree turn around of all three combined stages 

f. Firing of the 10. 5 K LEM engine at apogee of orbit 

1. 3. 4 Apollo /MOL/Saturn IB Designed. Specifically to 

MOL Requirements 

A preliminary layout of an Apollo/MOL configuration with all the MOL 

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i 

experiments integrated into a laboratory module is shown in Figure 1-11. 

The Apollo /MOL configuration consists of the Apollo command module with 

a short adaptor attached to the combined laboratory/service module. This 

configuration has .a single laboratory- compartment of approximately the same 

size and arrangement as the modified Concept II configuration (Figure 1-10). 

The "extravehicular" air-lock is relocated to allow for the placement of the 

LEM 10. 5 K rocket engine on the center directly aft of the laboratory module 

bulkhead. 

The separate service module required for both the NAA Concept II and the 

NAA optimized. Concept II is eliminated and is combined with the laboratory 

module. The subsystems, the propulsion system, and the experiments are 

identical with the NAA modified Concept II but are rearranged and relocated 

within the space external to -the laboratory to suit the aft mounted propulsion 

system. This arrangement allows the rocket motor to be operated to 

circularize into the final orbit at apogee without requiring the repositioning 

maneuver of the command and service modules during the coast period as in 

the case of both. Concept II configurations. 

The Apollo /MOL configuration with the subsystems integrated into .the laboratory module requires only one repositioning,maneuver after -attaining final 

orbit with no definite time limit imposed for performing the repositioning 

maneuver. Since the subsystems are-rearranged, the umbilical connections 

to provide environmental control and life support to the laboratory as used in 

Concept II are unnecessary. The crew can shorten the mission duration at 

any time and separate from the laboratory module without interrupting the 

functioning of the laboratory since the subsystems are self-contained. 

This Apollo/MOL/Saturn IB configuration has a growth capability to 

accommodate larger crews and to extend mission duration. 

1. 4 PERTURBATION OF THE APOLLO /MOL CONFIGURATION 

1: 4: 1 Configuration. Arrangements and Laboratory Size 

The laboratory volume can be varied to suit:• the mission requirements, the 

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crew size, the MOL experiment requirements, and the payload capability of 

the launch vehicle. The arrangement of the Apollo /MOL configuration can 

also be varied as shown in Figures 1-1, 1-2, and 1-3. Figure 1-4 shows the 

same arrangement as Figure 1-3, but the laboratory size is decreased to a 

cylindrical habitable compartment 150 inches internal diameter and of varying 

height, resulting in a volume of 1230 cubic feet. 

1.4.2 Methods of Crew Transfer 

The method of crew transfer has been considered in terms of a docking 

procedure and entry through the use of air-locks.for which the Apollo vehicle 

was designed .(References 1-3 and 1-4). Other methods of crew transfer are 

shown in Figures 1-5 through 1-8 as perturbations of Figure 1-4. Figure 1-5 

shows a modified Apollo capsule and retro package with the central portion of 

the heat shield hinged out of the way to allow fora pressurized crew transfer 

tunnel to connect the laboratory with the Apollo capsule. This configuration 

would require an. air-lock hatch both in the floor of the capsule and possibly 

in the laboratory bulkhead, which would be sealed before separation. The 

sealing of the hinged heat shield would be a technological problem to be solved 

before safe re-entry could be effected. The laboratory would need an air-lock 

hatch only, in case of extended missions requiring resupply or crew rotation. 

Figure 1-6 shows a modified. Apollo capsule with a pressurized inflatable 

tunnel which connects the side hatch of the Apollo vehicle with the side wall 

of the laboratory module. A hatch in each vehicle, at the entrance•to the 

tunnel, would be required to make a satisfactory design. The laboratory 

could. be modified .for extended mission and crew rotation by adding a central 

docking cone and an air-lock hatch. It is doubtful that the pressurized tunnel 

can be suitable for crew transfer after the separation of the first Apollo capsule 

unless a rear docking technique is combined with a method of securing the 

tunnel to the side hatch of subsequent Apollo capsules. 

Figure 1-7 shows a modified Apollo capsule supported on trunnions from the 

laboratory module. Through the use of a flexible control wire or -by radio link, 

the trunnion mechanism can be operated.to separate the Apollo capsule from 

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•4• • • ■ • . . 

• the launch supports of the laboratory module a sufficient distance to allow the 

mechanism to rotate the Apollo capsule 180 degrees and draw it into the airlock hatch of the laboratory module. Crew transfer is effected_through the 

air-lock hatches. 

The mechanism can be precisely controlled and programmed to operate at the 

pressing of a single button. The addition of the trunnion mechanism would be 

a .weight penalty but would insure a safe and a highly reliable system for 

docking and crew transfer. Weight could be saved by designing the mechanism 

to accomplish the rotation of the Apollo capsule at a slow rate and by cushioning against resilient material on stopping and on drawing together. 

Figure 1-8 shows a modified Apollo capsule in which crew transfer is effected 

by the "extravehicular" system. The crew leaving the Apollo capsule moves 

by means of hand holds. to the laboratory and enters through a tunnel serving 

as an air-lock. 

1.4.3 Extension of Mission Duration • Extending the mission beyond the 30-day requirement involves both the 

physiological and endurance capability of the crew members and the quantity 

of supplies aboard the laboratory or service module. The crew members 

could likely endure being confined in a laboratory provided .their duty cycle, 

recreation period, and rest periods are well balanced. The unknown aspect 

is the physiological and psychological effect of weightlessness for extended 

periods. It may be that even 30 days is too long a period to remain weightless, 

thus requiring crew rotation or periods of "artificial gravity" treatment to 

rejuvenate the crew. 

The increased supply requirements can be handled by making the life support, 

power supply, attitude control systems, etc., of sufficient capacity to be ample 

for the duration of the mission or a resupply technique could also be used. 

By using the resupply technique, the original launch weight could remain the 

same as for shorter duration missions, but launching, rendezvous and resupply 

procedures would be required for the mission duration. 

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One aspect to be considered in extending mission duration is the effect of the 

"hard-vacuum" on the parachute recovery system of the. Apollo vehicle. Since 

the parachutes consist of nylon fibers and certain plasticisers, evaporation 

effects on these constituents could result in recovery system failure. It may 

be necessary for extended missions to repackage the entire recovery system 

within pressure tight cannisters to overcome vacuum effects, to redesign the 

entire recovery system, or to use some other material less affected by vacuum. 

1. 4. 4 Addition of Artificial Gravity Capability 

There are many schemes for inducing artificial gravity, but they all require 

some means of rotating the crew or the entire station. The laboratory module 

could be modified to contain a large centrifuge below the "floor" of the 

laboratory or without much modification, a small centrifuge can be used in 

the laboratory which could be stored out of the way until needed. The type of 

machine and the requirements for artificial gravity have not been established 

and hence will not be considered until the need for artificial gravity is clarified. 

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1 . 5 CONCLUSIONS 

The following conclusions have been_derived from this study: 

1. The Apollo/MOL/Titan IIIC configuration is limited both in volumetric 

and payload weight capability in carrying experiments. Experiments P-1, 

P-3, P-6, P-11, and P-12 must be carried entirely within the capsule; 

portions of experiments P-2, P-4, P-7, P-8, and P-10 maybe carried in 

the unpressurized area. Some value, however, can be attributed to this 

configuration because the astronauts do not have to leave the capsule to 

perform the experiments.. 

2. The Apollo /MOL/Titan IIIC configuration, because of the limited weight 

and space available for experiments, would require more flights than the 

Apollo/MOL/Saturn 1B configuration to cover the range of .MOL experiments. 

3. All of the Apollo /MOL/Saturn 1B configurations appear satisfactory in 

volume capability for containing all of the MOL experiments in one flight. 

4. The growth potential and mission duration capability of the Apollo/MOL 

is dependent on the launch vehicle payload capability and on updating of the 

information on space effects. 

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LINCLAS166mana 

• 

Figure 1-1. Minimum Modification Command Module, Service 

Module and Laboratory Module Using NAA Concept II 

Geometry 

Figure 1-2. Command Module, Service. Module and Laboratory 

Module (Modified NAA Concept II) 

• 

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INNIMMIDMIL 

Figure 1-3. Command Module and Combined Laboratory/Service 

Module with the Propulsion System Aft 

Figure 1-4. Command Module and Combined Laboratory/Service 

Module Reduced in Volume with the Propulsion 

System Aft.. 

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Figure 1-5. Crew Transfer Effected by a Tunnel through Command 

Module Heat Shield 

— 

' 

C, Mt 

Figure 1-6. Crew Transfer Effected by an Inflatable Pressurized 

Tunnel from the Command Module to the Laboratory 

Module 

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Figure 1-7. Crew Transfer Effected by the Repositioning of the 

Command Module Using a Mechanical Trunnion • 

Figure 1-8. Crew Transfer Effected by Extravehicular Excursion 

of Crew Members through Air Locks 

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• 

REFERENCES 

1-1 "MOL Preliminary Experiments Data Book," Aerospace Report, 

TOR-269(4107-35)-i, 30 June 1964 (Secret). 

1-2 "Extended Mission Apollo Study," North Am-erican Aviation Inc. , 

Report SID63-1370, November 1963 (Confidential). 

1-3 NASA Report, "Project Apollo - Spacecraft Description," i September 

1963 (Confidential). 

1-4 "Modified Apollo Logistics Spacecraft," North American Aviation Inc., 

Final Review, Report No. SID63-1459, 3 December 1963 (Confidential). 

1-31 

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• 

SECTION 2 

WEIGHTS 

C. Stewart 

SUMMARY 

Preliminary weight estimates are computed for a space station 

system using an Apollo re-entry vehicle, a service module and an experimental 

laboratory-that would orbit at 160 nautical miles. 

These weight estimates enable the weight. available for experiments 

to be determined for various space station configurations and mission durations, 

and for two launch vehicles: the Titan IIIC and the Saturn M. 

The results show -that, for the space station configuration launched by 

the Titan IIIC, there. is an experiment weight capability of approximately 2280 

pounds, and for the various space station -configuratinns-.1aunched by the 

Saturn IB this experiment weight capability ranged from approximately 180 

pounds to 6255 pounds. All configurations include an additional velocity increment of 200 feet per second that is available for experiments. 

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CONTENTS 

2. WEIGHTS 

2. 1 INTRODUCTION 

2.2 PRELIMINARY WEIGHT STUDY FOR APOLLO/MOL 

USED WITH TITAN IIIC 

Page 

2-5 

2-5 

2-5 

2. 2. 1 Vehicle Description 2-5 

2. 2. 2 Basis for Weight Analysis 2-6 

2. 2. 3 Results 2-6 

2. 2. 4 Conclusions 2-7 

2. 3 PRELIMINARY WEIGHT STUDY FOR APOLLO/MOL 

USED WITH SATURN IB 2-7 

2. 3. 1 System Description 2-7 

2. 3. 2 Experimental Payload Definition 2-8 

2. 3. 3 Mass Property Data 2-8 

2. 3. 4 Basis for Weight Analysis 2-9 

2. 3. 5 Results 2-9 

Z. 3. 6 Conclusions 2-10 

REFERENCES 2-23 

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4.61.ASSak 1 i 

TABLES 

2- 1 

2-2 

2-3 

2-4 

2-5 

2-6 

Weight Summary Apollo/MOL Configuration for 

Titan IIIC Launch Vehicle 

Weight Statement Apollo Crew Module Apollo/MOL 

Configuration for Titan MC 

Weight Statement Service and Equipment Module 

Apollo /MOL Configuration for Titan IIIC 

Moment of Inertia for Minimum Modification Vehicle 

Mass Properties 

Estimated Weight Statements for Apollo /MOL on 

Saturn 1B 

Weight Available for Experiments and/or Maneuver 

Page 

2-11 

2-12 

2-13 

2-14 

2- 15 

2-19 
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• 

2. WEIGHTS 

2.1 INTRODUCTION 

Preliminary weight estimates are made for Apollo/MOL space station configurations using, as a launch vehicle, the Titan IIIC in one case and the Saturn IB 

in various other configurations. These weight estimates present -the weight 

available for experiments, in the various space station systems, all of which 

orbit at 160 nautical miles. The basis for these estimates is data from North 

American Aviation supplemented by analytical weight estimating techniques. 

2.2 PRELIMINARY WEIGHT STUDY FOR APOLLO/MOL USED WITH 

TITAN IIIC 

Preliminary weight estimates were computed to assess the experimental payload capability of a MOL system using a modified. Apollo crew module and a 

service module. The mission postulated for this study was an integral launch, 

east from ETR into a 160-nautical mile orbit using the Titan IIIC launch 

vehicle. (See page 7-62, Section 7. ) 

2.2.1 Vehicle Description 

2.2.1.1 Apollo Crew Module 

The Apollo re-entry vehicle.considered-for this study was based on the standard 

NASA vehicle modified to meet the following requirements: 

a. The crew size was reduced from three to two men. 

b. The vehicle equipment and structural heat shield were modified 

for the earth orbit mission. 

c. The mission duration was set at 30 days. 

d. A de-orbit system was required in the re-entry vehicle. 

• • 2-5 

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2.2.1.2 Service Module 

This vehicle was a new design and its size. was constrained by the booster 

capability and the..weight of the Apollo re-entry vehicle. .It was. designed to 

contain the service equipment required for the 30-day mission as well as -the 

volume of the proposed experiments. 

2.2.2 Basis for Weight Analysis 

The weight of the crew module ,(Apollo), was based on data contained in 

Reference •2-1 supplemented by more recent weight data (Reference 2-3) 

obtained from NASA, Cape Kennedy on the Block 2-vehicle system. This 

supplementary data listed a weight increase of approximately 1300 pounds 

over that quoted in the November 1963 monthly weight status, Reference 2-1. 

This increase was due to the addition of docking requirements, an exit heat 

shield, and a new maintenance concept. The weight increments for the 

reduction in heat shield for the low orbit mission and the addition of a de-orbit 

system were based on data shown in Reference 2-2. 

The weight statement includes a 2000-pound growth contingency in order that 

a payload comparison can be made between this configuration and the Gemini/ 

MOL design. 

The weights for the service module section were based upon the dimensional 

data as depicted in Section 1 of this report and analytical weight estimating. 

methods. The weights for the equipment in the service module were based 

on data appearing in Reference 2-1, modified to include a fuel cell power 

supply capable of operating for 30 days. 

2.2.3 Results 

A weight summary showing the weight of the Apollo/MOL configuration is 

presented in Table 2-1. 

Weight statements for the Apollo crew module and the service module are 

shown in Tables 2-2 and 2-3. 

2-6 • 

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• 

. 2. 2. 4. Conclusions 

The study. indicated an experimental payload capability of :2280 pounds for the 

Apollo/MOL/Titan IIIC launch configuration. This weight compared with 3432 

pounds of experiments carried on a Gemini B/MOL/Titan IIIC configuration for 

a 30-day mission. These results suggest that more flights would be required 

to accomplish the same program covered by the Gemini B/MOL, or . that a 

reduced program of experimentation could be performed. with. the Apollo/MOL 

system. 

The weights derived for this study were based on preliminary data and should 

be regarded as approximate. 

2. 3 ,PRELIMINARY WEIGHT STUDY FOR APOLLO /MOL USED WITH 

SATURN IB 

A preliminary weight estimate was made of three MOL system configurations 

designed around the present Apollo command module and boosted by the Saturn 

IB launch vehicle. The study included perturbations of the three baseline 

configurations for mission durations of 30, 60, and 120 days using alternate 

power supply system concepts. A nominal experimental payload weight of 

4395 pounds was used throughout the study. 

2. 3. 1 System Description 

The three baseline configurations are described in the following brief summaries. 

These studies were made for mission durations of 30, 60, and 120 days. 

2. 3. 1. 1 Minimum Modification Apollo /MOL 

The minimum modification configuration is composed of the present Apollo 

command module modified for a two-man crew, the present service module 

with a new propulsion system, and a la:boratOry, modiile. This system is shown 

in Figure 1-9, Section 1. 

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Z. 3. 1. 2 Modified Apollo/MOL 

The modified configuration is similar except the command, module heat shield 

is reduced to earth orbit requirements, the service module is shortened, a 

smaller laboratory module is utilized, and the electrical power system is 

changed. Studies were made on the modified version, designed for mission 

durations of 30 to. 120 days, comparing current technology fuel cells and a 

LiOH carbon dioxide removal system with a system using solar panels and 

molecular. sieves. An outline drawing showing this configuration is contained 

in Figure 1-10. 

2. 3. 1. 3 Alternate Configuration 

This configuration is similar to the modified version except that the propulsion 

system is moved from the service module and mounted on the aft bulkhead of a 

combined laboratory and service module. This configuration is shown in 

Figure 1-11. 

2. 3. 2 Experimental Payload Definition 

The experimental payload of 4395 pounds was established for this study by .the 

Program Office and is based on estimates made by the Weight Prediction Section 

for the Gemini/MOL Program. This .nominal payload weight was used so that 

a relative comparison could be made among .the various configurations studied. 

2. 3. 3 Mass Property Data 

The mass property data shown in Table 2-4 include an approximate estimate of 

the inertia values made to provide basic data for control system analysis. The 

longitudinal centers of gravity for vehicle conditions 1 and 2, shown in the 

table, were measured forward of the space station to S-IVB booster interface. 

These values for the vehicle in condition 3 were measured aft of the crew module 

nose leading edge. The reason for this change in reference datum between 

vehicle conditions 1, 2, and 3 is due to the 180° rotation of the Apollo re-entry 

2-8 

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vehicle during the docking maneuver. The vertical centers of gravity were 

measured from the vehicle centerline and in all cases were below this centerline reference datum. 

2.3.4 Basis for Weight Analysis 

The weights for this study were based on configuration drawings shown in 

Section 1. The weight used for the Apollo Command Module. was based on 

Reference 2-1, the North. American Aviation monthly weight status. of 

November, 1963, supplemented by more recent weight data (Reference 2-3) 

obtained from NASA on the Block 2 vehicle system. The supplemental weight 

data included approximately 1300 pounds increase in weight for docking 

capability, an exit heat shield and a new maintenance concept. In this study 

the Apollo re-entry heat shield weight was reduced by 600 pounds for the low 

orbit mission and a de-orbit retro rocket system was added. These data were 

based on the North American Aviation XMAS study, Reference 2-2. The space 

station system service and supply section weights were derived for mission 

periods of 30,. 60, and 120 days considering fuel cell and solar array power 

supplies as alternatives. (A weight allowance of 650 pounds was included for 

the effective weight of the abort tower. ). A weight contingency of 2000 pounds 

based on that used in the Gemini/MOL weight studies was included to facilitate 

the comparison to be made between these studies and the Gemini/MOL studies. 

2.3.5 Results 

The results of this weight study are presented in Table 2-5 as estimated weight 

statements for the various configurations. Included for comparative purposes 

were the reference North American Aviation configurations for 3 man 14-day 

and 3 man 120-day missions. This table includes a .fixed experiments weight 

of 4395 pounds and indicates the total weights of all the configurations for this 

experimental weight. Table 2-6 is prepared to show the available experimental 

weight capability for a Saturn IB booster capability of 33,460 pounds. Also 

included for all configurations is a velocity increment of 200 feet per second, 

available for the experiments. 

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2.3.6 Conclusions 

Configurations'D, 0, H, and K are capable of carrying the proposed payload 

of 4395 pounds. 

Configurations C, D, and K may be compared on a mission basis with Gemini/ 

MOL; of these, D and K have 5860 and 5460 pounds available for experiments • 

(after considerable modification to the Apollo system) versus 3432 for the 

Gemini B./MOL/Titan IIIC; configuration C (relatively unmodified. Apollo 

system) can only carry .180 pounds of experiments. 

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Table 2-1 

Weight Summary 

Apollo /MOL Configuration for Titan MC Launch Vehicle 

Item Weight (lb) 

Crew Module 12,300 

Apollo (including 2 men and retro section) 

Experimental payload 

11,070 

1,230* 

Service Module 6,400 

Structure and equipment 5,350 

Experimental payload 1,050* 

Payload Contingency 2,000 

Total Booster Payload 20,700 

*

These values differ very slightly from those presented in Volume I 

because they are based on later information. 

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Table 2-2 

Weight Statement 

Apollo Crew Module 

Apollo /MOL Configuration for Titan IIIC 

Item Weight (lb) 

Structure 4,000 

Crew Systems 210 

Communications 302 

Instruments 132 

Control and Displays 160 

Guidance and Navigation 282 

Stability and Control -220 

Reaction Control 323 

Electrical Power 480 

Environmental Control 300 

Transfer Provisions 210 

Earth Landing System 753 

Weight Growth since contract 1,300 

CREW MODULE - EMPTY WEIGHT 8,672 

Crew Systems 705 

Reaction Control 270 

Environmental Control 153 

Experimental Payload 1,230 

CREW MODULE - RE-ENTRY WEIGHT 11, 030 

Retro-Rockets 936 

Retro Package Structure 102 

_External Power Supply 232 

CREW MODULE - DE-ORBIT WEIGHT 12, 300 

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Table 2-3 

Weight Statement 

Service and Equipment Module 

Apollo /MOL Configuration for Titan MC 

Item W eigh't (1b) 

Structure 1,550 

Electronics ' 177 

Reaction Controls, Fixed 400 

Electrical Power, Fixed 1, 020 

Environmental Control, Fixed 150 

Experimental Payload 1,050 

SERVICE MODULE - EMPTY WEIGHT 4,34.7 

Reaction Control - Useful Load 500 

Electrical Power - Useful Load 1,160 

Environmental Control 393 

SERVICE MODULE - LOADED WEIGHT 6,400 

• 2-13 

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Table 2-4 

Moment of Inertia for Minimum Modification Vehicle 

Mass Properties 

Weight Center of Gravity Moment of Ipertia 

Vehicle Condition Pounds Inches Slug/ft 

Longitudinal Vertical Pitch Yaw Roll 

1. Apollo/MOL - 35, 695 293* 2. 4*** 216, 785 216, 285 33, 200 

After Docking 

2. Apollo/MOL - 34, 275 285* 2. 5*** 204, 010 203, 650 33, 195 

Less Mission 

Propellant 

3. Apollo - 21, 540 150** 4. 0*** 30, 305 29, 840 10, 200 

At. Docking 

Maneuver 

Measured forward of space station. to S-IVB/booster interface. 

Measured aft of the command module nose leading edge. 

Measured from vehicle longitudinal centerline. 

2-14 

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4111111111111111111111 

REFERENCES 

2-1. Monthly Weight and Balance Report for Apollo Spacecraft (C), S. I. D. 

62.99.21, dated 1 November 1963, North American Aviation, Inc. 

2-2. Extended Mission Apollo Study, S. I. D. 63.1370, dated November 

1963, (C), North American Aviation, Inc. 

2-3 Informal Weight. Data on Apollo Block 2 Vehicle, NASA, F. C. Strible, 

dated 3 August 1964 

2-2.1 

A 

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UNCLASSIFIED 

SECTION 3 

EXPERIMENTS INTEGRATION 

J. J. Fastiggi 

SUMMARY 

The Apollo/MOL/ Saturn 1B concept appears to be roughly competitive with Gemini/MOL. The Apollo/MOL/Titan IIIC is feasible but clearly not 

competitive with Gemini/MOL. Depending on the configuration utilized, the 

MOL mission objectives may be met with from three to nine launches of the 

Saturn 1B. Some of the Apollo/MOL/Saturn 1B configurations are capable of 

carrying all the Air Force primary experiments and a few of these configurations 

are capable of carrying additional experiments. Crew duty cycles may pose 

problems when all experiments are to be performed on a two-man 30-day mission. 

However, the available data on experiment duty cycles was found to be incomplete 

at this time, thereby prohibiting a complete evaluation of this problem. 

3-1 

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t➢ NCLASSIFIED 

CONTENTS 

Page 

3. EXPERIMENTS INTEGRATION 3-3 

3. i INTRODUCTION 3-3 

3.2 DISCUSSION 3-3 

3. 2. i Configurations 3-3 

3. 2. 2 Experiments 3-4 

3. 2. 3 Allocation Constraints and Ground Rules 3-5 

3.2,4 Experirmnt Allocation 3-6 

3:3 CONCLUSIONS 3-8 

3.4 RECOMMENDATIONS 3-8 

REFERENCES 3-23 

3-2 

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3.• EXPERIMENTS INTEGRATION 

3. 1 INTRODUCTION 

It is the primary purpose of this study to investigate the feasibility of utilizing 

the Apollo command module, Apollo subsystems, and Apollo hardware as a 

vehicle system capable of performing the MOL mission objectives. To meet 

this end, from an experiment integration standpoint, the number of flights of 

a boost vehicle - MOL combination required to perform the mission objectives 

has been determined. The boost vehicles considered in this study were the 

Titan IIIC and Saturn 1B. 

Since weight is probably the most significant parameter, experiments were 

allotted to each flight primarily on a weight basis. In all but a few cases, 

the laboratory and experiment configurations, which were generated to determine the minimum number of flights required, do not exceed the anticipated 

booster capability. Because of the limited time available and the preliminary 

design nature of the equipment, it was not possible to include all design parameters and possible configurations. 

In addition to the weight problem, crew duty cycles (defined as the time required 

daily to perform experiments) must be considered. However, crew duty cycles 

are mainly supposition at this time and many experiments have yet to be 

completely defined. Therefore, only a cursory investigation of this problem 

has been performed. 

It should be emphasized that the detailed data presented in Table 3-1 are based 

on preliminary and incomplete information; and the summaries presented in 

Table 3-6 are subject to the same limitations. 

3.2 DISCUSSION 

3. 2. 1 Configurations 

The configurations used for this study are described in Section 1. Included 

in these are Apollo /MOL/Titan IIIC, Apollo/MOL NAA Concept II (minimum 

modification), Apollo/MOL NAA Concept II (modified), and Apollo/MOL Aerospace 

UNCLASSIFIED 3-3 

4111N1111111111111111M 

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ill111.111.111110 

alternate configuration. All the basic configurations listed above utilize LiOH 

to remove CO2' and fuel cells as a power source for a 30-day mission. Perturbations were introduced into the Apollo/MOL NAA Concept II (modified) 

configuration with the substitution of molecular sieves in lieu of the LiOH system, 

and solar cells in lieu of fuel cells. 

3. 2. 2 Experiments 

In order to generate a realistic comparison with the Gemini B/MOL system, 

only those experiments considered in previous integration studies have been 

included in this study. These include the following Air Force primary experiments: 

P-1 Acquisition and Tracking of Ground Targets 

P-2 

P-3 Direct Viewing of Ground and Sea Targets 

P-4 Electromagnetic Signal Detection 

P-5 Eliminated 

P-6 Extravehicular Activities 

P-7 

P- 8 Autonomous Navigation 

P- 9 Deleted 

P- 10 Multiband Spectral Observation 

P- 11 General Human Performance in Space 

P- 12 Biomedical Experiments 

No attempt has been made to integrate secondary experiments although some 

configurations provide weight allowances for additional experiments. Table 3-1 

provides a complete component weight breakdown for all experiments. Data 

in Table 3-1 were compiled from previous Gemini B/MOL studies and Ref. 3-1. 

3-4 

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3. 2. 3 Allocation Constraints and. Ground Rules 

All experiments and experimental components possess characteristics which 

influence the allocation method. These characteristics usually impose a 

certain amount of constraint on where, how, and when the experiment may be 

allocated. Constraints may be imposed by some or all of the following 

characteristics: 

1. Weight 

2. Size 

3. Volume 

4. Propulsion Requirements 

5. Duty Cycle Requirements 

6. Power Requirements 

7. Environmental Criteria 

8. Directional Constraints 

9. Aperture Requirements 

10. Mounting and Location Requirements 

11. Heat Output 

12. Stability Limits 

13. Maintenance Requirements 

14. Spare Part Requirements 

15. Pyrotechnic Hazards 

In addition to basing the experiment allocation on the above constraints, 

certain ground rules must be followed. These ground rules are determined 

by Air Force mission requirements and in some cases by common sense. 

They are as follows: 

1. Each experiment is allocated to at least two flights in order to provide 

a backup experiment in the event of first flight failure. 

2. Experiments P-11 and P-12 are to be allotted to all flights since they 

are concerned with human performance and as large a population of data 

points as possible should be examined in order to produce meaningful data. 

• 

3-5 

UNCLASSIFIL 

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18, U.S.C., Section 793 and 794, the transmission or revelation of which in any manner to an unauthorized person is prohibited by law. 

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3. Secondary experiments as such are not considered at this time due to 

payload restrictions. 

4. Experiments utilizing common components are allocated to the same 

flight when possible. 

5. On-orbit propulsion, which is required for Experiments P-2 and P-8, 

is provided when total payload weight permits. 

6. The Apollo power system provides the basic power needs of the 

experiments. 

7. The Apollo environmental control system is utilized to cool the experiments except on Experiment P-10 where a cryogenic system is provided. 

8. The Apollo telemetry system is utilized as part of the experiment 

telemetry system. 

9. The Apollo radio links are used by the experiments for oral communication with ground systems and the astronauts performing extravehicular 

activities. 

!O. For the Apollo/MOL/Titan IIIC system, all equipment which does not 

require access or a conditioned environment is located in the unpressurized 

service module. 

11. The first flight in any flight series is reserved for systems checkout. 

3. 2. 4 Experiment Allocation 

Using the above constraints and ground rules, an experiment allocation was 

performed. All the ground rules were adhered to, but it was not considered 

feasible to attempt to satisfy all the listed constraints at this time. The first 

3-6 

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• 

six constraints were considered as the most important with major emphasis 

placed on the first three. Except for the Apollo/Titan IIIC configuration, size 

and volume presented no problems. Weight is, probably the most significant 

parameter is this study. The experiments are currently in a preliminary 

design or conceptual stage, thereby prohibiting the acquisition of good, useful 

data on the remaining constraints. Therefore, this experimental allocation 

study was primarily based on weight. 

In order to perform an experiment allocation on a weight basis, the boost 

vehicle payload capability and the space station weight less experiments and 

experimental propulsion must be known. These weights are obtained in 

Section 2. With the weights known, it is a relatively simple, although tedious 

procedure to allocate experiments to flights. The experiment allocations for 

each configuration considered feasible are shown in Tables 3-2 through 3-5. 

When a particular configuration indicated that only the two biomedical and one 

other experiment may be flown, the configuration was considered impractical 

Table 3-6 presents a complete allocation summary, condensing the information contained in Tables 3-2 through 3-5. The boost vehicle payload in-orbit 

capability is cross tabulated with the various concepts and the required 

number of flights indicated in their appropriate positions. This table presents 

a number of configurations which are impractical at the lower booster capability 

and a number of configurations which are capable of performing the mission 

with three flights. Configurations which would allow the inclusion of a third 

crew member are so noted in Table 3-5. 

In the event the mission is planned for the minimum number of flights (three), 

the crew may be duty-cycle limited. Section 4 of this report indicates the 

two-man Apollo /MOL crew to have 11 man-hours per day in which to perform 

experiments. The times required to perform experiments were taken from 

Reference 3-1 and are tabulated in Table 3-7. Since the experiments are not 

all. completely defined, the times referred to from Reference 3.-1 are by no 

means exact, and the times obtained from Section 4 are also inexact. 

Therefore, no firm commitment can be stated regarding duty cycles. Based on 

• 

UNCLASSIFIED 

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

19, U.S.C., Section 793 and 794, the transmission or revelation of which in any mono*, to an unauthorized person is prohibited by law. 
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available information, however, it appears that duty cycles are marginal for 

an all-experiment flight. No problems are expected on flights which do not 

carry the full complement of experiments or which provide three crew members. 

3.3 CONCLUSIONS 

It should be noted that this is a "first cut" at MOL experiments integration for 

the Apollo/MOL concept and has been performed for the purpose of determining 

the feasibility of such a system. However, based on the preliminary and 

conceptual data available, the following conclusions may be made: 

I. The Saturn IB/Apollo/MOL is a. feasible system with the present-day 

anticipated boost vehicle payload capability. 

2. For a 30-day duration flight, the MOL mission objectives may be realized 

with half as many flights utilizing the Saturn IB/Apollo/MOL concept as with 

the Gemini B/MOL/Titan IIIC system. Thus, Apollo /MOL appears roughly 

competitive with Gemini/MOL. 

3. The Apollo/MOL/Titan IIIC is limited in both payload weight and 

volume to perform the MOL experiments. A minimum of nine flights are 

required to meet the MOL objectives of performing all experiments at 

least twice. 

• 

4. Two- and three-man flights of 120-day durations are not feasible for 

this program unless molecular sieves are employed in the life support system 

and solar cells or are utilized in the power system. 

3.4 RECOMMENDATION 

A complete integration study should include the effects of the reliability and 

the cost implications. It is therefore recommended that the data presented 

herein be weighted by the appropriate reliability and cost parameters to 

provide a more meaningful comparison to Gemini B/MOL/Titan 

Since the data utilized in this study has been based on concepts and preliminary 

experiment data, an upgraded integration study should be performed when 

more complete data becomes available. This would allow the remaining 

constraints to be included in the study. • 

UNCLASSIFIED3-8 

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Ta

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Flight Allo

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• 
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TABLE 3.7 

3-21 pms ee,sS 

. - 

• 

PROPOSED DAILY TIME. ALLOCATION FOR EXPERIMENTS 

(Based on an approximate 11 hour duty cycle) 

Experiment 

Time Req'd. 

to Conduct 

Experiment 

Frequency of 

Experiments 

.for 30 Day 

MissibP 

Day Allocated for Performance of 

Experiments During a 30 Day Mission 

1-20 21-23 24425 26-28 29&30 

P-1 0:50 20 Times 0:50 

P -2 

P-3 0:30 20 Times 0:30 

P-14. 0:30 50 Times 1:00 1:00 ,1:00 

(a) (a) (a) 

P-5 

P-6 3:00 6 Times 3:00 3:00 

P-7 

P-8 1:35 Daily 1:35 1:35 1:35 1;35 1:35 

P-9 

P-10 2:00 Daily 2:00 2:00 .2:00 2:00 2:00 

P-11 1:00 Daily 1:00 1:00 .1:00 1:00 .1:00 

P-12 3:10 Daily 3:10. 3:10 .3:10 3:10 3:10 

TOTAL 10:05 11:35 10:50 11:45 11:00 

All times are in hours and minutes 

(a) This experiment is performed twice daily, 

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REFERENCES 

3-1. "MOL Primary Experiments Data Book," TOR-269(4107-35)-1, 

Aerospace Corporation, El Segundo,' California, 30 June 1964. 

323 

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SECTION 4 

CREW ACTIVITY TIME ALLOCATIONS 

E. F. Schmidt 

F. E. Cook 

SUMMARY 

The time available for the performance of experiments is presented 

for two-, thiee-, and four-man crews. These estimates are based largely on 

a comparative evaluation of the results of several previous industry and 

Aerospace studies of crew time allocations for various manned systems. It 

is found that the time available for experiments is 11 hours for both the twoman Gemini /MOL or Apollo /MOL. 

• 4-1 

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CONTENTS 

4, CREW ACTIVITY TIME ALLOCATIONS 

Page 

4-7 

4. 1 INTRODUCTION 4- 7 

4. 2 DISCUSSION 4- 7 

4. 3 CONCLUSIONS 4-12 

REFERENCES 4- 13 

4-3 

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TABLES 

Page 

Two-Man. Crew Work Load 4-9 

Apollo/MOL Crew Work Load .. . .. . . . . . . . . . ... . . . 4-10 

Crew Work. Load 4-11 

4-5 

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ellag1111111111, 

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4. CREW ACTIVITY TIME ALLOCATIONS 

4.1 INTRODUCTION 

A brief investigation was made to determine reasonable time allocations for 

the various duties of two-, three, and four-man crews for Apollo/MOL. The 

primary objective was to determine the time available for the performance of 

experiments. The approach taken was to review the results of the various 

studies already performed in this area (References 4-1 and 4-2 and previous 

Aerospace studies) and to attempt to resolve the differences in the time allocated to various crew duties, where differences existed. Allocations were then 

established for Apollo/MOL, primarily on the basis of the existing information 

that was judged to be the best supported. 

4. 2 DISCUSSION 

Table 4-1 is a tabulation of the information extracted from the various 

references. An attempt was made to break down the time allocations to a 

common level of detail, but this was not possible in all cases. The XMAS 

allocations for exercise and recreation were combined, but the total is in 

good agreement with all of the others excepting the MORL figure. Although 

the latter is comparatively large, a degree of confidence is inspired by the 

fact that Douglas appears to have performed more detailed study and experiments in this area than have the others. However, confidence in any figuile 

can only be expressed relatively until actual flight data is obtained. The 

allocation for exercise and recreation for Apollo /MOL of six hours might 

be considered to represent a weighted average of the others, with more 

confidence having been attached to a single high allocation of MORL than 

to the several low values for the other systems. 

Large discrepancies appear in the values assigned by the various studies to 

station-keeping and maintenance. However, the functions included in these . 

categories differ from one system to another. For example, the MOL figures 

for station-keeping include monitoring time, whereas for MORL monitoring 

is considered a part of maintenance. The differences between the total times 

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for the two categories are not.large except for the XMAS values. The latter 

appear to be unreasonably low, and the allocations assigned to Apollo/MOL 

reflect confidence in the higher time estimates. A more complete breakdown 

of the latter, given in Table 4-2, supports the use of the larger number. The 

time available for experiments for the various systems is quite consistent, 

except for the XMAS values. This discrepancy may be attributed primarily 

to the low figure used for station-keeping and maintenance, as discussed 

above. 

The column headed "Gemini B/MOL" represents the values used by the MOL 

Program Office in a recent briefing. 

Since both the MORL and XMAS concepts are oriented primarily toward biomedical and behavioral experiments, an attempt was made to allocate the 

Apollo/MOL time for these experiments (P-11 and P-12) separately. The 

allocations given are based on recent discussions with the MOL Program. 

Office. The time allocated to P-11 of one man-hour per day is an average, 

and represents the performance of two one-half hour experiments every other 

day by each astronaut. The P-12 allocation is also an average value based 

on conversations with Air Force Aeromedical Division personnel. Each of 

the tests included in P-12 is not performed every day, nor are both astronauts 

sampled each day. The daily time required will, therefore, vary. 

Table 4-2 tabulates the time allocations for two-, three-, and four-man crews. 

The task breakdown is detailed to the level required to determine the allocations which increase with crew size. The estimated time requirements for 

biomedical and behavioral experiments for the three- and four-man crews 

were based on discussions with the MOL Program Office. An increase in time 

available for other experiments of nine man-hours for each additional crew 

member over the initial two is realized. 

Table 4-3 summarizes the Apollo/MOL two- and three-man work loads, and 

compares the allocations for various categories with those given by the 

Gemini/MOL Program Office. 

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111111111111111111111 aTable 4-2. Apollo/MOL Crew Work Load. 

2 3 4 

Task Astronauts Astronauts Astronauts 

Sleep 16:00 24:00 32:00 

Eat 2:00 3:00 4:00 

Personnel Hygiene 2:00 3:00 4:00 

Exercise 3:00 4:30 6:00 

Recreation 3:00 4:30 ' 6:00 

Station Keeping 5:30 6:00 6:30 

Data Management (2:40) (3:00) (3:20) 

Station Management (1:50) (2:00) (2:10) 

Orbit Keeping & Navigation (1:00) (1:00) (1:00) 

Maintenance 5:30 5:30 5:30 

Preventive & Repair (1:30) (1:30) (1:30) 

Checkout & Monitoring (4:00) (4:00) (4:00) 

Experiments 11:00 21:30 32:00 . 

Biomedical (4:10) (5:40) (7:10) 

P-11 (1:00) (1:30) (2:00) 

P-12 (3:10) (4:10) (5:10) 

Other (6:50) (15:50) (24:50) 

TOTAL 48:00 72:00 96:00 

*

All time is expressed in hours and minutes. 

4-10 

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ifidikl mai ASS 

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4.3 CONCLUSIONS• 

The time available for experiments in an Apollo /MOL mission with a two-man 

crew appears to be approximately the same as that -for Gemini /MOL. If a 

third man is added, the available time appears to be almost doubled. 

4-12 

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• 

REFERENCES 

4-1 "Report on a System Comparison and Selection Study of a 

Manned Orbital Research Laboratory, " Volume I, Appendix 

3, Douglas Missile and Space Systems Division Document 

SM-44599 

4-2 North American Aviation SID 63-1370-1, "Extended-Mission 

Apollo,Study, " Final Report, Volume 1, 24 November 1963 

4-13 • UNCLASSIFIED 

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SECTION 5. 

POWER SYSTEMS 

H. J. Killian 

SUMMARY 

A study of electrical power systems for an Apollo /MOL spacecraft has 

be6n performed and is reported herein. Objectives of the study were to (1) 

define an electrical power system for Apollo/MOL making maximum use of 

the present Apollo fuel cell hardware, (2) select and define the best electrical 

subsystem for Apollo /MOL on the basis of comparative evaluations of candidate 

power system concepts, (3) evaluate the growth potential of an Apollo/MOL 

electrical power subsystem for extended mission durations, and (4) compare 

the best Apollo/MOL electrical power subsystem with that recommended for 

(Gemini) MOL in previous studies. Apollo fuel cells, current technology fuel 

cells, solar cells and were evaluated as candidate power 

systems on the basis of a set of power system requirements formulated as a 

part of this study. It was concluded that the current technology fuel cells were 

preferable for the early (two-year subsystem availability) missions even though 

their mission duration capabilities were limited to 30 to 40 days. Follow-on 

missions could be extended in duration to 120 days by incorporation of 

Solar cells appeared to be generally 

unsatisfactory for the Apollo/MOL,mission. Power system characteristics 

for the five Apollo/MOL configurations of interest were defined assuming use 

of the recommended power system(s). 

,5-1 

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CONTENTS 

5.1 INTRODUCTION 

Page 

5-7 

5.1.1 Candidate Power Systems 5-7 

5.1.2 Scope of Comparison Effort 5-7 

5.1.3 Objectives 5-8 

5.2 DESIGN REQUIREMENTS 5-9 

5.2. 1 Electrical Power Requirements 5-9 

5.2.2 Mission Duration 5-10 

5.2.3 Orbit 5-10 

5.2.4 Orientation 5-10 

5.2.5 Life Support Integration 5-10 

5.2.6 Reliability 5- i 

5.2.7 Auxiliary Power 5-11 

5.2.8 Availability 5-11 

5.3 APOLLO FUEL CELL POWER SYSTEM 5-12 

5.3.1 Basic Fuel Cell Module Configuration 5-12 

5.3.2 Weight 5-17 

5.3.3 Reliability 5-20 

5.3.4 Volume 5-20 

5.3. 5 Costs 5-23 

5.3.6 Availability 5-23 

5.4 COMPARISON OF CANDIDATE POWER SYSTEM CONCEPTS 5-23 

5.4.1 Quantitative Comparison 5-25 

5.4.2 Qualitative Comparison 5-29 

5. 5 RECOMMENDED POWER SYSTEMS 5-30 

5.6 COMPARISON WITH GEMINI/MOL POWER SYSTEM 5-33 

5.7 CONCLUSIONS 5-34 

UNCLASSIFIED 5-2 

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CONTENTS (Continued) 

APPENDIX A: APOLLO/MOL POWER SYSTEM USING CURRENT 

Page 

TECHNOLOGY FUEL CELLS 5-35 

APPENDIX B: APOLLO/MOL POWER SYSTEM USING SOLAR CELLS . 5-41 

APPENDIX C: APOLLO/MOL POWER SYSTEM USING 

5-49 

APPENDIX D: QUALITATIVE COMPARISON OF FUEL CELLS 

AND SOLAR. CELLS FOR APOLLO/MOL 5-59 

REFERENCES 5-69 

• 

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TABLES 

5-1 

5-2 

Apollo/MOL Estimated Power Requirements 

Power System Weight Summary Using Apollo 

Page 

5-9 

Fuel Cells (pounds) 5-19 

5-3 Power System Gross Reliability Using Apollo Fuel Cells . . 5-21 

5-4 Power System Volume Using Apollo Fuel CellS 5-22 

5-5 Power System Cost Using Apollo Fuel Cells (M = Million) . . 5-24 

5-6 Quantitative Comparison of Candidate Power Systems 5-26 

5-7 Summary of Qualitative Comparison Results Between 

Ftel Cells and Solar Cells for Apollo/MOL 5-31 

5-8 Recommended Power Systems for Apollo/MOL 5-32 

A-1 Power System Weight Using Current Technology Fuel Cells. 5-37 

A-2 Power System Reliability Using Current Technology 

Fuel Cells 5-38 

A-3 Power System Volume Using Current Technology Fuel Cells 5-39 

A-4 Power System Costs Using Current Technology 

Fuel Cells (M = Million) 5-40 

B-1 Power System Weight Using Solar Coils (lb) . . . 5-43 

B-2 Power System Reliability Using Solar Cells 5-46 

B-3 Power System Volume Using Solar Cells 5-47 

B-4 Power System Costs Using Solar Cells (M = Million). . . 5-48 

C-1 Potential Fuels for Power System 5-50 

C-2 Power System Weight Using 5-52 

-C-3 Power System Reliability Using 5-56 

C-4 Power System Volume. Using . . . . . 5-57 

C-5 Power System Costs Using 5-58 

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ILLUSTRATIONS 

5-1 Apollo. Fuel Cell Module Performance Constraints 

Page 

(Pratt and Whitney Data) 5-14 

5-2 Estimated Effect of Apollo Fuel Cell Module 

Power on Module Life (Pratt and Whitney Data) 5-15 

5-3 Reactant and Tankage Specific Weight for the 

Apollo Fuel Cell 5-16 

5-4 Reliability of Apollo Fuel Cell Modules for 

30-Day Mission 5-18 

5-5 Apollo/MOL Power Systems Weights 5-27 

B-1 Solar Cell Array Weight Comparisons 5-42 

C-1 

5-55 

• 

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5. 1 INTRODUCTION 

5.1.1 Candidate Power Systems 

Much of the prior work done at Aerospace in connection with Gemini /MOL was 

directly applicable to the analysis of power systems for Apollo/MOL. In particular, 

the same basic power system concepts under consideration for Gerniiii/MOL 

were competitive for application to Apollo/MOL; viz., fuel cells, solar cells, 

and power systems. Solar dynamic power systems were not 

evaluated since they are not current state-of-the-art and are not expected to 

be operational in this decade. Each of the three power system concepts evaluated 

has special advantages and shortcomings. A fuel cell system has a low fixed 

weight, but consumes stored reactants and therefore becomes quite heavy for 

longer mission durations. Solar cell systems have a high fixed weight, but this 

weight remains substantially constant for longer duration high altitude missions. 

This left fuel cells and solar cells as the primary candidates 

which received major emphasis for the Apollo /MOL power system study. 

5.1.2 Scope of Comparison Effort 

It was evident from the onset of the power subsystem study that the Apollo fuel 

cell power subsystem would require substantial reconfiguration to adapt it to 

Apollo/MOL needs and that solar cells were very competitive with fuel cells 

for this mission. Consequently, it was realized that technical justification for a 

selection of one in preference to the other would involve extensive analysis of 

both power systems. For unmanned spacecraft above a couple of hundred 

5-7 

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nautical miles altitude, the fuel cell, because of its continuous reactant consumption, is not weight competitive with solar cells beyond about two weeks. 

However, for manned spacecraft, a fuel cell system may integrate well with the 

life support system (thereby saving weight), whereas no integration is possible 

with a solar cell system. Also, for orbit altitudes below a couple Iwndred 

nautical miles, aerodynamic drag on solar cell panels creates propulsion weight 

penalties in favor of the fuel cell system which does not interface with propulsion. The net effect is that the MOL mission is in an area of overlapping 

applicability of both fuel cells and solar cells and analyses to select between 

the two must consider the sometimes subtle involvement of many other subsystem 

areas. 

5. 1. 3 Objectives 

Specific objectives of the electrical power subsystem study were as follows: 

1. Define an electrical power system for Apollo/MOL making maximum use of 

the present Apollo fuel cell hardware. 

2: Select and define the best electrical power subsystem(s) for Apollo/MOL 

on the basis of comparative evaluations of candidate power system concepts. 

3. Evaluate the growth potential of an Apollo/MOL electrical power subsystem 

for extended mission durations. 

4. Compare the best Apollo/MOL electrical power subsystem with that 

recommended for Gemini/MOL in previous studies. 

Design requirements which formed a base for the power system analyses are 

presented below followed by the definition of a minimum modification Apollo 

fuel cell system for Apollo/MOL. A subsequent comparison of all promising 

power system concepts for Apollo/MOL leads to recommended power systems 

for the several Apollo/MOL system configurations of interest. These results 

are compared to characteristics and capabilities of the (previously recommended) 

Gemini/MOL electrical power subsystem, and over-all study conclusions are 

drawn. 

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5.2 DESIGN REQUIREMENTS 

5.2.1 Electrical Power Requirements 

Estimated Apollo/MOL electrical power requirements are shown in Table 5-1. 

These estimates are based on the continuous utilization of Apollo and all of its 

subsystems which would be retained for the Apollo/MOL mission. In addition, 

the continuous maintenance of a "shirt sleeve" environment in an attached 

laboratory module for the housing and conduct of experiments was assumed. 

Experiment power is based on estimates made in connection with Gemini/MOL 

of the average power required for all primary experiments. These power 

estimates typify expected requirements with either a two- or three-man crew. 

Table 5-1 was prepared before the final results from other subsystem studies 

were available. Consequently, the subsystem power requirements indicated 

may differ slightly from those finally determined in the subsystem studies. 

The total average power requirements shown, however, are consistent with 

similar data from numerous Aerospace Corporation and industry studies of 

comparable missions. They will therefore result in valid power system analyses 

pertaining to Apollo /MOL itself and, it is believed, in valid power system 

comparisons with (Gemini) MOL or other mission concepts. 

Table 5-1. Apollo/MOL Estimated Power Requirements. 

Watts 

Apollo: 

Average Peak 

Stabilization and control 200 210 

Environmental control and life support 380 420 

Guidance and navigation 50 500 

Control and displays 170 270 

Illumination 80 100 

Commun. and instru. (including T/M) 190 490 

Subtotal 1070 1990 

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Watts 

Laboratory Module: 

Average Peak 

Environmental control and life support 300 500 

Illumination 100 300 

Experiments 300 1000 

Subtotal 700 1800 

Total 1770 3790 

5.2.2 Mission Duration 

Mission duration was a key design parameter covering the. range 30 to 120 days. 

5.2.3 Orbit 

Circular orbit altitudes between 100 and 200 nautical miles were considered. 

An easterly launch from AMR (32. 5-degree orbit inclination) was assumed where 

necessary for power system definition or optimization. It resulted, however, 

that the power systems defined in the study were adequate for any orbit 

inclination. 

5.2.4 Orientation 

An earth-stabilized spacecraft (roll axis aligned with velocity vector and yaw 

axis aligned withlocal vertical) was assumed to exist throughout the mission. 

5.2.5 Life Support Integration 

A fuel cell system would supply more than enough water for crew needs as a 

natural by-product of its power generation. Complete water recovery was 

assumed in connection with solar cell and power systems. However, 

the weight and power penalties which would be associated with a recoverable 

water supply were not assessed against the power systems. Therefore, in the 

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• 

power supply comparisons presented in this report, fuel cells do not receive 

credit for the advantage they offer in connection with water supply. Past studies 

have shown that this credit would amount to an equivalent weight of 100 to 150 

pounds. On the other hand, no penalties were assessed against fuel cells for 

the much greater thermal load they present to the spacecraft cooling system. 

Radiator, plumbing and control hardware weights associated with fuel cell 

cooling would exceed 50 pounds. These two counteracting omissions are both 

small enough so as to not have significant quantitative impact on the power 

system comparisons presented $ierein. 

5.2.6 Reliability 

A 30-day reliability goal of 0.9995 was assumed. This number was based on a 

preliminary subsystem reliability allocation made early in the (Gemini)MOL 

program. It was used to establish levels of redundancy for various power 

system elements so as to better definitize power system weight and volume 

characteristics and to provide a common reliability base for comparison of 

the candidate power system concepts. 

5.2.7 Auxiliary Power 

Power for Apollo during the launch and prei-station activation period would be 

supplied from the MOL. Also, an emergency power capability corresponding 

to full power (without experiments, 1.5 kilowatts) for 1.5 hours or reduced 

power for a longer period was assumed as a requirement for the over-all 

Apollo/MOL spacecraft. In addition to this emergency power, an emergency 

capability would be provided to obtain power in the Apollo command module 

only for an extended period after abandonment of the MOL and during re-entry. 

An emergency is defined as a total loss of power from the regular power source. 

5.2. 8 Availability 

Man-rated and flight-ready .hardware within two years from go-ahead was 

assumed as a requirement for the initial development program consistent with 

(Gemini) MOL planning. This limits design technology to essentially current • state-of-the-art for at least the 30-day (initial) MOL. 

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5.3 APOLLO FUEL CELL POWER SYSTEM 

The Apollo fuel cell qualified for special consideration as a power source in 

Apollo /MOL because of the study objective to make maximum use of existing 

hardware. It was decided that qualification under this objective limited such 

special consideration to "essential modification" versions of the Apollo fuel 

cell. Assumptions of repackaging, life extension or other desirable but not 

essential developments in connection with the Apollo fuel cell would render it 

no more extant than other fuel cell concepts which are important competitors. 

No major performance extensions, then, beyond what is presently specified 

for the Apollo program, were permitted the Apollo fuel cell in this "existing 

hardware" category. 

5.3.1 Basic Fuel Cell Module Configuration 

An Apollo fuel cell module can continuously deliver power at any level within 

a large' range: approximately 560 watts to almost 2000 watts. In view of the 

1.77 kilowatts of average power required by Apollo/MOL, it must be determined 

whether one, two, or three modules should be operated simultaneously to 

deliver this much power. 

A major factor that influences modules requirements is the 15-day life of 

Apollo fuel cells as compared to the 30-day and longer Apollo /MOL missions 

being considered. Obviously, module wearout and replacement with spare 

modules must form the basis for module requirements. 

Selection of a module configuration was based on consideration of the following 

constraints: 

1. Delivery of 1.77 kilowatts average power while keeping the primary DC bus 

voltage within a four-ito five-volt range. 

2. Minimize over-all weight. 

3. Achieve a 30-day reliability of approximately 0.9999 for the fuel cells. 

5-12 

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A secondary goal was to maximize peak power capability although this aspect 

of power delivery can be conveniently handled by secondary (rechargeable) 

batteries fed from the fuel cell(s). 

Figure 5-1 shows how the Apollo fuel cell module operating temperatures 

are constrained (in Apollo) by voltage tolerances and other constraints. Clearly, 

the generation of higher power levels requires higher fuel cell (electrodeelectrolyte) temperatures to stay with the voltage limits. However, module 

life is almost solely dependent on temperature with the result that it varies 

with module power output (in Apollo) as shown in Figure 5-2. Assuming that 

the Apollo fuel cell meets qualification test requirements, a relative life of 1.0 

corresponds to 400 hours (^- 15 days) and higher average power outputs would 

degrade this capability substantially. Although a variety of factors such as 

relaxation of voltage requirements, restriction of minimum power to higher 

levels and addition of another cell (32 total instead of 31), different temperature 

modulation, etc., could ameliorate this power output-life relationship, it is 

still highly questionable whether a single module could, with "minimum 

modification", deliver an average power of 1.77 kilowatts for as long as 15 days. 

Aside from intrinsic module capability, simple operational considerations 

indicate the desirability of at least two modules operating at all times. If 

one module fails, then the other can handle the full load until a spare module 

can be switched onto the line (approximately one hour warm-up time required). 

Two modules operating simultaneously can handle twice the peak load that a 

single module could. Two modules could deliver 2840 watts without falling 

below 27 volts and could supply 4400 watts at 22 volts. 

Figure 5-3 shows the weight,of cryogenic reactants and their tankage for 30 days 

per kilowatt of net module output (net output = gross output - parasitic load) 

versus average module power output. This weight begins to increase below a 

5-13 

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certain power level despite improving fuel cell efficiency because of the 

increasing importance of the fixed parasitic power losses (150 watts). It is 

clear from Figure 5-3 that no benefits are obtainable in this area by operating 

three units. simultaneously. 

It was concluded that two fuel cell modules operating at all times was the 

optimum module arrangement. This means that two modules would be started, 

would operate for 15 days (until wearout) and would then be replaced by two 

new modules. Standby replacement modules would then• be added to the basic 

four modules to obtain the necessary reliability. Figure 5-4 shows 30-day 

reliability versus total number of modules. On the basis of this graph, a total 

of eight modules was selected as the minimum necessary for 30 days of operation. 

5.3.2 Weight 

Table 5-Z•gives a' tabular summary of total Apollo /MOL power system weight 

when using Apollo fuel cells. Brief comments are made below for each 

• weight item. 

Modules: Each module weighs 210 pounds. The number of modules was determined as cited previously. 

Startup Batteries: Each module (except the initial two started prior to launch) 

requires 3500 watt-hours of energy for warmup. This would require about 

43 pounds of silver-zinc battery to be able to start up each module on-orbit. 

Water Recovery: Apollo experience 

Oxygen: Approximately 0.82 pound per kilowatt-hour including a two percent 

purge allowance. 

Oxygen Storage and Supply: Tankage, insulation, valving, etc., weights varied 

from 0.24 to 0.28 times oxygen weight depending on total oxygen weight. 

Hydrogen: Based on 0.103 pound per kilowatt-hour which includes a two 

percent purge allowance. 

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taiiASSIIFED miuma

• 

Hydrogen Storage and Supply: Tankage, insulation, valving, etc., weights varied 

from 1.3 to 2.2 times hydrogen weight depending on total hydrogen weight. 

Power Conditioning, Control and Distribution: A standard set of power 

conditioning, control and distribution equipment was defined for use with all 

types of power sources considered in this study. The "set" was comprised of: 

A 1.25-kilowatt rated regulator plus two spares 73 lb 

A 1.25-kilowatt rated inverter plus two spares 111 lb 

Fault protection and automatic switching 

Instrumentation and control 306 lb 

Distribution, installation provisions, misc. 

Justification of the spares provisioning is given in the section on reliability. 

Support Structure: Weight was computed at eight percent of the combined weight 

of the fuel cells, reactants and tankage. 

Auxiliary Battery: A nominal 50-pound allowance was made for a battery plus 

a charge regulator to accommodate extreme emergency (total loss of complete 

power system) and re-entry power requirements. 

5.3.3 Reliability 

Table 5-3 shows the results of a gross reliability analysis estimating the power 

system reliability for successful completion of a 30-day mission. The purpose 

of the analysis was to allow power system weight estimates to reflect reasonable 

levels of redundancy where needed and to establish a common reliability base 

for comparison among different power system concepts. 

5.3.4 Volume 

Results of volume calculations are shown in Table 5-4. Actual volumes occupied 

by the equipment items were multiplied by a factor estimated to account for 

unusable volume associated with the installation of items in addition to the 

volume of the items themselves. 

• 

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• 

5.3.5 Costs 

Cost estimates (see Table 5.-5) were divided into development costs and "unit" 

costs where "unit" refers to a full set of flight hardware. Fuel cell costs 

were based on informal data from Pratt and Whitney and cryogenic system 

costs were based on informal data from AiResearch. Uilit costs proved to be 

relatively high because of the large number of fuel cell modules involved. The 

200,000-dollar cost per module value used is believed to be conservative (law). 

Cost estimates for the categories entitled "electrical system engineering" and 

"other development and test" (including hardware) are compatible if not 

identical to similar cost categories used with other power system cost estimates 

(in this report). Consequently, cost comparisons with other power systems 

should be reasonably valid. 

5.3.6 Availability 

Earliest possible availability for an Apollo/MOL power system using Apollo 

fuel cells was estimated to be 18 months under the reasonable assumption that 

the fuel cells themselves would be the pacing items. The basis for the 18-month 

estimate was the following milestone schedule constructed from discussions 

with Pratt and Whitney. 

Define and engineer necessary design changes 3 months 

Procure material and parts 9 

Qualification testing 3 

Production and delivery 3 

18 months 

5.4 COMPARISON OF CANDIDATE POWER SYSTEM CONCEPTS 

Current technology fuel cells,* solar cells and are alternative 

power sources to the Apollo fuel cell that were considered in detail in this study. 

"Current technology" is a term that has been adopted to refer to fuel cells 

incorporating existing technology which is quite advanced over the Apollo 

fuel cell concept. Refer to Appendix A for more discussion of this point. 

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Appendices A, B and C summarize the results of the work on Apollo /MOL 

power systems incorporating these sources. They are defined in these 

appendices in approximately the same detail as the Apollo fuel cell power 

system was defined. 

As indicated previously, system availability would be such that 

it could not be considered for the initial Apollo /MOL design. It could, 

however, be made available for follow-on Apollo /MOL missions . Because of 

this and because of its very attractive characteristics, it was compared in .a 

direct manner with the more readily available fuel cells and solar cells. A 

constraint on its selection, however,' was that it could only be recommended 

for follow-on (generally thought of as longer duration) applications. 

5.4.1 Quantitative Comparison 

Table 5-6 compares total power system weights (plus aerodynamic-dragassociated weight penalties in the case of the solar cell system), volumes 

and costs. As indicated, these data are based on a common 30-day reliability 

of about 0.9995 for each system. Figure 5-5 illustrates the weight comparison 

data. 

Fully oriented (two-degree of freedom) solar arrays were used to define the 

solar cell system on the basis of their lighter weight for almost all altitudemission duration regimes of interest. An earth-oriented spacecraft was 

assumed at all times. Substantial aerodynamic drag would be incurred by' 

solar arrays so that a propulsion penalty was associated with the solar cell 

system. The amount of this penalty was, of course, a function of the altitude 

that would be maintained by the Apollo /MOL spacecraft. Consequently, solar 

cell power system weight is shown as a function of orbit altitude. Weights of 

both the fuel cell and systems would be altitude independent since 

they do not include drag inducing parts. 

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Assessing power system weights on the basis of a relatively high reliability 

proved quite a bit more penalizing to some systems than to others. Reliability 

improvement with the solar cell and current technology fuel cells up to the 

0.9995 level resulted in modest weight increases for these systems: roughly 

100 and 300 pounds, respectively. This same change cost almost 400 pounds 

with the system and about 950 pounds with the Apollo fuel cell 

system. Although no two of these power system types would receive the same 

reliability allocation, even in identical applications, there was no satisfactory 

alternative in this study but to compare the different systems on the basis 

of equal reliabilities for all, keeping in mind what this reliability cost in 

each case. 

Volume data for the fuel cell systems reflect the, relative compactness of the 

current technology fuel cells. Total volume of reactants was essentially 

identical for both fuel cell systems. The 200-nautical mile, 30-day solar 

cell system volume is indicative of the volume of the solar cell system itself. 

Larger values for other mission conditions include appreciable amounts of 

propellant and associated hardware volume. 

Cost estimates in Table 5-6 are rough at best and probably are all optimistic. 

In essence, they indicate that all of the power systems would cost about the 

same to develop. "Unit cost" refers to the cost of a full set of flight hardware. 

Differences in unit costs among the systems are more pronounced than 

development cost differences and realistically so. An important point to 

remember when considering the cost of developing a new power source for 

follow-on missions is that, within the guidelines of this study, all parts of 

the power system other than the source would not need redeveloping. Therefore, only a portion of development cost indicated for each system in Table 5-6 

UNCLASSIFIED 5-28 

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L UNCLASSIFIED itiall11111111.1.1. 

would be incurred if that system were to replace another system for 

follow-on missions. 

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• 

development (13 million dollars) plus the initial set of flight hardware 

(1. 5 million dollars). 

It is reasonably clear from the quantitative comparison data presented 

in Table 5-6 and Figure 5-5 that the Apollo fuel cell system is not very 

competitive for Apollo /MOL application. Also, fuel cells of either type 

studied are not competitive for longer duration missions out to 120 days. 

Factoring in the availability constraint on the system leaves 

the following possibilities which remain (at this point) attractive. 

1. Two-year availability (emphasis on 30-day mission capability) 

current technology fuel cells 

solar cells 

2. Three- to four-year availability (emphasis on growth to 120-day mission 

capability) 

- solar cells 

5.4.2 Qualitative Comparison 

A. great many factors are important to power system selection other than 

the few just discussed in quantitative terms. A large number of factors 

associated with development, mission and operational flexibility could be 

importantly influential in selection decisions. Specifically, prelaunch opera:. 

tions, launch interfaces, laboratory activation, on-orbit operations interfaces, 

experiments interference, emergency power, crew safety, crew demands, 

complexity, growth, and development risk are all categories that could have 

important bearing on such decisions. All of these factors are discussed 

in a qualitative manner in Appendix D as they would be affected by selection 

of a solar cell power system and by selection of a fuel cell power system and 

comparisons are made. These discussions were prepared originally for the 

• 

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1111111•1111MIL 

(Gemini)1MOL program office and are directly applicable to this study effort. 

A tabular summary of the discussion and comparison results is given in Table 5-7. 

Table 5-7 shows the over-all preponderance of practical advantages that would 

accrue to Apollo/MOL from the selection of a fuel cell power system in preference 

to a solar cell power system. A similar comparisOn between and 

solar cell systems was not prepared. Study of Appendix D will show, however, 

that all bad features of solar cells would still exist in a comparison with 

systems and many of the good features of fuel cells apply to the 

system. Also, it is clear from Appendix D that, if fuel cells were 

selected to meet the two-year availability requirement (as this study shows they 

should be), redesigning the power system for a longer mission capability would 

be immensely more complicated if solar cells were selected for the growth 

system instead of 

5.5 RECOMMENDED POWER SYSTEMS 

As a result of the power system comparisons in the previous section, the 

recommended Apollo /MOL electrical power systems are as follows: 

1. Two-year availability; select current technology fuel cells. 

2. 

These recommendations are summarized in Table 5-8 which characterizes the 

power systems selected for each of the five Apollo/MOL configurations specified 

by the program office. Also shown in Table 5-8 are the power system characteristics in Apollo and in (Gemini)MOL. 

An important over-all program recommendation rendered ffom the power system 

study is that early (two-year availability) missions should be designed around 

a 30-day capability. If a long duration (e. g., 120-day) capability had to be 

designed within two years, the power system would have to be built around solar 

cells (as indicated in Table 5-8). Since this study indicated that a solar cell 

power system would be basically undesirable for the Apollo/MOL mission, it 

follows that the deferment of a 120-day design until a system could 

be obtained would be preferable. 

this document contains Information affecting the national defense of the United States within the meaning of the Espionage Laws. Title 

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• 

5. 6 COMPARISON WITH GEMINI/MOL POWER SYSTEM 

Power system requirements with Apollo/MOL differ in no essential way 

from those in Gemini/MOL. Probably the only difference of importance is 

that "maximum use of existing hardware" in Gemini/MOL causes special 

consideration to be given to the Gemini fuel cell instead of to the Apollo 

fuel cell. 

Studies performed for Gemini/MOL to select a power system have been 

very similar in approach to that for Apollo/MOL described.in this report. 

In particular, the results have been the same as regards recommendation -

of fuel cells over solar cells as the primary power source. Much more 

detailed attention has been given to the question of which fuel cell, however. 

As yet no decision has been made as to which fuel cell concept(s) will be 

developed for (Gemini) MOL. In fact, several Air Force funded studies 

by industry are in progress with the objective, of providing a proper basis 

for making such a decision. 

Regardless of which power source might be expected for use with Gemini/ 

MOL, it is logical to suppose that the best power system for Apollo/MOL 

is also best for Gemini/MOL and vice versa. This logic stems from the 

close similarity in power system design requirements in the two cases. 

Viewed in this way, there should be no significant differences between 

an Apollo/MOL power system and a Gemini/MOL power system. 

5-33 

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5.7 CONCLUSIONS. 

1. Power source candidates for Apollo/MOL are limited to solar. cells, 

' fuel cells and Of these the fuel cell should be considered 

as a.limited duration (30 to 40 days) source, whereas the solar cell and 

systems qualify readily as 120-day power sources. 

although they appear to offer 

the best all around power system by a wide margin. Consequently, only 

solar cells and fuel cells are contenders for early missions. 

3. Current technology fuel cells are probably superior to the present Apollo 

fuel cells. Comparisons between current technology fuel cell power systems 

.and a solar cell power system indicated that many practical advantages would 

accrue to Apollo/MOL if the fuel cell system were selected. Consequehtly, 

the current technology fuel cell system is judged preferable for the initial 

Apollo/MOL missions. 

4. If a long duration (120-day) Apollo.•/MOL had to be available with a twoyear lead time, the power system would have to be built around solar cells 

since Since 

this study indicated that a solar cell power system would be basically undesirable for Apollo/MOL, it was concluded that the 120-day design should 

be deferred 

5-34 

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APPENDIX A 

APOLLO/MOL POWER SYSTEM USING CURRENT 

TECHNOLOGY FUEL CELLS 

An important reason for the selection of Pratt and Whitney's- "Bacon" 

fuel cell for the Apollo mission was its high heat rejection temperature. At 

the time of selection, it was planned that the Apollo spacecraft would land 

on the lunar surface where heat rejection would be required in spite of lunar 

surface temperatures as high as 250 F. Since that selection, which occurred 

when fuel cells were little more than laboratory curiosities, considerable 

insight has been gained into which electrochemical concepts offer the best fuel 

cell performance and life characteristics. In particular, the medium temperature (185 - 200 F) catalyzed electrode approach is currently demonstrating 

major advances over the Bacon (375 - 425 F) and ion exchange membrane 

(75 - 100 F) fuel cells being developed for the Apollo and Gemini missions, 

respectively. Better known examples of medium temperature fuel cells are 

the Allis-Chalmers fuel cell, the Union Carbide fuel cell, and the Pratt 

and Whitney '!compact" fuel cell. 

It is entirely possible that the Apollo and Gemini fuel cell concepts may 

be at higher states of development than the medium temperature approach. 

Because of this possibility, they may be preferable for use with Apollo /MOL 

despite their indicated performance inferiority. At any rate, the medium 

temperature fuel cell, hereafter referred to as the "current technology" 

fuel cell, is an important contender for application to the Apollo/MOL mission. 

Data used in the following definition of a power system, using "current 

technology" fuel cells are based on the Allis-Chalmers fuel cell for which 

the most abundant supply of information was available. 

A. Availability 

Fully qualified flight hardware could be made available in 18 to 24 months. 

This includes all power system hardware other than the fuel cells since it 

should not take any longer to develop this other hardware for current technology 

fuel cells than the 18 months maximum estimated in connection with the 

Apollo fuel cell power system. 

B. Basic Module Configuration 

A single module such as that designed for the Saturn IVB can deliver 

more than enough average power (two kilowatts). Also, this fuel cell has 

a life capability equal to or greater than the 720-hour basic mission duration 

of interest. As with the Apollo fuel cell, however, it is desirable to have 

a pair of modules always functioning so that if one fails, theLother can handle 

the entire load without any power interruption. 

UNCLASSIFIED 5-35 

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Unlike the Apollo fuel cell, a second operating fuel cell is not 

necessary to satifactory power system performance over the entire mission. 

Thus, a second operating current technology fuel cell is in active redundancy 

and contributes markedly to reliability. Assuming each module had the 

same failure rate (82 x 10-6failures/module-hour) as the Apollo fuel cell 

gives a 30-day reliability for two operating modules of 0.9967. Adding a 

third module as a non-operating spare gives a 30-day reliability of 0.99992. 

It was concluded that a total of three- modules would be required for a 

30-day mission, two operating and one held in passive standby, in order to 

meet the power system reliability objective established for this study. 

C. Weight 

A summary of the weight versus mission duration of a current 

technology fuel cell system is given in Table A-1. Note that the current 

technology fuel cell is self-starting; that is, it does not require an energy 

input to warm up and begin delivering power. Hydrogen and oxygen reactant 

consumption rates proved to be almost identical to those of the Apollo fuel 

cell. All other weight items (except, of course, the fuel cell modules 

themselves) are the same or equivalent to the Apollo fuel cell power system. 

D. Reliability 

Table A-2 shows the power system success probability determination 

fbr a 30-day mission. It is important to note that the module reliabilities 

are based on a failure rate assumed equal to that of the Apollo fuel cell. 

In view of the great amount of development money that has been spent on the 

Apollo fuel cell, this may be an optimistic expectation. 

E. Volume 

Volume values for the current technology fuel cell power system are 

the same as for the Apollo fuel cell except for the former's smaller and 

fewer modules and no start-up batteries. Table A-3 gives volume estimates 

for the current technology fuel cell system. 

F. Costs 

Power system costs are estimated in Table A-4 using current technology 

fuel cells. Fuel cell module costs are based on estimates from Allis-Chalmers. 

All other costs are the same as for the Apollo fuel cell system. 

UNCLASSIRE6 36 

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APPENDIX B 

APOLLO/MOL POWER SYSTEM 

USING SOLAR CF.LLS 

A. Array Orientation 

One of the most fundamental decisions to be made in configuring a 

solar cell power system is how the solar arrays will be oriented. A study 

ground rule was that the spacecraft would be earth oriented (roll axis 

coincident with velocity vector and yaw axis aligned with Local -veitical). 

A variety of associated array orientation control modes has been studied 

in the past (in connection with MOL) in which the arrays had zero, one, 

and two degrees of freedom relative to the spacecraft. Only two cases 

proved to be of interest from a weight standpoint; two degrees of freedom 

(full array orientation) and one degree of freedom in which the large array 

surfaces were always para.P.el to the velocity vector (roll control). 

Figure B-1 shows the altitude - mission duration regimes in which 

each of these array1.-types would be lighter weight. This weight comparison 

includes consideration of the arrays themselves, batteries, altitude 

sustenance propulsion due to aerodynamic drag on the arrays, attitude 

control propellant required because of array drag, and array stowage, 

deployment, support and orientation weights. It can be seen that roll 

control arrays would be lighter weight only with long duration, low altitude 

missions which are of doubtful over-all spacecraft capability and, therefore, 

of doubtful interest. Since the fully oriented arrays would be most competitive with other power system types, they were selected for use in defining 

the characteristic., of a solar cell system. Total area of fully oriented 

arrays was 605 ft". 

B. Weight 

A weight summary for an Apollo /MOL solar cell power system is 

presented in Table B-1. Important facts and assumptions underlying the 

weight determinations are discussed below for each item. 

Solar arrays: Solar arrays weights are based on Program 461 experience 

corrected for differing charged particle radiation environments. Program 461 

solar arrays are the largest arrays which have been flown and closest of any 

designed to meeting the Apollo/MOL requirements from the standpoint of 

stowage, deployment and orientation. Therefore, they are the best available 

indicator of what could be achieved with Apollo /MOL. Despite this, there 

is little doubt that an analysis of the dynamic response requirements of the 

arrays, as dictated by attitude control needs, would show the array weights 

in Table B-/ to be grossly underestimated. Such analyses have been conducted 

in connection with (Gemini) MOL and have shown that obtaining the required 

array rigidity would increase array weights over those shown by something in 

excess of several hundred to several thousand pounds. An alternative method 
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of obtaining the required dynamic decoupling would be to increase the flexibility of the arrays with a design akin to a window shade arrangement. Such 

a design would be highly experimental and non-conforming to the "current 

state-of-the-art" design requirement. 

Array Stowage, Deployment and Support: These weights were estimated 

on the basis of folded arrays stowed under external fairings during launch 

and subsequently deployed by lazy tong. devices. 

Array Orientation: The weight estimated includes drive members, 

gear trains and a drive motor with two spaies for reliability purposes. Two 

sets of orientation gear were assumed consistent with having two arrays, 

one on each side of the spacecraft. 

Secondary Batteryi.' Silver-cadmium batteries at 15.watt-hours per 

pound were sized on the basis of a 50 percent depth of discharge. Reliability 

calculations subsequently showed the need for a 75 percent redundancy level 

in connection with the batteries which was then added in. 

Battery Charge Regulators: Four secondary batteries each with its own 

charge regulator was assumed as the basic battery arrangement. A 

redundancy of 100 percent on the regulators gave a total of eight of these 

devices. 

Reserve and Pyrotechnic Batteries: Reserve batteries would be 

required to power Apollo /MOL during the period between launch and power 

system activation on orbit. A total of 15,000 watt-hours was allocated for 

this purpose. This same battery (or a portion thereof) could be used to 

supply power to the re-entry vehicle after separation from the laboratory. 

Consequently, no additional weight was included for this function. A special 

(high current) pyrotechnic battery at 15 pounds was included to handle squib 

firings, motor startings and the like. 

Power Conditioning, Control: and Distribution: The same standard 

"set" of this type of equipment that was used in connection with the Apollo 

fuel cell system was applied to the solar cell system. 

Associated Penalties: Aerodynamic drag on the solar arrays could 

significantly influence the functions of altitude sustenance and spacecraft 

attitude control performed by the propulsion and reaction control subsystems, 

respectively. These affects were evaluated for various circular orbit 

altitudes with varying mission durations. Propellant requirements for 

altitude sustenance (attributable solely to the solar arrays) were calculated 

from Reference i data and from drag coefficient data used for - 

Gemini /MOL. These propellant weights were then multiplied by 

. . 

UNCLASSIIIFBED 5-44 

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correction factors to account for tankage and other propulsion hardware 

associated with propellant utilization. Alternatively, it would be possible 

to launch to a sufficiently high initial altitude such that continuous decay of 

the orbit would not cause re-entry over the duration of the mission. 

However, the presence of solar arrays would cause the drag and decay 

rate to increase such that a higher initial altitude would be required with 

a solar cell system than with a fuel cell or system. This 

higher initial altitude requirement creates a payload loss which would have 

to be taken as a weight penalty for the solar cell system. Analysis showed 

that the weight penalty so obtained was equivalent to the solar array propulsion weight penalty obtained by maintaining a constant orbit altitude of 

about 170 nautical miles. To obtain attitude control penalties, maximum 

disturbing torques of solar array drag origin were assumed to be continuously 

balanced by thrusting from the reaction control subsystem. The propellant 

requirements resulting from this ultra-conservative analysis were small 

enough so that refinement of the analysis was not warranted. 

C. Reliability 

An estimation of the over-all reliability of a solar cell power system 

is shown in Table B-2. Redundancy was added in a manner that increased 

weight the least until the goal of 0.9995 for 30 days was attained. Array 

reliability stemmed from having sufficient excess operating solar cells so 

that array degradation could occur without power output "failure". Array 

deployment and orientation reliabilities were based solely on probabilities 

of drive motor failures. Power conditioning equipment reliabilities were 

taken from Westinghouse data while the other basic reliability data were 

from various industry sources. Extension of the 30-day estimate to obtain 

a 120-day estimate was done using the simple exponential model R = 

D. Volume 

Power system and associated equipment volumes are tabulated in 

Table B-3. Note that the internal volume consumed by the basic power 

system is relatively small but that the contribution to total volume from 

"power system associated" propulsion can be quite large for longer duration 

missions. 

E. Costs 

Estimated solar cell power system costs are shown in Table B-4. 

Costs in the "other" and "AGE° categories are essentially the same as those 

estimated for other power system types in these or equivalent categories. 

5-45 

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S 

APPENDIX D 

QUALITATIVE COMPARISON OF FUEL CELLS AND SOLAR CELLS 

FOR APOLLO /MOL 

Elements of the Apollo /MOL mission which could be affected by power 

system selection through factors other than weight, cost, volume, and 

reliability have been categorized as follows: 

1. Prelaunch Operations 

2. Launch Interfaces 

3. Laboratory Activation 

4. On-Orbit Operations Interfaces 

5. Experiments Interference 

6. Emergency Power 

7.. Crew Safety 

8. Crew Demands 

9. Complexity 

10. Growth 

11. Development Risk 

Each of these categories is discussed below in terms of the use of fuel cells 

or solar cells for Apollo/MOL, followed by a tabular summary of the 

qualitative discussion results (presented in the main body of the report). 

1. Prelaunch Operations 

No unusual burdens from either fuel cell or solar cell systems are 

foreseen for prelaunch (or postlaunch) ground support activities: A singularly 

important reliability advantage accrues to the fuel cell in this category, however, 

because of its amenability to thorough preflight checkout. In fact, operation 

of a fuel cell system would be started some number of hours before launch 

and would continue up to, through, and after launch. A solar cell system, on 

the other hand, cannot be checked out on the pad prior to flight. The 'necessarily 

flimsy panels cannot support their own weight under one-g)and solar simulation• and illumination of even large panel segments would require prohibitively 

elaborate equipment provisions. Because of the great risk to damage involved 

with even superficial prelaunch check-out of solar cell panels, current practice 

in one major Air Force program has resulted in no prelaunch check-out, 

except for simple electrical continuity checks, once the panels have left the 

fabricating contractor's facility. 

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2. Launch Interfaces 

Fuel cell systems would be entirely contained within Apollo/MOL 

and would have no significant launch-peculiar interfaces. The dynamic 

environment during a launch apparently poses no threat to fuel cell operation. 

Both Gemini and Apollo fuel cells have successfully undergone "double severity" 

vibrations tests 'Without a trace of trouble. Solar cell panels, it is assumed, 

would be folded and housed externally to Apollo/MOL beneath protective fairings. 

Two important interfaces are created with this approach; viz., aerodynamics 

of the launch vehicle/payload with side protuberances, and structural dynamics 

of such a configuration. Present rough estimates indicate that the problems 

created by these interfaces may not be important. However, they wourd 

definitely constitute amengineering and design burden of significant magnitude 

to ensure that they would not be a source of mission degradation or failure. 

3. Laboratory Activation 

As was noted earlier, a. fuel cell system would be operating 

throughout launch and ascent. Consequently, laboratory (and Apollo) power 

will exist continually and will not be a problem for the crew during laboratory 

activation. General Dynamics recently reported on a sequence-time study of 

solar cell power system activation for (Gemini) MOL which showed that 222 

minutes of elapsed time were required before the power system could be 

declared fully operational. A substantial portion of the 222 minutes (for at 

least one crewman) was devoted exclusively to the solar cell power system. 

4. On-Orbit Operations Interfaces 

a. Environmental Control System - A fuel cell would have an 

important interface with the ECS since about 30 percent of all heat rejected 

by the ECS would come from cooling the fuel cell. This burden is offset 

to some extent by storing excess water from the fuel cell for use in the ECS 

water boiler during periods of high heat load. Also, some alleviation is 

possible by utilizing the heat capacity of the cryogenic reactants consumed 

by the fuel cell. The net effect of using a fuel cell power system, however, 

is a substantial thermal burden on the ECS. A solar cell system would 

necessitate battery and battery charge control unit temperature control which 

would be a minor interface with ECS. 

b. Life. Support - A beneficial interface of fundamental significance between fuel cells and life support is water supply. A fuel cell 

system would provide more than enough water for all crew drinking water 

and wash water needs. The interface would simplify water management in 

Apollo/MOL by eliminating water purification provisions. With a fuel cell, 

the only provisions required would be for collecting the water as it comes 

from the fuel cell. Another interface which could be obtained if it proved..: 

desirable would be to integrate life support and fuel cell oxygen storage. A 

solar cell power system would not ititeract with the life support system. 

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• 
c. Maneuvering - Very limited, if any, spacecraft maneuvering 

would be possible with fully deployed solar panels because of their frail construction. Methods for retraction of panels to permit maneuvering are conceivable 

but would be complex and time-consuming during operation. A fuel cell 

power system would not affect maneuvering. 

d. Attitude Control - A solar cell system would have a large 

impact on attitude control and vice versa. The Apollo /MOL spacecraft roll 

moment of inertia would be increased markedly by solar cell panels, thereby 

increasing propellant consumption for stabilization about the roll axis. Exhaust 

impingement on the solar panels from the reaction control thrustors would 

constrain the design and possibly the relative positioning of these equipments. 

Aerodynamic drag on the paddles would disturb attitude control about the pitch 

and yaw axis. Whether this disturbance would be beneficial or detrimental as 

regards spacecraft stabilization is not known at this time. In the other 

direction, precision attitude control requirements would dictate sufficient 

solar panel stiffness (substantially more than is otherwise required) so that 

panel flutter induced by reaction control pulses would damp out rapidly. A 

fuel cell system would not interact significantly with the attitude control system. 

e. Orbit Keeping Propulsion - Aerodynamic drag on solar cell 

panels would create a need for larger amounts of propellant and more frequent 

thrusting if a nominal orbit altitude were to be maintained. Beyond this, 

acceleration forces from thrusting to regain lost altitude would be a constraint 

on the design of solar cell panel structure and possibly panel orientation control 

mechanisms. No propulsion interfaces would be involved with the use of a fuel 

cell power system. 

Communication - The requirement for broad beam antenna 

patterns from Apollo/MOL to ensure adequate ground coverage guarantees 

communications interference from solar cell panels. The degree of interference would vary with panel orientation and could result in periodic waxing 

and waning of signal strength. How severe this interface problem would 

actually be cannot be assessed at this time. A fuel cell system would not 

interface with communications. 

g. Rendezvous and Docking -Close-in maneuvers by a ferry 

vehicle to achieve rendezvous and docking with an Apollo/MOL spacecraft 

would almost certainly be influenced by the presence of large fragile solar 

panels extending 15 to 20 feet from either side of Apollo/MOL. It is not 

possible at this time to estimate reliably how rendezvous and docking might 

be constrained .by such influence or the importance of this interface. However, 

no such interface would exist in this area for fuel cells. 

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taremmala 

5 Experiments.Interference 

Probably the most conclusive factors against the use of solar cell 

panels for Apollo /MOL arise from considerations of experiment interference. 

A sketch is shown below of the Apollo/MOL spacecraft with solar cell panels 

and the range of panel positions which would occur as the panels are oriented 

for maximum insolation. Panel size and spacecraft size are in proper 

proportion in the sketch for about 1.5 kilowatts average power and fully 

oriented panels (roll-only orientation of panels would necessitate panels 

almost twice as large as shown). Note that the theoretical angle range (2 8) 

swept out by each solar cell panel is symmetric about local vertical (neglecting 

solar eclipse). For purposes of illustration in the sketch, it is convenient 

to assume that the panels are parallel to the flight direction (roll axis). 

However, the same basic panel motion will result from full panel orientation 

in which such parallelism is not maintained. Neglecting eclipse affects would 

give a maximum value for 8 of 90 degrees. Considering that panel orientation 

is not required during solar eclipse periods limits the maximum value of 8 

in the non-cross-hatched zones to 50 to 60 degrees while leaving e in the crosshatched zones unchanged. 

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\ • \\\\ 

\ \ \ 

UNCLASSIFIED 5-62 

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Experiments P1, P2, and P3 

Experiment P7 - 

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1441111111111.110, ,.„.„ 1111

• 

Experiment interference for each of the experiments currently 

in planning for Apollo/MOL is considered below in context with the above 

skeWi. 

Experiment P4 - This experiment may involve a five- to eightfoot diameter parobolic receiving antenna deployed immediately beneath the 

Apollo /MOL spacecraft. Both physical interference and signal interference 

from a solar panel would be a problem with this antenna. 

Experiment P5 - Deleted 

Experiment P6 - Extravehicular activity is the subject of this 

experiment and would involve an astronaut leaving via an escape hatch in the 

top of the spaceciaft. The presence of an overhead solar panel could interfere with such an operation. 

Experiment P8 - Autonomous navigation by the Apollo/MOL crew 

is attempted in this experiment. At least one facet of this experiment which 

would be confronted with interference by solar cell panels is manual star 

"shooting" through a view port in the top of the spacecraft. 

Experiment P9 - Maintenance - Undefined. 

Experiment P10 - A manually directed radiometer would be 

pointed towards earth in this experiment both at preplanned targets and at 

targets of opportunity. Obscuration of the field of view of the radiometer 

would be a problem with solar cell panels for this experiment. 

Experiment P11 - General (astronaut) performance in a military 

space system - no interference. 

Experiment P12 - Biomedical - no interference 

A fuel cell power system poses no experiment interference problems 

for Apollo /MOL. 

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6. Emergency Power 

Power system sizing was accomplished on the basis of essentially 

equivalent emergency power provisions for both the fuel cell and solar cell 

system. No advantage is believed to accrue to either power system in this 

area. 

7. Crew Safety 

Neither power system type possesses features that would significantly affect cretv 'safety. 

8. Crew Demands 

It was pointed out in Section 4 that a substantial amount of crew 

time and effort would be involved in activating a solar cell power system. 

This is the only unique demand on crew time that can be foreseen. All other 

demands such as periodic power system check-out should be about the same 

for both systems. ThiS assumes that solar paddle orientation is fully automated throughout the mission and that the fuel cells purge themselves 

• automatically. 

9. Complexity 

Types of failures and failure modes are difficult to predict even 

in the late stages of mission development. Failures which can result from 

inadequate quality control, improper handling, misuse or reasons unknown 

are not accounted for in most system failure models. Consequently, preliminary reliability analyses, which are necessarily based on failure models, 

are generally insufficient to establish relative probabilities of failure between 

two systems. An additional measure of failure probability is the quantity of 

possible failure sources, i. e., complexity. A solar cell system would 

contain, for example, many gear trains and mechanical linkages whose 

predictable probabilities of failure could be negligible. Nonetheless, they 

are potential failure sources whose satisfactory performances are reasonably 

sensitive to quality control, handling, etc. Similar comments are applicable 

to the valving and complex structuring in fuel cells. It is estimated that fuel 

cells and solar cells are about equivalent over-all relative to such complexity 

considerations and that arguments favoring either power system in this regard 

would be uncertain at best. 

10. Growth 

a. Different Altitudes - Fuel cell system characteristics are 

independent of altitude. A solar cell system penalizes Apollo /MOL from an 

altitude sustenance propulsion standpoint as explained earlier. Consequently, 

lower altitude Apollo/MOL missions would be substantially compromised by 

a solar cell power system in terms of smaller payload, shorter duration or 

in some equivalent manner due to increased propulsion requirements. On the 

other hand, of course, a solar cell powered Apollo /MOL could decrease its 

UNCLASSME 5-64 

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tNtetIIIIIIIIIIIIIIIIIINPIIIIL 

e 

altitude sustenance propulsion for higher altitude orbits. Higher altitudes are 

not so. attractive , however, because of the personnel shielding requirements 

that arise, and because of reduced resolution for observations of earth. 

b. Longer Mission Duration - A solar cell system has a life 

capability well beyond 30 days and could perform for many months without 

significant design or electrical performance changes. A fuel cell power system 

is much more limited in this regard. Since a fuel cell system can, within 

bounds, trade power level for mission duration, modest (e. g., 10 to 15 days) 

extensions of mission duration capability could be obtained by reducing 

average Apollo /MOL power consumption to some- appropriate level.. Beyond 

this, reactant resupply from a ferry vehicle would be required to render a 

basic 30- to 40-day fuel cell design adequate for long mission extensions. 

c. Different Power Levels - A fuel cell consumes reactants 

only as needed and at a rate roughly linear with power output. Also, they 

can operate over a wide range of power densities. Fuel cells readily adapt, 

therefore, to either lower or higher average power levels and to either lower 

or higher peak power levels. In every case, the volume and weight cost per 

unit of energy delivered is (within bounds) roughly the same. A solar cell 

system must be redesigned with larger solar panels to produce a higher 

average power output. Higher sustained peak power levels also require a 

redeSign involving larger batteries and possibly larger solar panels. Lower 

power levels can be accommodated without solar cell system redesign with, 

of course, no reduction in power system- weight. 

d. Polar (Operational) Missions - Fuel cell performance is 

independent of orbit parameters. Similarly, a solar cell power system would 

produce its design power at any orbit• inclination. Therefore, either type 

of power system is amenable to Apollo/MOL adaption to operational missions. 

e. Vulnerability (Operational Missions) - An important advantage of fuel cells in an operational mission environment is their much superior 

vulnerability characteristics. Solar cell panels would increase the radar 

cross-section of Apollo /MOL by a factor of at least two or more, thereby 

facilitating missile tracking. Also, the power generating properties of solar 

cells can be readily destroyed by distant nuclear detonations. Fuel cells would 

have no influence on the basic Apollo/MOL spacecraft vulnerability. 

11. Development Risk 

a. Power System Equipment - Both fuel cell and solar cell 

technologies are sufficiently advanced to warrant their selection for an 

accelerated Apollo/MOL development program. It can be seen from the previous 

discussions, however, that a solar cell power system involves much greater 

complexity in operating procedures and interfaces with other subsystems. 

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Consequently, the number of opportunities Cor development troubles (i. e. , 

development pitfalls) is estimated to be much greater for a solar cell 

system than for a fuel cell system. For example, large appendages housing 

stowed panels on the sides of Apollo/MOL may prove dynamically unsound. 

Also, current experience in the deployment and orientation of solar panels 

has proven so troublesome that a major Air Force program has abandoned 

oriented panels in favor of fixed panels, and is starting, for the second time, 

with development of conceptual deployment schemes. On the other hand, 

there is limited zero-g experience with both fuel cells and their associated 

cryogenic tankage. While there is not any serious question of the ability of 

either of these equipments to function in a zero-g environment, the degree 

to which their in-orbit performance (e. g., specific reactant consumption, 

reactant boiloff rate) will meet expectations will not be known for some time. 

Thus, in comparing solar cell risks to the fuel cell risks, the possibility 

of obtaining, at the worst, a few days less mission duration capability from 

fuel cells is preferable to the large number of uncertain design problems and 

interfaces, any one or all of which could be much more harmful to the 

Apollo/MOL mission. 

b. Apollo/MOL Power Requirements - Power requiremerits 

for an electrically complex system such as Apollo/MOL become accurately 

defined late (if ever) in the design of the spacecraft. At this conceptualplanning stage for Apollo /MOL, it is a certainty that the eventual power 

requirements for Apollo /MOL will be importantly different from current 

power profile projections, with a probable trend to higher power levels. 

Solar cell power systems have a fixed power generating capacity once their 

design is frozen. Higher (average) power from them requireS redeSign with 

resultant reverberations through all of its interfaces. To avoid this, either 

a highly conservative solar cell system design would have to be employed 

initially, the design would somehow have to be kept flexible until late in 

development, or power requirements would have to be frozen early in 

development. None of these undesirable design constraints is necessary with 

a fuel cell system which simply consumes reactants faster if more power is 

drawn from it. A conservative cooling design in the fuel cell units plus a 

trade-off of mission duration for higher average power requirements on a 

one-for-one basis causes power profile variations to be of only modest design 

significance for .fuel cell powered spacecraft. 

UtIcLASSIFIEb" 

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TABULAR SUMMARY 

Table 5-7 in the main body of the report summarizes the results of the 

foregoing comparisons. No attempt was made ttt assigning relative leirels 

of importance to the various comparison factors; some should undoubtedly 

be considered, more important than others in selecting the Apollo /MOL 

power system. 

CONCLUSIONS 

It is concluded from the foregoing discussions and the revealing summary 

of Table 5-7 that a fuel cell power system for Apollo/MOL has an overwhelming 

number of advantages in comparison to a solar cell system. Taken one at a 

time, each solar cell system disadvantage probably could be conceptually 

circumvented or judged tolerable. Viewed en masse, as in this appendix, such 

arguments become rather academic to the over-all preponderance of practical 

advantages that would accrue to Apollo /MOL from the selection of a fuel cell 

power system. 

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• 

REFERENCES 

5-1. Buglia, J. J. and J. F. Newcomb, "A Metliod for Determining the 

Propellant Required to Maintain an Earth Orbital Vehicle in a 

Useful Orbit Between Two Specified Orbits," NASA TN D 1817, 

August 1963. 

5-69 

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SECTION 6 • LIFE SUPPORT 

C: C. Wright 

SUMMARY 

A study was conducted of the life support subsystems for various 

Apollo/MOL vehicle configurations. The weights, volumes and power 

requirements of two candidate systems were estimated and compared 

with those of the Gemini B /MOL. In addition, the life support system 

characteristics, location of equipment in the vehicle, compartment 

volumes and major features for each vehicle configuration were 

identified and compared. 

The two candidate systems selected for study were (1) lithium 

hydroxide method of CO2 removal with water supply from the fuel cells 

and (2) molecular sieve method of CO removal system with water re- 2 

clamation and a solar cell power supply system. The cross-over point 

between these two systems for two men at an orbit of 160 nautical miles 

(with altitude sustenance and one repressurization) is about 35 days. 

The lithium hydroxide CO2 removal system has the lowest weight for 

mission durations less than 35 days whereas the molecular sieve CO2 

removal system is lighter for mission durations greater than 35 days. 

There are no significant differences for the life support and 

environmental control systems between the Gemini/MOL and Apollo/MOL. 

• 

6-1 
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• 

• 

CONTENTS 

Page 

6.1 INTRODUCTION 6-7 

6. 2 BASIC WEIGHT AND BULK DENSITY DATA 6-8 

6. 3 POWER REQUIREMENTS AND WEIGHT PENALTY 6-8 

6. 4 WEIGHT VERSUS MISSION DURATION 6-11 

6. 4. 1 LiOH CO2 Removal System, 5. 0 Psia Oxygen 

Atmosphere, Fuel Cell Power and Water Supply 6-12 

6. 4. 2 Molecular Sieve CO2 Removal System, 7 Psia 

Oxygen/Nitrogen Atmosphere, Solar Cell Power 

With Water Reclamation 6-13 

6.5 COMPARISON OF LIFE SUPPORT SUBSYSTEM 

DATA 6-15 

6. 5. 1 Vehicle Configuration, Mission Duration, and 

Number of Men 6-16 

6. 5. 2 Power Supply 6-16 

6.5.3 Atmosphere 6-16 

6. 5. 4 Compartment Volumes 6-20 

6. 5. 5 Location of Equipment 6-21 

6. 5. 6 Method of CO2 Removal and Water Supply 6-21 

6. 5. 7 Contaminant Control 6-21 

6. 5. 8 Leakage Rate 6-21 

6. 5. 9 System Weights, Volumes and Power Requirements 6-21 

6. 6 CONCLUSIONS 6-22 

REFERENCES 6-23 

• 

6-3 
TABLES 

6-1 

6-2 

6-3 

Page 

6-9 

6-10 

6-17 

Basic Weight and Bulk Density Data, Apollo/MOL 

Life Support Subsystems 

Estimated EC/LS System Power Requirements, 

Apollo/MOL Study 

Comparison of Life Support Subsystem Data 

Apollo/MOL Study 

ILLUSTRATIONS 

Page 

6-1 Life Support Subsystem Weight Versus Mission 

' Duration 6-14 

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• 
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6.1 INTRODUCTION 

A study was conducted of the life support subsystem characteristics of two 

candidate systems for various possible Apollo/MOL configurations. Weights, 

volumes, and power requirements were estimated and compared with those of 

the Gemini B/MOL. Life support subsystem characteristics, location of 

equipment, compartment volumes and major features for each configuration 

were also identified and compared. 

In accordance with the Apollo/MOL study ground rules, .the various vehicle 

configurations and crew requirements that were studied are as follows: 

1. Basic Apollo, three men for 14 days 

2. NAA Concept II, two men. for 30 days 

3. Semi-optimum NAA Concept II, two men. for 30 days 

4. Semi-optimum NAA Concept II, two men for 120 days 

5. Semi-optimum NAA Concept II, three men for 120 days 

6. Aerospace Recommended Configuration, two men for 30 days 

7. Gemini B/MOL, two men for 30 days 

8. Gemini B/MOL, two men for 90 days 

• The vehicle configuration for the North American Aviation Concept II consists 

of the Apollo command module, the Apollo service module and a large pressurized laboratory module in the LEM adapter. The atmospheric control 

system is located on the command module. This configuration is based on the 

NAA extended mission Apollo Study (Reference 6-1) and is described in 

Section 1. The semi-optimum configurations (Section 1) have shorter service 

modules- and smaller laboratory modules. In the Aerospace recommended 

configuration (Section 1) the service and laboratory modules are combined, 

the cryogenic storage tanks and propulsion system are relocated in the aft 

section, and the laboratory contains its own environmental control and life 

support subsystem similar to the Gemini B/MOL. All Apollo/MOL configurations use the Saturn IB launch vehicle, whereas the Gemini B/MOL uses the 

Titan-IIIC launch vehicle. 

6-7 
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The two candidate EC/LS subsystems selected for analysis are (1) lithium • 

hydroxide type of CO2 removal system with water supply from the fuel cells 

and (2) molecular sieve type of CO2 removal system with water reclammation 

and a solar cell power supply system. 

In the first system, the CO2 expired into the air by the astronauts is removed 

by a chemical reaction in LiOH beds. No regeneration is possible, and the 

beds must be replaced with a fresh charge at regular intervals. In the second 

system, the CO2 is removed from the air, which is first dried with silica gel, 

by absorption on the surface of zeolite granules called molecular sieves. Here, 

regeneration (desorption) of the beds is possible by periodic exposure to the 

vacuum of space. Two silica gel-molecular sieve units operating cyclically in 

parallel are required so that while one unit is drying the air and absorbing CO2, 

the other unit is being desorbed by CO2 and water. Although the CO2is lost to 

space, the water, which was absorbed by the silica gel bed ahead of the 

molecular sieve, is returned to the purified air stream. 

6. 2 BASIC WEIGHT AND BULK DENSITY DATA 

The weights and volumes of the various fixed and variable items of the Apollo/ 

MOL life support subsystems may be estimated from the basic weight and bulk 

density data shown in Table 6-1, which was obtained or adapted from various 

sources, principally References 6-1 and 6-2 and the Gemini/MOL program. 

6. 3 POWER REQUIREMENTS AND WEIGHT PENALTY 

The estimated power requirements for the various Apollo/MOL study configurations are summarized in Table 6-2. These numbers are crude estimates, 

but they should be adequate for preliminary comparison purposes. 

According to Section 5, the total weight power penalty for present improved 

version fuel cells in the 1. 5- to 2. 0-kilowatt range is 

Wfc 

= sum of the fixed power system weight, replacement 

weight, and variable weight, lb 

Wfc 

= 500 + 200 PN + 34. 8 PD for 1. 5<2. 0 kilowatts (1) 

6-8 
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Table '6- Basic Weight and Bulk Density Data 

Apollo/MOL Life Support Subsystems. 

Item Weight Bulk Density 

LiOH CO2 removal granules, , 

charcoal filters and canisters 

3.0 lb/man-day 25 lb 

Metabolic oxygen 2. 0 lb /man-day 

Cabin leakage (assumed) 2.0 lb/day at 5 psia 

2.8 lb /day at 7 psia 

Repres surization(1 ) 144 PV/RT lb 

Cryogenic oxygen tankage 

(supercritical at 1500 psia) 

0.48 lb/lb fluid 0. 0/75 ft3 /lb 07 

(inc. 5% ullager 

Food (freeze dried, including 

storage containers) 

1.5 lb/man-day 45 lb/ft3 

Sanitation and clothing 0.58 lb/man-day 40 lb/ft3 

Fixed Hardware 

(2 - 3 men) 

LiOH CO2 

Removal System 390 lb (2) • 

35 lb 

Molecular Sieve 

CO2 Removal System 500 lb(2) . 

Water Reclamation 

System . 50 lb 

V = total volume to be pressurized, ft

3 

R = gas constant (for 02= 48.3 ft#/lb-R ) 

T = cabin temperature (nominal value = 530 R) 

2. In all Concept II versions of Apollo/MOL an additional 225 lb of thermal 

control equipment should be added to the laboratory module. Atmospheric control is accomplished with the CO2 removal equipment in the 

command module plus booster fans and ducts. 

Notes: 

1. P = cabin pressure, psia 

6-9 
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Apollo/MOL Study. 

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•el • 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

• where P = average power level, kw 

N = number of 30-day periods 

D = mission duration, days 

Assuming a total average power level of 1.77 kilowatts (which is typical for 

the MOL mission), the weight penalty chargeable to the EC/LS system for a 

fuel cell power supply is 

Wfc = (Pec/Is/P1.77) (500 + 200 PN + 34.8 PD) 

= Pec/ls (283 + 200 N + 34.8 D) 

where 

Pec/ls = the power required for the EC/LS system 

D = mission duration, days 

In the case of solar cell power, the weight power penalties for an average 

power level of 1.77 kilowatts. (Section 5) are: 

1. 125 nautical miles with altitude sustenance: 

WSC = sun fixed solar cell system weight plus variable weight 

WSC = 200 +. 18.2 D for P = 1.77 kw 

2. 200 nautical miles with altitude sustenance: 

WSC = 1968 + 1.8 D for P = 1.77 kw 

The weight penalty chargeable to the life support system is Pecn1.77 

times the above values. 

6. 4 WEIGHT VERSUS MISSION DURATION 

The total weight of the life support subsystem may be computed from the data 

in Tables 6-1 and 6-2 together with the appropriate power penalty equation. 

Assuming a 10 per cent allowance for expendable reserves and an additional 

• 

6-11 
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RELEASE 1 JULY 2015 

10 per cent allowance for cryogenic residuals, the weight versus mission 

duration of the two candidate life support subsystems may be computed 

as follows: 

6. 4. 1 LiOH CO2 Removal System, 5. 0 Psia Oxygen Atmosphere, 

Fuel Cell Power and Water Supply 

W = Wfixed + 1.10 (W + W LIOH food + Wsanitation and clothing) 

+ 1.20 (W02 + Wtankage) + Wpower penalty 

Assuming one complete repressurization (2800 ft3), the EC/LS weights for the 

NAA Concept II version of the Apollo/MOL are: 

Wfixed = referring to Table 6-1, the fixed hardware weight in 

command and laboratory modules is = 390 + 225 = 615 lb 

wLiOH = 3. 0 MD, lb 

= 1. 5 MD, lb Wfood 

W sanitation = O. 58 MD, lb 

and clothing 

W0 = combined weight of metabolic oxygen, leakage and 

2 repressurization gas for 2800 ft at 5. 0 psia 

2.0 MD + 2. 0 D+ 144 x 5 x 2800/48. 3 x 530 

W = 2.0 MD + 2. 0 D + 78.7 lb 

02 

w tankage chargeable to the EC/LS system tankage 

0. 48 x W

o 

= 0. 96 MD + 0. 96 D + 37. 8 lb 

2 

w O. 510 (283 + 200 N + 34. 8 D) = 144 + 102 N + 17.75 D 

power 

penalty 

Substituting these weights into Equation (5), the total weight is: 

W = 615 + 1.10 x 08 x MD + 1.20 (116. 5 + 2.96 MD +2.96 D) 

+ 144 + 102 N + 17.75 D 

• 

(5) 

• 

6-12 
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Combining similar terms, the total weight of the LiOH CO2removal EC/LS 

subsystem, including power penalty, is 

W = 899 + 102 N + 9. 14 MD + 21. 3 D lb (6) 

where 

D = mission duration, days 

M = number of men 

N = number of 30-day periods (maximum useful life of fuel cells because 

of reliability consideration is 30 days) 

For 0 to 30 days, N = 1; for 30 to 60. days, N = 2, etc. 

A plot of Equation (6) for a two-man system is shown in Figure 6-1. 

6. 4. 2 Molecular Sieve CO2 Removal System, 7 Psia Oxygen./Nitrogen 

Atmosphere, Solar Cell Power With Water Reclamation 

In this case, the various itemized weights are: 

Wfixed = fixed hardware weight in command and laboratory modules 

500 +. 225 = 775 lb 

WLiOH = emergency weight for two days = 3 x 2 x M lb 

Wfood = 

1 5 MD lb 

= O. 58 MD lb sanitation 

and clothing 

0 MD + 2. 8 D + (144 x 7 x 2800)/(51. 5 x 530) 

WO2 

W = 2. 03VID + 2. 8 D + 103. 5 lb 

°2 

= 0. 48 W = 0. 96 MD + 0. 96 D + 49. 7 lb tankage 0

2 

Referring to Equation (3) and Table 6-2, the power penalty for a 125-nautical 

mile orbit with altitude sustenance is: 

w = (0.73/1.77) (2004 + 18.2 D) = 828 + 7.51 D 

power penalty 

125 nautical miles 

• 

6-13 
6000 

5000 

3000 

2000 

1000 

0 

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8000 

7000 

1 

1, WEIGHT INCLUDES POWER 

PENALTY CHARGEABLE TO 

THE EC/ LS SUBSYSTEM 

2. NUMBER OF MEN : 2 

FUEL CELL 

LI OH CO2 

SYSTEM 

POWERED, 

REMOVAL 

../ 

N tt. 

20° 

,.. 

CROSS—OVER 

POINT AT 

160 N MI 

CELL POWERED, 

SIEVE 

SYSTEM 

MOLECULAR 

REMOVAL 

— 

,---- 

SOLAR 

CO2 

20 40 60 80 100 120 140 

MISSION DURATION, DAYS 

Figure 6-1. Life Support Subsystem Weight 

Versus Mission Duration. 

6-14 
= 1787 + 6 M + 5. 84 MD + 12. 48 D lb (7) 

= 1771 + 6 M + 5. 84 MD + 5.7.1 D, lb (8) 

W naut mi 

W200 naut mi 

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Referring to Equation (4), the power penalty for a 200-nautical mile orbit with • altitude sustenance is 

power penalty = .(0.7311. 7.7) (1968 + 1..8.0) = 81.2 + 0,74 D 

ZOO nautical miles 

Combining similar terms in accordance with Equation (5), the total weight of 

the molecular sieve EC/LS systems, including power penalty, is 

where 

D = mission duration, days 

M = number of men 

A plot of Equations (.7) and (8) for a two-man system is shown on Figure 6-1. 

Inspection of this figure reveals that the cross-over point between a fuel cell 

powered, lithium hydroxide CO2removal system with water supply from the 

fuel cells and a solar cell powered, molecular sieve CO2removal system with 

water reclamation is between 32 and 46 days for 200-nautical mile and 125-

nautical mile orbits, respectively. The cross-over point at an orbit of 160 

nautical miles with solar cell altitude sustenance is about 35 days. 

6. 5 COMPARISON OF LIFE SUPPORT SUBSYSTEM DATA 

A comparison of the weights, volumes and power requirements of the life 

support subsystem for each study configuration is shown in Table 6-3. In 

addition, the life support system characteristics,' location of equipment in 

the vehicle, compartment volumes, and major features are identified and 

compared. In accordance with Figure 6-1, which shows the cross-over point 

between the two candidate life support systenis at about 32 to 46 days, all 

configurations having a 30-day mission were assumed•to be supplied with a 

fuel cell powered, LiOH CO2 removal system whereas all configurations 

having a 90- or 120-day mission were assumed to be supplied with a solar. 

cell powered,, molecular sieve CO2removal life support system. 

6-15 

• 
Inspection of Table 6-3 reveals the over-all characteristics, similarities and 

dissimilarities of the life support systems of each configuration. Note that 

• 

the Aerospace recommended version of the 30-day, two-man Apollo/MOL is 

very similar to the 30-day, two-man Gemini B/MOL, except for cabin volume. 

The differences in the life support subsystem data between the various Apollo/ 

MOL configurations in Table 6-3 may be explained as follows: 

6. 5. 1 Vehicle Configuration, Mission Duration, and Number of Men 

Specified as a study ground rule. 

6. 5. 2 Power Supply 

See Section 5. A fuel cell power supply was selected for the 30-day missions, 

whereas a solar cell or power supply was selected for the 

120-day missions. 

6. 5. 3 Atmosphere 

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Selecting an atmosphere involves trade-off considerations between the man subsystem, the environmental subsystem, and the power system. The constraints 

are (1) physiological considerations, (2) mission safety, (3) reliability, (4) 

hardware availability, (5) thermal management consideration, and (6) vehicle 

penalty comparisons. 

In order to satisfy man's physiological requirements, the alveolar oxygen 

partial pressure must be greater than about 100 mm Hg in order to prevent 

hypoxia. The corresponding inspired oxygen partial pressure should be about 

185 mm Hg. This number represents present space suit conditions and is a 

good basis for the transition from a one- to two-gas atmosphere. Also required is a water partial pressure of about 3 to 28 mm Hg and a carbon dioxide 

partial pressure of about 1 to 8 mm Hg. A 260 mm Hg (5 psia) oxygen atmosphere is probably satisfactory for the 30-day MOL mission, but longer missions 

may be detrimental because of oxygen toxicity. 

An inert gas, such as nitrogen, helium or neon, may also be beneficial. It is 

also thought that man may, in some unknown way, require nitrogen for body 

chemical reactions. However, there may be difficulties with the use of 

6-16 
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• 

• 

nitrogen as a diluent because of the possibility of the "bends" occurring when 

changing from a higher pressure two-gas system to a. 185 mm Hg "pure oxygen" 

suit system. Helium may be physiologically more attractive than nitrogen. 

In regard to mission safety, both explosion and fire hazards must be considered. 

The possible hazard caused by explosions is reduced (the minimum spark ignition 

energy for explosive combustion of gas mixtures is increased) and the rate of 

consumption of materials in the case of fire is decreased by (1) lowering the 

oxygen partial pressure while maintaining the total pressure constant, or (2) 

by raising the total pressure while maintaining the oxygen partial pressure 

constant. Combustible mixtures can also be reduced or eliminated by using a 

catalytic burner contaminant control. 

Low fire hazard materials have a high thermal conductivity, a high ignition 

temperature,. and a low heat of reaction with oxygen. Experiments conducted 

by F. A. Parker of General Electric (Reference 6-3) showed that stainless steel 

can burn in an atmosphere of 85 per cent 02and 13 per cent N2at a total 

pressure of 750 mm Hg. Aluminum is a better material because of its higher 

thermal conductivity (heat is conducted away from reaction zone, thus reducing 

the heat required to sustain the reaction) and its tough oxide coating. It should 

be noted, however, that materials that burn well in a free convection environment 

(under the influence of gravity) may not be capable of a self-sustaining reaction in 

a zero gravity environment because the products of combustion tend to dilute the 

reaction constituents. The result is the material temperature at the reacting surface may drop.below the self-sustaining level. In general, a mixed gas atmosphere (with inert mixing gases) is safer than a pure oxygen atmosphere, a lower 

oxygen partial pressure atmosphere is safer than a higher partial pressure 

atmosphere, and the possibility of a. fire, especially an uncontrollable fire, can 

be minimized by utilizing low fire hazard materials. 

Reliability considerations indicate that, on the basis of hardware requirements, 

a. mixed gas atmosphere tends to have a slightly lower reliability than a pure 

oxygen atmosphere. Nevertheless, a two-gas atmosphere can be made just as 

reliable if sufficient effort is expended. In fact, the over-all reliability of a 

combined man-equipment system may be higher with a properly selected mixing 

gas atmosphere than with a pure oxygen atmosphere. Also, during an emergency 

mode of operation both mixed gas and pure oxygen atmospheres have the 

same reliability. 

6-19 
From the point of view of available hardware, there appears to be no compelling 

reason for selecting one type of atmosphere over another. Even helium-oxygen 

atmospheres-are within present state-of-the-art capability. Thermal management considerations indicate that the order of preference is 02

/He, 02 /N2 

'and pure 02atmospheres. Finally, over-all vehicle considerations indicate 

that the desired pressure level range for minimum vehicle weight is about 5 to 

9 psia with the minimum occurring at about 7 psia. Below'5 psia the pumping 

power penalty becomes excessive and above 9 psia the structural weight is 

too large. 

Although the type of atmosphere for the Gemini/MOL vehicle has not been 

established, the present industry EC/LS study contracts are based on a 5 psia 

one-gas atmosphere (pure oxygen) for 30-day missions and a 7 psia two-gas 

atmosphere (46 per cent 02and 54 per cent N2) for 90-day missions. Other 

atmospheres are also possible. For example, a recent preliminary qualitative 

study indicates that the optimum atmosphere for the MOL mission.is a 5 psia, 

O2He mixture having a composition of 165 mm Hg of 02, 75 mm Hg of He, 

15 mm Hg of H2O and 5 mm Hg of CO2 . 

For purposes of the Apollo/MOL study the atmosphere for 30 days or less was 

assumed to be 5.0 psia "pure oxygen", whereas for longer durations the atmosphere was assumed to be a 7 psia, two-gas atmosphere consisting of 50 per cent 

02and 50 per cent N2. Except for leakage and repressurization requirements, 

variations from the above assumptions will not significantly effect the weight 

and volume estimations presented-in-Table 6-3. 

6. 5. 4 Compartment Volumes 

Compartment volumes were estimated from the sketches of Reference 6-1 and 

Section 1. • The difference between the free volumes of the command module in 

the•basic Apollo and Apollo/MOL configurations is due to the electronic equip-

, 

ment and electronic cold plates associated with the Apollo mission. Much of 

this hardware is not necessary for the Apollo/MOL mission. 

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6-20 
• 6. 5. 5 Location of Equipment 

The equipment location becomes fairly obvious by referring to Reference 6-1 

and Section 1. It should be noted, however, that the CO2 removal equipment 

in the semi-optimum versions of Apollo/MOL is located in the command module. 

In the Aerospace recommended version, CO2removal equipment is also located 

in the laboratory module. This eliminates the need for ducts and blowers (plus 

interface connecting fittings at the junction between the command and laboratory 

modules) for circulating the atmosphere between the two compartments. 

6. 5. 6 Method of CO2-Removal and Water Supply 

Inspection of Figure 6-1 reveals that the weight cross-over point between the 

two candidate combined life support-power systems is about 35 days at an 

altitude of about 160 nautical miles. The use of power would give 

about the same cross-over point as fuel cells. Using this as a criteria for 

selection, all 30-day or less missions would use lithium hydroxide CO2removal equipment with a fuel cell water supply. On the other hand, all missions 

greater than 30 days would use a molecular sieve CO2removal system with • 

water reclamation. Water reclamation is required in the latter case because 

water is not produced as a by-product from the power supply as in the case 

with fuel cells. 

6. 5. 7 Contaminant Control 

For mission durations greater than 30 days, it appears as though catalytic 

burners will be required to prevent excessive accumulation of contaminants. 

6. 5. 8 Leakage Rate 

The figure of two pounds per day for a 5 psia atmosphere was assumed to be 

the same as for the Gemini B/MOL EC/LS study. At higher pressure levels 

the leakage rate is directly proportional to the pressure. 

6. 5. 9 System Weights, Volumes and Power Requirements 

These items are adequately explained in the sections of this memorandum on 

weight, bulk density and power requirements. All of the data was either 

computed or estimated from the reference reports. Note that all variable 

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weights (food, oxygen and tankage, sanitation and clothing) are directly • 

proportional to the number of men and mission durations. Also, the power 

requirements include lighting requirements. 

6.6 CONCLUSIONS 

1. The cross-over point between a fuel cell powered, LiOH type of carbon 

dioxide removal system with fuel cell supplied water and a solar cell powered, 

molecular sieve type of carbon dioxide removal system with water, reclamation 

is about 35 days at a_160-nautical_mile circular orbit. Missions less. than 35 

days in length should use the LiOH type of CO2removal system whereas 

missions of longer length should use the molecular sieve type of CO2removal 

system. 

2. The life support subsystem in the Aerospace Corporation recommended 

version of the 30-day, two-man Apollo /MOL is very similar to the 30-day, 

two-man. Gemini BJMOL. 

3. There are no significant differences for the life support and environmental 

control systems between the' Gemini /MOL and the Apollo /MOL. 

6-22 

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• 

REFERENCES 

6-1. "Extended Mission Apollo Study," Volume V, Environmental Control 

and Life Support, North American Aviation, Inc., Space and 

Information Systems Division, Report No. 51D 63-1370-5, 

November 1963. 

6-2. Coe, C. S., J. Rousseau, and A. Shaffer,-"Analytical Methods for 

Space Vehicle Atmospheric Control Progresses, Part II," ASD 

Technical Report 61-162 (AiResearch Manufacturing Company 

Report SS-695-R, July 1962). 

6-3. Parker, F. A. and D. R. Ekberg, "Selecting the Space Station 

Atmosphere," Astronautics and Aerospace Engineering, 

August 1963. 

• 

6-23 
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• 

SECTION 7. 

PERFORMANCE 

K. N. Easley 

J. M. Dreyfus 

SUMMARY 

Abort boundaries for the Apollo re-entry vehicle have been computed 

for use in ascent flight performance calculations. The boundaries were 

computed for re-entry vehicle weights of 9000 and 12,000 pounds with liftto-drag ratios of 0. 4 and 0. 5 and a maximum total load factor of 13 g's. 

The four boundaries which result are presented in graphical form. 

A parametric study of re-entry trajectories for the Apollo re-entry 

vehicle is presented. Deboost velocity increments of 500 and 1500 feet 

per second, re-entry vehicle weights of 9000 and 12, 000 pounds, and liftto-drag ratios of 0, 0. 25,. and 0. 5 are used. The results are plots of 

velocity, altitude, and density versus time. 

The ascent performance of an Apollo/MOL system, using the Titan 

IIIC and Saturn IB launch vehicles, has been evaluated. The Apollo abort 

recovery ceiling and re-entry performance have also been evaluated using 

the parametric studies of Sections 7.1 and 7. 2. The results are presented 

in graphical form. 

• 7-1 

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CONTENTS 

Page 

7. PERFORMANCE 7-7 

7.1 PRELIMINARY APOLLO /MOL RE-ENTRY VEHICLE 

ABORT BOUNDARIES 7-7 

7.1.1 Introduction 7-7 

7.1.2 Analysis 7-7 

7.1.3 Results 7-8 

7.2 PRELIMINARY APOLLO/MOL RE2p\ITRY TRAJECTORIES . 7-8 

7.2.1 Introduction 7-8 

7.2.2 Analysis 7-8 

-7.2.3 Results 7-9 

7.3 PRELIMINARY APOLLO /MOL ASCENT AND RE-ENTRY 

PERFORMANCE 7-10 

7.3.1 Introduction 7-10 

7. 3. 2 Titan IIIC Ascent Performance 7-10 

7.3.2.1 Method of Analysis 7-10 

7.3.2.2 Results 7-11 

7.3.3 Saturn IB Ascent Performance 7-12 

-7.3.3.1 Method of Analysis 7-12 

7.3.3.2 Results 7-13 

7.3.4 Apollo Abort Recovery Ceilings 7-14 

7.3.5 Apollo Re-entry Performance 7-15 

7.4 CONCLUSIONS. 7-15 

REFERENCES' 7-69 

7-3 

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TABLES 

Page 

7-1 Titan IIIC/Apollo Sequence of Events 7-16 

7-2 Titan IIIC/Apollo Sequential Weights 7-17 

7-3 Titan IIIC/Apollo Propulsion Data 7-19 

7-4 Saturn LB Sequence of Events 7-20 

7-5 Saturn IB Sequential Weights 7-21 

7-6 Saturn IB Propulsion Data 7-22 

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ILLUSTRATIONS 

Page 

7-:1 32° Inclination Apollo Abort Boundaries L/D = 0.5 7-23 

7-2 32° Inclination Apollo Abort Boundaries L/D = 0.4 7-24 

7-3 Re-entry Time History 7-25 

7-4 Re-entry Time History 7-28 

7-5 Re-entry Time History 7-31 

7-6 Re-entry Time History 7-34 

7-7 Re-entry Time History 7-37 

7-8 Re-entry Time History 7-40 

7-9 Re-entry Time History 7-43 

7-10 Re-entry Time History 7-46 

7-11 Re-entry Time History 7-49 

7-12 Re-entry Time History 7-52 

7- /3 Re-entry Time History 7-55 

7-14 Re-entry Time History . . . . . . 7-58 

7-15 Titan IIIC/Apollo Aerodynamic Characteristics . . . 7-61 

7-16 Titan IIIC Payload Capability Versus Circular Orbit 

Altitude 7-62 

7-17 Saturn IB/Apollo Aerodynamic Characteristics 7-63 

7-18 Saturn II3 Payload Capability 7-64 

7-19 Comparison of Apollo and Gemini B Abort Ceilings 7-65 

7-20 Launch Vehicle Apogee Traces and Apollo Abort Ceilings. 7-66 

7-21 Typical Apollo Re-entry Trajectory 7-67 

7-22 Apollo Landing Footprint 7-68 

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• 7. PERFORMANCE 

7.1 PRELIMINARY APOLLO/MOLL RE-ENTRY VEHICLE 

ABORT BOUNDARIES 

7.1.1 Introduction 

Abort boundaries corresponding to a maximum total load factor of 13 g's for 

the Apollo re-entry vehicle are presented in this section.. Re-entry :Vehicle 

weights of 9000 and 12,000 pounds were used with lift -to-drag ratios of 0.4 

and 0.5. The computed boundaries are shown in the accompanying plots. 

7.1.2 Analysis 

The two re-entry vehicle weights used here are the expected upper and lower 

limits of the weight for the Apollo /MOL configuration. The two lift-to-drag 

ratios are representative of the aerodynamic capabilities of the configuration. 

The following data and assumptions were used in computing these boundaries: 

1. Spherical, rotating earth; RE= 20.9029 x 106It, go= 32.2284 ft /sect 

2. ARDC 1959 standard atmosphere 

3. Initial latitude = 32 degrees, Initial azimuth = 90 degrees 

4. Constant mass (i. e., no inert weight loss due to ablation) 

5. Neither total heat nor heating rate were considered. 

6. Aerodynamic data (taken for Mach number 4.65) 

L/D = 0.4, CL= 0.46 = constant 

CD= 1.15 = constant 

LID = 0.5, C = 0.505 = constant L 

CD= 1.01 = constant 

Reference area = 133 ft2 

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7.1.3 Results 

Figure 7-1 illustrates the abort boundaries for both 9000 and 12,000 pounds 

with a LID of 0.5. The boundaries corresponding to the LID of 0.4 with the 

same weights are shown in Figure 7-2. 

These apogee boundaries represent upper limits which must not be exceeded 

by the ascent flight instantaneous apogees to ensure that the re-entry vehicle 

will not exceed the stated load factor during re-entry. The instantaneous 

apogee during the ascent results from the current booster velocity plus the 

velocity increment due to the abort rocket if it is assumed that the re-entry 

vehicle trajectory is not affected by the atmosphere between separation from 

the booster and attainment of the maximum altitude. Note that the ascent 

flight instantaneous apogee could be lowered by executing a pullover maneuver 

prior to exit from the atmosphere. The lower bound on the plots is merely 

the constant dynamic pressure line corresponding to the stated limit load 

factor. 

7.2. PRELIMINARY APOLLO/MOL RE-ENTRY TRAJECTORIES 

7. 2. 1 Introduction 

Re-entry trajectories of the Apollo re-entry vehicle from a 160-nautical 

mile circular orbit are presented in this report. The trajectories were 

obtained for deboost velocity increments of 500 and 1500 ft/sec using re-entry 

vehicle weights of 9000 and 12,000 pounds. Lift-to-drag ratios of 0, 0.25, 

and 0.5 were used. The time histories of velocity, altitude, and density are 

presented in graphical form for use in conducting heating analyses on the 

twelve .trajectories. 

7. 2. 2 Analysis 

Tee two re-entry vehicle weights used here are the expected upper and lower 

limits of the weight for the Apollo /MOL configuration. The three values of 

lift-to-drag ratios are taken to represent the vertical components of L/D where 

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• 

the vertical lift is modulated by banking the vehicle. In each of the cases, 

the drag corresponds to a total LID of 0.5. All of the trajectories were 

initiated by the application of a straight-back retro-thrust to deboost from 

the 160-nautical mile circular orbit. Velocity increments due to retro-thrusts 

of 500 and 1500 ft/sec were used here. 

The following data and assumptions were used in computing these boundaries. 

1. Spherical, ,rotating earth; RE= 20.9029 x 106 ft 

2. ARDC 1959 standard atmosphere; no atmospheric effects above an altitude 

of 400,000 feet. 

3. Trajectory computations were initiated at the 400,000-foot altitude and 

terminated at 70,000 feet. 

4. Initial latitude = 32 degrees, Initial azimuth = 90 degrees 

5. Constant mass (i. e., no inert weight loss due to ablation) 

6. Neither total heat nor 'heating rate were considered. 

7. Aerodynamic data (taken for Mach number 4.65) 

C

D = 1.01 = constant 

C

L = 0, 0.2525, and 0.505 for L/Dis of 0; 0.25, and 0.5, respectively 

Reference area = 133 ft2 

7. 2. 3 Results 

The velocity, altitude, and density time histories are presented in the accompanying plots. A summary of the input parameters is given in the following 

table. 

Figure No. Deboost %IV Weight L/D 

7-3a, b, c 500 9000 0.5 

7-4a, b, c 500 9000 0.25 

7-5a, b, c 500 

9 000 

0 

7-6a, b, c 500 12000 0.5 

7-7a, b, c 500 12000 0.25. 

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Figure No. Deboost AV Weight LID 

7-8a, b, c 500 12000 0 

7-9a, b, c 1500 9000 0.5 

7-10a, b, c 1500 9000 0.25 

7-11a, b, c 1500 9000 0 

7-12a, b, c 1500 12000 0.5 

7-13a, b, c 1500 12000 0.25 

7-14a, b, c 1500 12000 0 

For each figure number, the velocity, altitude, and density are plots a, b, and 

c, respectively. 

7.3 PRELIMINARY APOLLO/MOL ASCENT AND RE-ENTRY 

PERFORMANCE 

7.3. i Introduction 

Ascent and re-entry performance for an Apollo/MOL system has been evaluated 

and is reported in this section. The section is divided into four parts: the 

Titan IIIC ascent performance with an Apollo payload, the Saturn IB ascent 

performance, the Apollo abort recovery ceiling, and the Apollo re-entry 

performance. 

7.3.2 Titan IIIC Ascent Performance 

7.3.2.1 Method of Analysis 

Titan IIIC ascent performance is based on the Revision VI definition of the 

launch vehicle, Reference 7-2, with suitable corrections to account for the 

Apollo spacecraft payload. A sequence of events, a sequential weight statement, and a propulsion summary for this launch vehicle are presented in 

Tables 7-1,. _ 7-2 and 7-3. Note that the 7050-pound Apollo launch escape 

system is included in the weight statement and is jettisoned with the solid 

motors. The axial force coefficient data for the Titan IIIC/Apollo was obtained 

from the Fluid Mechanics Department (see Section 10. 3) and is presented in 

Figure 7-15. 

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The ascent trajectory consisted of a 10-second vertical rise followed by 

a kick (an instantaneous rotation of the missile attitude and velocity vector) 

into a gravity turn trajectory. After the solid motors are jettisoned, a 

constant inertial pitch rate is established for the remainder of the Core 

Stage I flight. A different constant inertial pitch rate is introduced at Core 

Stage II ignition and maintained to Core Stage'III (transtage) initial burnout. 

Values for the kick angle and the two constant pitch rates were automatically 

determined by the computer so that the burnout flight path angle was horizontal 

at a specified burnout altitude (60 or 100 nautical miles) and the aerodynamic 

heating index, $ q Vdt, was 0.95 x 108 - the Titan IIIC limit. Reference 7-3 

indicates that the latter constraint, fq Vdt = 0. 95 x 10

8

, maximizes the 

ascent performance, without violating the launch vehicle aerodynamic heating 

or dynamic pressure limits. 

Actual burnout of the transtage occurred when the velocity required to ascend 

to a desired orbit altitude was achieved. The payload weight was adjosted so 

that sufficient propellant remained in the transtage at the initial burnout in 

order to perform the final injection maneuver at the desired orbit altitude. 

In addition; .a propellant margin was included at the final burnout into orbit 

to account for -3cr dispersions in the laurich vehicle performance. This 

margin was based on the root-sum-square of three percent of the ideal 

velocity of the individual stages. 

7.3.2.2 Results 

Maximum payload capability as a functioh of final orbit altitude is presented 

in Figure 7-16 and is based on a 106-degree azimuth launch from ETR with 

an initial transtage burnout into a 60- or a 100-nautical mile perigee. At 

apogee, the transtage is restarted and injects the payload into the final 

circular orbit. 

The payload data presented in Figure 7-16 is based on optimum transtage 

propellant loading. The transtage has-a maximum capacity (or 22,841 pounds 

of usable propellant; however, off-loading between 4000 to 7500 pounds of 

7-11 

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transtage propellant may result in as much as a 2.3 percent improvement 

in payload capability. The optimum amount of propellant off-loading and 

resulting performance improvement vary with burnout altitude and final 

orbit altitude. 

The payload estimates for the Apollo/ MOL configuration are approximately 

700 pounds lower than those for Gemini/MOL. It is estimated that half of 

the difference is due to the inclusion of the 7050-pound launch escape system 

for the Apollo spacecraft. The remainder may be attributed to the increased 

drag resulting from the 13-foot diameter Apollo spacecraft. 

7.3.3 Saturn IB Ascent Performance 

7. 3.3. 1 Method of Analysis 

A sequence of events, a sequential weight statement, a propulsion summar y, 

and the aerodynamic drag characteristics for the Saturn IB launch vehicle 

are presented,in Tables 7-4, 7-5, 7-6 and in Figure 7-17, respectively, and 

are based on the data obtained from Reference 7-4. This configuration is 

characterized by a 32,500-pound minimum (-3o-) payload capability for a 

105-nautical mile circular orbit (based on a 72-degree azimuth launch from 

ETR.). 

The ascent trajectory consisted of a 25-second vertical rise followed by 

a kick (an instantaneous rotation of the missile attitude and velocity vector) 

into a gravity turn trajectory. Following ignition of the S-IVB stage, a 

constant inertial pitch rate is introduced and maintained for 260 seconds of 

the second stage flight. A different constant pitch rate is introduced at this 

time and maintained to burnout of the S-IVB stage. 

Values of the kick angle and the two constant pitch rates were automatically 

determined by the computer so that the burnout velocity was maximized (for 

a fixed payload weight) for a specified burnout altitude and a zero-degree flight 

path angle. Burnout of the S-IVB stage occurred with sufficient propellants 

remaining on board to account for -3o- dispersions in the launch vehicle 

performance. This propellant margin was based on the root-sum-square of 

2.333 percent of the ideal velocity of the individual stages. 

7-12 

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Maximum burnout velocity was- determined as a function of-payload weight 

for burnout altitudes of 60, 80 and 100 nautical miles. From this, apogee 

altitude could be calculated as a function of payload weight. Note that the 

Saturn burnout was into a perigee condition. At apogee, the propulsion system 

included in the Apollo/MOL payload is used for injection into-the final circular 

orbit. The LEM descent engine was assumed for this maneuver and is 

characterized by 10,500-pound thrust level and a 305-second vacuum specific 

impulse. 

7.3.3.2 Results 

Payload propellant expended to achieve the final orbit and payload weight in this 

final orbit are presented in Figure 7-18 as a function of the final orbit altitude. 

This data is based on a 106-degree azimuth launch from ETR and considers 

burnout altitudes of 60, 80 and 100 nautical miles. Note that the sum of the 

propellant weight expended and the payload weight in orbit constitute the.Saturn 

boosted payload weight. 

The performance data presented in Figure 7-18 is based on a Saturn M burnout 

into the perigee of an elliptic transfer orbit. At apogee, the payload propulsion 

system injects the payload into the final circular orbit. Another ascent mode 

requiring two injection maneuvers by the payload propulsion system was also 

considered. In this mode, the Saturn launch vehicle burnout is into a circular 

parking orbit. The payload propulsion is used to inject into the elliptic 

Hohmann transfer orbit and subsequently for injection into the final circular 

orbit. 

This type of ascent mode increases the payload weight in the final orbit by only 

300 pounds for a 300-nautical mile final orbit altitude, and by only 100 pounds 

for a 150-nautical mile orbit while approximately doubling the payload 

propellant requirement. While this ascent mode does offer a small payload 

gain, it has two disadvantages: (1) it requires an additional propulsive 

maneuver, and (2) the parking orbit ascent mode is undesirable at altitudes 

below 100 nautical miles. For these reasons, this ascent mode was dropped 

• from consideration. 

7-13 

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• 

There are several factors in addition to maximum payload capability that 

must be considered in selecting a Saturn burnout altitude. The burnout 

altitude must be high- enough so that the spacecraft will have sufficient life 

in orbit (in this case, the elliptic tranfer orbit) to effect a satisfactory 

recovery in the event of an abort. It is estimated that at least two or three 

orbits are required in order to accurately predict the spacecraft impact 

point. Secondly, the launch vehicle aerodynmic heating and load limits 

may determine a minimum burnout altitude, below which these constraints 

are violated. Neither of the above considerations has been examined in 

this study. 

7.3.4 Apollo Abort Recovery Ceilings 

Manned payloads generally limit the launch vehicle ascent trajectory by the 

requirement that the spacecraft and crew be recovered from any abort 

situation which could occur during the ascent to orbit. The abort recovery 

constraint is based on spacecraft design and crew tolerance limits and is 

usually presented in the form of an abort ceiling - a limiting curve of apogee 

altitude versus apogee velocity. Any combination of velocity, flight path 

angle, and altitude which has an apogee condition in excess of this ceiling 

will result in a spacecraft abort trajectory profile that exceeds the spacecraft design or crew acceleration tolerance limits associated with the abort 

ceiling. 

Abort ceilings for the Apollo spacecraft and the Gemini B spacecraft are 

compared in Figure 7-19. These curves are based on a 13-g maximum re-entry 

load factor limit and, in the case of Gemini, an afterbody temperature limit. 

In the region of interest, no afterbody temperature limit exists for the 

Apollo spacecraft. Apollo abort ceiling data was determined in Section 7. i 

for aerodynamic lift-to-drag ratios of 0.4 and 0.5 with re-entry. vehicle 

weights of 9000 and 12,000 pounds. It was found that these aerodynamic and 

weight variations resulted in only minor perturbations to the abort ceiling.. 

The more conservative ceiling based on the L/D of 0.4 and a re-entry vehicle 

weight of 12,000 pounds was used in Figure 7-19. 

7-14 uNCLASSURED 

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• The ascent trajectory apogee traces for the Saturn IB and Titan IIIC launch 

vehicles are compared to the Apollo abort ceiling in Figure 7-20. - Burnout 

altitudes of both 60 and 100 nautical miles-are considered in this figure.. The 

apogee trace is merely the locus of apogee altitude and apogee velocity points 

calculated each instant along the ascent trajectory. Note that the apogee trace 

for. the Saturn IB ascent trajectory to a 100-nautical mile burnout altitude 

exceeds the abort ceiling. This violation disappears when an abort ceiling 

based on a 10,500-pound re-entry vehicle weight is used. 

7. 3.5 Apollo Re-entry Performance 

The Apollo re-entry from orbit is initiated by firing the retro-rockets.. 

According to Reference 7-5, the nominal Apollo re-entry trajectory would 

have an effective bank angle of 45 degrees. If no lateral range is required, 

the vehicle is alternately rolled from left to right in order to maintain the 

trajectory in the orbital plane. A trajectory time history for this nominal 

re-entry is presented in Figure. 7-21 and was taken from the above reference. 

Lateral maneuvering is achieved by biasing the bank angle to one side. Maximum re-entry range is achieved by a zero-degree bank angle while minimum 

range is achieved by a 90-degree bank angle. The landing footprint for the 

spacecraft, also obtained from Reference 7-5, is presented in Figure 7-22 

for aerodynamic lift-to-drag ratios of 0.4 and 0.5. 

The Apollo spacecraft for the lunar mission has a trim lift-to-drag ratio 

of 0.5; however, removing unnecessary ablation material from the heat 

shield for orbital missions causes the spacecraft center of gravity to move 

aft and results in a trim lift-to-drag ratio of 0.4. 

Finally, the re-entry trajectories used for Figures 7-21 and 7-22 were based on 

a -2 degree re-entry flight path angle. Re-entry trajectories for other re-entry 

angles, spacecraft aerodynamics, and spacecraft wing loadings are presented 

in Section 7.2. 

7.4 CONCLUSIONS 

1. Within the conditions of the study, it appears that near-optimum ascent 

trajectories may be used with either Titan IIIC or Saturn IB as the launch 

vehicle. 

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2. Recovery conditions for the Apollo/MOL appear to be within acceptable. bounds. 

UNCLASSIFIED 7-15, 

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Table .7-1. Titan IIIC /Apollo Sequence of Events. 

Event Time, sec 

Liftoff 

End Vertical Rise, Start Pitch Over 

Solid Motor Web Burnout 

0 

10.00 

104.20 

Start Staging Sequence (SS) at a Sensed Axial SS 

Load Factor of 2.34 g's, Jettison Boattail 

Core Stage I Ignition SS + 1.00 

Solid Motor Burnout 114. 80 

Jettison Solid Motors and Launch Escape Tower SS + 10.00 

Core Stage I Burnout SS + 152.96 

Jettison Core Stage I and Core Stage II Ignition SS + 153.96 

Core Stage II Cutoff Signal, Start of Tailoff SS + 359.94 

Core Stage II Burnout and Jettison SS + 369. 94 

Core Stage III Ignition SS + 371.94 

Core Stage III Burnout SS + 806. 77(1) 

Note: . (1) Core Stage III burnout time based on complete 

depletion of usable propellants. 

7-16 

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Table: 7-2. Titan IIIC /Apollo 

Weight at Liftoff 

Solid Motor Propellant 

Solid Motor Expended Inerts(I) 

Sequential Weights. 

Component Weight 

(lb) 

Total Weight 

(lb) 

838,000 

6,75 

1,400,933 

Expended TVC Injectant(1) 21,594 

Core:Stage I and II Engine Bleed( 1) 36 

Core Stage I Boattail(2) 1,058 

Core Stage I Start Charge(3) 9 

Core Stage I Propellant(4) 14,921 

Weight at Solid Motor Jettison 518,563 

Solid Motor Burnout Weight 157,716 

Apollo Launch Escape Tower 7,050 

Weight After Solid Motor Jettison 353,7.97 

Core Stage I Propellant(4) 237,018 

Core Stage I and II Engine Bleed(5) 22 

Weight at Core Stage I Burnout 116,757 

Core Stage I Burnout Weight 15,647 

Core Stage II Start Charge 3 

Weight at Core Stage II Ignition 101,107 

Core Stage II Propellant 66,627 

Core Stage II Expenplqd Inerts 

and Engine Bleedtui 

67 

7-17 

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UNCILASa.q1111.111111111111111111111111 

Table 7-2 Titan IIIC/Apollo Sequential 

(Continued). 

Weight at Core Stage II Cutoff 

Weights 

Component Weight 

(lb) 

Total Weight 

(lb) 

34,413 

Core Stage II Shutdown Propellant 186 

Weight at Core Stage II Burnout 34,227 

Core Stage II Burnout Weight 6,455 

Control Module Settling Propellant 8 

Weight at Core Stage III Ignition 27,764 

Core Stage III Propellant 22,841 

Core Stage III Expended Inerts(7) 30 

Weight at Core Stage III Burnout 4,893 

Core Stage III Burnout Weight 2,531 

Control Module Weight at Burnout 2,362 

Payload Weight 0 

Notes: (1) Expended over 104.20 seconds of Solid Motor Web Action Time. 

(2) Jettisoned at a sensed axial load factor of 2.34 g's. 

(3) Expended 1 second after boattail jettison. 

(4) Core Stage I propellant expended both before and after solid 

motor jettison. 

(5) Expended at a constant rate over Core Stage I burning. 

(6) Expended at a constant rate over Core Stage II burning. 

(7) Expended at a constant rate over Core Stage III burning. 

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Table 7-3. Titan IIIC /Apollo Propulsion 

Solid Motors (All Data for Two Motors) 

Data(.'). 

Total Vacuum Impulse, lb/sec 219.59527 x 106 

Average Vacuum Specific Impulse, sec 262. 04686 

Impulse Propellants, lb 838";00Q 

Web Action Time, sec 104'. 20. 

Total Action Time, sec 114.80 

. Total Nozzle Exit Area, m.2 17,840 

Nozzle Cant Angle, deg 6.0 

Core Stage I, II and III 

Core Stage I II III 

Vacuum Thrust, lb 474,000 100,890 16,000 

Sea Level Thrust, lb 430,000 

Nozzle Cant Angle, deg, 2. 0 0 0 

Propellant Vacuum I 

sp

, sec 285.9 311.9 305.0 

Propellant Flow Rate, lb/sec 1657.797 323.469 52.459 

Total Flow Rate, lb/sec(2) 1657.942 323.794 52.528 

Notes: (1) All data along nozzle centerline, thrust levels 

should be corrected for nozzle cant angle. 

(2) Includes engine inert and engineJ bleed flow rates. 

7-19 

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Table 7-4. Saturn IB Sequence of Events. 

Event Time, sec 

Liftoff 0 

End Vertical Rise, Start Pitch. Over 25.000 

Shutdown S-IB Stage 4 Inboard Engines 147.328 

S-IB Stage Burnout 153.328 

Jettison S-IB Stage and S-IVB Stage Ignition 158.828 

Jettison Launch Escape System and 168.828 

Ullage Rocket Cases 

S-IVB Stage Burnout 627.153 (1) 

Note: (1) Complete depletion of S-IVB stage usable propellants. 

7-20 

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Table 7-5. Saturn IB Sequential 

Weight at Liftoff 

Weights. 

Component Weights 

(lb) 

Total Weight 

(lb) 

1,245,718 

Propellant Expended 883,178 

Frost, Fuel Additive, and Lube Oil(i) 1,640 

Weight at S-IB Stage Burnout 360,900 

S-IB Stage Burnout Weight 102,935 

S-IB/S-IVB Interstage 5,600 

S-IVB Ullage Rocket Propellant 182 

Weight at S-IVB Stage Ignition 252,183 

Ullage Rocket Cases(2) 213 

Launch Escape System(2) 6,600 

Propellant Expended(3) 219,871 

Weight at S-IVB Stage Burnout(3) 25,499 

S-IVB Stage Burnout Weight 22,939 

Vehicle Instrument Unit 2,660 

Payload Weight 0 

Notes: (1) Expended at a constant rate over 153.328 seconds of S-IB 

stage burning. 

(2) Jettisoned 10 seconds after S-IVB stage ignition. 

(3) S-IVB stage burnout weight should be adjusted so that sufficient 

propellant remains at burnout to provide a velocity margin against 

-3cr performance. This velocity margin should be equal to the 

root-sum-square of 2.333 percent of the ideal velocity of each stage. 

7-21 

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Table 7-6. Saturn IB Propulsion Data. 

S-IB Stage (i) S-IB Stage (1) 

Inboard Engines Outboard Engines • S-IVB Stage 

Vacuum Thrust, lb(2) 853,050 853,050 200,000 

Sea Level Thrust, lb(2) 752,000 752.000 

Nozzle Cant Angle, deg 3.0 6.0 0 

Flow Rate, lb/sec 2937. 500 2948.196 (3) 469.484 

Nots: (1) 

(2) 

(3) 

The S-IB stage has four H-1 inboard engines and four H-1 

outboard engines. 

Thrust data is along nozzle centerline and should be corrected 

by the cosine of the nozzle cant angle. 

Includes 10.696 lb/sec of frost, fuel additive, and lube, oil 

expended during S-lB stage burning. 

7-22 

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4 8 12 16 20 24 

VELOCITY, 1000 FT/SEC 

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• 

uNcLAsplf.tal 

I I I 

MAX LOAD FACTOR: 13 

L/Dt: 0.5 

R/V WEIGHT 

-- •----- 9000 

os s 

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ABORT 

APOGEE 

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// 

ft....ft......... .... 

se 

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Figure 7-1. 32° Inclination Apollo Abort Boundaries L/D = 0. 5 

7-23 

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2200 

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ALTITUDE,1000 FT 

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MAX LOAD FACTOR:13 (I's 

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Figure 7-2. 32° Inclination Apollo Abort Boundaries L/D = 0.4 

7-24 

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7-67 

LASSIFIED 

This document contains information affecting the national defense of the United States within the meaning of the Espionage Laws, Title 

18, U.S.C., Section 793 and 794, the transmission or revelation of which in any manner to on unauthorized pinyon is prohibited by law. 

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REFERENCES 

7-1 NASA TM X-603, "Static Longitudinal Aerodynamic Characteristics of a 

O. 07 Scale Model of a Proposed Apollo Spacecrat at Mach Numbers 

of 1.57 to 4.65," by James R. Morgan and Roger H. Fournier, 

September 1961 (Confidential). 

7-2 "Program 624A Standard Data Book for Vehicle Performance Calculations, 

Revision VI," TOR-269(4116-22)-4, 15 June 1964 (Unclassified). 

7-3 "Effect of Gemini B Abort Ceilings on Titan IIIC Ascent Performance," 

TOR-269(4535-15)-2, 29 May 1964, J. M. Dreyful3 (Unclassified). 

7-4 "Saturn IB Launch Vibliicle Design Data,: Revision VII," IN-P and VE-V63-1, 23 March 1964 (Confidential). 

7-5 "Extended Mission Apollo Study, Volume III," SID63-1370-3, 24 November 

1963 (Confidential). 

7-69 

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18, U.S.C., Section 793 and 794, the transmission or revelation of which in any manner to an unauthorized person Is prohibited by law. 
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SECTION 8 

TEST OPERATIONS 

L. Cooper 

SUMMARY 

This section contains a comparison of various significant test operations 

aspects of the Gemini/MOL and Apollo /MOL programs, as currently 

envisaged, covering the period from final countdown to re-entry. It considers 

the characteristics of the ground support network required for Apollo /MOL 

and possible conflicts with other programs in its use. These conflicts can 

occur throughout the Apollo /MOL programs, but only in the early part of the 

Gemini /MOL program. Also, the unmanned development flight requirements 

are examined, and it is estimated that one such flight is required, as compard 

with four for Gemini /MOL. 
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CONTENTS 

Page 

8. 1 INTRODUCTION 8-3 

8. 2 ASSUMPTIONS AND -CONSTRAINTS 8-3 

8. 3 TEST PLANNING CONSIDERATIONS 8-3 

8. 3. 1 Mission Control Center 8-3 

8. 3. 2 Trajectory 8-4 

8. 3. 3 Ground Support Network 8-4 

8. 3. 3. 1 Stations 8-4 

8. 3. 3. 2 Visibility Considerations 8-5 

8. 3. 3. 3 Security Considerations 8-5 

8. 3. 3. 4 Schedule Interference 8-5 

8. 3.4 Recovery Sites 8-6 

8. 3.5 Development (Unmanned) Flights 8-7 

8. 3.6 Earliest Manned Flight 8-7 

8. 3. 7 Launch and Orbital Operations 8-8 

8: 4 CONCLUSIONS 8-8 

8-2 

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• 

8. TEST OPERATIONS 

8. 1 INTRODUCTION 

This section is concerned with the test operations aspects of the Apollo/MOL, 

how they differ' from the Gemini/MOL and a comparison of these program 

considerations. 

The test operations area is considered to encompass flight operations, covering the period from final countdown to re-entry. 

8. 2 ASSUMPTIONS AND CONSTRAINTS 

In a preliminary study of this nature, any planning or definition must be based 

on a series of initial assumptions, since no portion of the system has been contractaully defined at this point. For the purposes of this report, the following 

assumptions have been made: 

1. Baseline configuration will be .the Saturn IB launch vehicle and the modified 

Apollo vehicle described in Section 1. 3.4, with minimum modifications. 

2. Payload capacity will be sufficient to carry all experiments on a single 

flight. 

3. Planning will be oriented to the earliest possible achievement of manned 

flight. 

4. Each experiment will be flown at least twice. 

8. 3 TEST PLANNING CONSIDERATIONS 

The following items delineate areas of primary consideration in flight test planning and comparisons of .these areas for the Apollo/MOL and Gemini/MOL 

programs. Summaries of these comparisons are presented in Table 8-1, Areas 

of Similarity, and Table 8r2, Areas ofiDi'ssiMilarity. 

8. 3. 1 Mission Control Center 

There are two possible choices for the location of the control center - the NASA. 

MSFC at Houston or the Mission. Control Center at CKAFS. The Gemini/MOL 

Program Office is currently studying these areas, and indications are that the 

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MCC at Cape.Kennedy will be selected. It should be noted that the considerations 

determining control center selection, such as interference with NASA operations 

and avilability of secure communications, are the same for both the Apollo/MOL 

and the Gemini/MOL programs. Therefore, this area does not affect the program 

comparison. 

8. 3. 2 Trajectory 

The trajectory considerations for both the Apollo/MOL and the Gemini/MOL 

programs are similar. Launch azimuths currently under consideration are 73 

and 106 degress, resulting in an orbital inclination of 32.5 degrees. Final 

selection of the launch azimuth will be based primarily on considerations of 

experiment, abort recovery, and range safety requirements. Injection into the 

final circular orbit, at an altitude of 160 nautical miles, will take place in the 

vicinity of Carnarvon, Australia. 

Since the trajectories are similar for both programs, this area does not affect 

the program comparison. 

8. 3. 3 Ground Support Network 

8. 3. 3.1 Stations 

Information supplied by the Gemini/MOL Program Office-indicates that current 

planning calls for a combined NASA/DOD ground network' comprised of the 

following stations: 

CKAFS 

Antigua 

Bermuda (NASA) 

Grand Canary (NASA) 

Carnarvon (NASA) 

Okinawa (to be implemented) 

Hawaii 

Vandenberg Tracking Station 

San Antonio 

Current planning for Apollo/MOL calls for utilization of the Unified S-Band 

communication system. Information, informally obtained,. indicates that the 

NASA Unified S-Band equipment is not compatible with the DOD Integrated S-Band 

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system. Therefore, it must be assumed that the Apollo/MOL program will 

utilize the NASA ground station network which will contain the following stations: 

Guam 

Hawaii 

Guaymas 

West Texas 

CKA FS 

Antigua 

Bermuda 

Ascension 

Carnarvon 

8. 3. 3. 2 Visibility Considerations 

An investigation of the Gemini/MOL network station visibility for each orbit 

is in progress, but has not been completed at this time. Preliminary results 

indicate that, while all stations do not see the vehicle on every orbit, the 

laboratory will be visible to at least one ground station on every orbit, for 

either the DOD/NASA network of the Gemini/MOL program, or the NASA Unified 

S-Band network of the Apollo/MOL program. 

8. 3. 3. 3 Security Considerations 

Since the MOL program is a military program, it would seem desirable, 

especially in view of the requirement for growth capability, to have secure 

communications available for the ground station network. The Gemini/MOL 

DOD/NASA network has this capability at the DOD stations, but not at the NASA 

stations. The Apollo/MOL network stations have no provisions for communication 

'security. 

8. 3. 3. 4 Schedule Interference 

The NASA Gemini flight test program .is scheduled for completion in. early 1967. 

Current Gemini/MOL schedules call for the first manned flight in mid-1967. 

If all schedules are met, there will be no ground station operational interference 

between the NASA Gemini and Gemini/MOL programs. Even if the NASA program 

slips, there will only be interference during the early Gemini/MOL flights. 

The NASA Apollo program will be in operation during the same time span (1968-

1970) as the Apollo/MOL program. Since the same ground stations will be used 

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for both programs, the probability of program interference must be considered. 

The schedules for Apollo and Apollo/MOL flight operations are based on the 

official NASA Apollo schedule, which contemplates four flights per year through 

1967 and six flights per year starting in 1968, and on amore probable Apollo 

schedule, determined from conversations with NASA officials, which contemplates a flight every four months. Based on the official schedule, and assuming 

30-day flight durations for the Apollo/MOL and 14-day flight durations for the 

NASA Apollo earth-orbit missions, there does not appear to be significant 

interference in ground station operations for the two programs.. This, however, does not allow for slippage in either program. The probable actual 

schedule indicates at least three Apollo/MOL flights on which both MOL and 

NASA Apollo vehicles will be in orbit at the same time supported by the same 

NASA ground stations. Schedule slippage could further aggravate this condition. 

Conversations with NASA Goddard SFC and NASA Hq. personnel indicate that 

turn-around time for a specific station (the minimum time required from lossof-sight of one orbital vehicle until a second vehicle can be acquired) could vary 

from a few minutes to a maximum of approximately half an hour. Therefore 

it seems possible to encounter orbits, when coverage is available from only 

one or two•stations due to visibility limitations, where ground station interference could result in complete loss of coverage for one orbit. This possibility 

could probably be reduced by careful scheduling, however. 

8. 3.4 Recovery Sites 

The minimum number and location of planned landing sites for a mission depends 

on such factors as the loiter capability, the orbital altitude and inclination, and 

the L/D ratio of the re-entry vehicle. For the purposes of this report, the 

orbital altitude and inclination can be considered the same for both programs, 

and eliminated from the comparison. The L/D for the Apollo/MOL re-entry 

vehicle is approximately 50 percent greater than for the Gemini/MOL vehicle. 

For a loiter capability of one orbit, the Gemini/MOL vehicle requires five 

planned landing sites, while the Apollo/MOL vehicle would probably require 

only four planned sites. If the loiter capability is increased to one half-day, 

the effect of the different LID is reduced, and the required number of landing 

sites is two for both programs. 

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• 8. 3. 5 Development (Unmanned) Flights 

Current Gemini/MOL planning, as stated by the Gemini/MOL Program Office, 

calls for four unmanned flights associated with the program. The first two 

flights will determine the effects of the addition of the laboratory on the 

Titan IIIC/Gemini configuration and will be instrumented to measure the flight 

environment. They are not considered part of the MOL program; they will 

utilize Titan III research and development launches which do not have assigned 

payloads. Although this provides a. substantial vehicle cost saving, it also 

allows possible conflict between Titan III research and development and MOL 

flight test objectives. The first two flights of the MOL program will also be 

unmanned and will verify the laboratory structure and the Gemini B capabilities 

(30-day life, crew safety and life support, etc. ). 

It is expected that an external configuration similar to the Apollo/MOL vehicle 

will be flown by the NASA Apollo program (Vehicle 206) in 1966 or 1967, prior 

to the first flight of the Apollo/MOL program. Since this flight, and subsequent 

• flights, will provide structural, flight control, and environmental data which 

is considered applicable to the Apollo/MOL vehicle, the two development 

flights which precede the Gemini/MOL program can be eliminated for the 

Apollo/MOL. In addition, since the laboratory structure will be based on the 

previously tested LEM adaptor structure, it seems likely that sufficient 

ground testing will reduce the second two development flights of the Gemini/ 

MOL to one flight for the Apollo/MOL program. On the other hand, this 

dependence on the NASA Apollo program for development data could result 

in program slippage if the NASA program slips significantly. In addition, 

present indications are that Saturn IB research and development flights will 

not be available to carry Apollo/MOL development payloads. Therefore, one 

or more additional Saturn vehicles will have to be purchased. 

8. 3. 6 Earliest Manned Flight 

According to the MOL Program Office, the first manned Gemini/MOL flight-is 

scheduled for mid-1967. The earliest time that a manned Apollo/MOL could be 

achieved is late 1967 or 1968, depending on the NASA Apollo launch schedule. 

It should be noted that the Titan III research and development program is 

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scheduled for completion approximately one year before-the Saturn IB research 

and development program, and one year before ,the first scheduled Gemini/MOL 

manned flight.. The Saturn IB research and development program is scheduled 

for completion less than one year.before the first ApollonvIOL manned flight. 

Therefore, schedule slippage in the Saturn IB program could have a greater. 

effect on the Apollo/MOL program than .a corresponding slip in the Titan. III 

research and development program would have on the Gemini/MOL program. 

8. 3. 7 Launch and Orbital Operations 

While the vehicles and, possibly, the crew size for the two programs under 

comparison vary considerably, the over-all launch operations and orbital 

operations (including range operations and housekeeping, maintenance, and 

experiment performance) considerations are, in general,' quite. similar. Therefore, this area does not affect.the program comparison. 

8.4 CONCLUSIONS 

1. Communication security may be a problem with Apollo/MOL because the 

NASA communication network will be used. 

2. It appears that Gemini/MOL might experience some ground station interference during early flights, while Apollo/MOL could experience some interference 

throughout the program. In both cases, the interference could probably be 

reduced, .if not eliminated, by careful scheduling. The effects of interference 

could be reduced or eliminated if the necessity for contact with..the orbiting 

vehicle at least once per orbit is modified. 

3. For short-time loiter capability, the Gemini/MOL configuration might 

require a greater number of planned landing sites than the Apollo/MOL. For 

longer loiter capability, the number of sites required is the same for both 

programs. 

4. The Gemini/MOL program requires two developmental flights (plus two 

preliminary "free" flights), while a total of one developmental flight is likely 

for the Apollo/MOL program. The Gemini/MOL might encounter interference 

problems with Titan III research and development objectives, and the Apollo/MOL 

program might be forced to slip if the NASA Apollo schedule slips. 

• 

8-8 
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• Table 8-1. Gemini /MOL Program and Apollo /MOL Program Comparison 

Test Operations Areas of Similarity. 

Mission Control Center 

Trajectory 

Launch Operations 

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616LAilair 

• 

SECTION 9 

RELIABILITY 

F. P. Klein 

SUMMARY 

Estimates of the reliability of the Apollo/MOL orbiting and re-entry 

system-s are made, using the standard exponential reliability model. They 

are combined with the Saturn 1B reliability, estimated in previous studies, to 

obtain an estimate of the overall reliability of the Apollo/MOL mission. 

Similarly, the reliability of the Gemini B and MOL laboratory are estimated 

and combined with that of the Titan IIIC to obtain the over-all Genaini/MOL 

reliability. The reliability estimates of the two systems are compared and 

found to be equal to within the accuracy of the estimates. 

9-I 

UNCLASSIFIED 

411111111111111111111 

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• 

CONTENTS 

9. RELIABILITY 

Page 

9-7 

9. 1 INTRODUCTION 9-7 

9. 2 MOL MISSION ANALYSIS 9-7 

9. 2. 1 Reliability Models 9-7 

9. 2. 2 Mission Definition and Time-Line Analysis 9-8 

9. 2. 3 Mission Success Definition 9-11 

9. 3 LAUNCH VEHICLES 9-11 

9. 4 GEMINI B/MOL RELIABILITY ANALYSIS 9-12 

9.5 APOLLO/MOL RELIABILITY ANALYSIS 9-14 

9. 6 CONCLUSIONS 9-17 

9.7 REFERENCES. 9-19 

• 

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1111101111111111111111111 

• 

TABLES 

Page 

9-1 MisM.on Profile 9-10 

9-2 Gemini B /MOL Reliability Analysis 9-13 

9-3 Apollo /MOL Reliability Analysis 9-16 

9-4 System Reliability 9-18 

• 

• 

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• 

9. RELIABILITY 

9. 1 INTRODUCTION 

The purpose of this study is to evaluate the capability of the proposed Apollo/ 

MOL configurations to achieve mission success and to compare the mission 

success reliability capabilities of the Apollo/MOL and the Gemini B/MOL. 

This section presents the prediction of system reliability for a 30-day MOL 

mission. The system as defined consists of launch-vehicle (all booster stages), 

spacecraft, and orbiting laboratory. The MOL mission as defined consists 

of powered flight, orbital insertion, orbital laboratory operation, de-orbit, 

re-entry, and landing. 

The study was limited to a reliability analysis of the launch vehicle, re-entry 

(command) module, and orbiting laboratory vehicle exclusive of experiments. 

The subsystems included in the laboratory vehicle analysis were limited to 

those necessary for normal vehicle operation (housekeeping) and service to 

the experiments. No maintainability capabilities or requirements were considered because the spares payload and available maintenance time have not 

been defined. 

Maximum use was made of previous reliability studies of the systems under 

consideration. The results of these studies were evaluated for applicability 

to the MOL mission configuration. In cases where subsystem configurations 

were identical, the reliability analysis was applied directly to the MOL 

evaluation. If the subsystem configurations were modified for the MOL 

mission, a new reliability model was constructed which incorporated all 

reliability data available from the previous studies. 

9.2 MOL MISSION ANALYSIS 

9. 2. 1 Reliability Models 

The reliability of an item may be represented by the exponential model 

R = e-

>t, where X is the failure rate and t is the operating time, if the failure rate is assumed to be constant during the operating time. This model 

9- 7 

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4 Equipment 2 

Equipment 3 

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is a useful approximation of complex equipment reliability•if the item being 

modeled does not have redundant elements. The reliability of a series of 

such 

independent. The product of subsystem reliabilities is used to determine the 

system reliability for the MOL mission. 

Where redundant equipment provides alternate paths to accomplish subsystem 

mission objectives, the reliability model must allow for this. In this case, 

we are concerned with the probability that either one or the other or both 

equipments will function and may be represented by the model R = 1 

-(1-e-X1t) (1-e-X2t) providing that exponentiality is assumed for each equipment. 

For a configuration such as: 

Equipment I 11...M.Imel•••••■•• 

Where 2 and 3 provide a redundant path, the system reliability may be represented by the-model, 

- R = e-X1t

E 1-(1-e -X2t) - x3t)1. 

If equipments 2 and 3 are identical and one is turned off (in -standby) when the 

other is operating, then the model will be 

R = t(1 + X2t), where X2 = X3* 

9-8 

UNCLASSIFIED 

0:7 

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MA. 

This assumes that the equipment has zero failure rate when inoperative. This 

assumption is made for all reliability analyses performed during this study. 

This type of modeling is an integral part of a reliability analysis, and sometimes is the most difficult to accomplish due to incomplete system description. 

The equipment must be defined so that all redundant paths are considered, the 

operating times of the individual equipment must be determined, and failure 

rate information must be available which is appropriate to each equipment in 

the operational environment. 

9.2.2 Mission Definition and Time-Line Analysis 

The basic MOL mission is defined to operate for 30 days at an orbital altitude 

of 150 to 250 nautical miles. Using these ground rules, a -simplified mission 

profile was developed and is shown in Table 9-1. The sources- of information 

for the mission profile are Reference 9-1, 9-2, and 9-3. The time required 

for each phase is an optimum figure which does not include holds for unscheduled events. Only the major events are considered for a simplified mission 

analysis. 

Each event is examined to determine the subsystems which are required for 

successful operation. An analysis of the subsystem operating -requirements 

is used to provide information to determine the operating time of each subsystem during the entire mission. Some subsystems are assumed to operate 

continuously throughout an entire mission phase, while other subsystems 

are assumed to operate in a regularly scheduled intermittent fashion throughout an entire phase. In order to simplify the reliability calculations for timedependent subsystems, simplified operating times were derived. Subsystems 

which were assumed to operate continuously throughout the entire mission 

were required to operate for 725 hours. 

UNCLASS1FOED 

9- 9 

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, . • . . r11 0 

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Time from 

L/0 to End 

of Phase 

(Hours) 

Lift off, powered flight, 10 1. 0 

Coast Orbital Injection 

1. Ascent & 

Injection 

Establish orbit, active 25 3. 5 

lab housekeeping equipment, enter lab. 

2. Pre-Laboratory Operational Orbit 

720. 0 

2.0 

723..5 

725. 5 

3. Laboratory 

Operational 

Orbits 

4. Pre-Separation Orbit 

Commence lab. operation, de-activate unused 

S/C equipment, perform 

experiments 

Shut down lab, activate 

all S/C equipments 

5. Separation 1. 7 727. 2 

and Re-entry 

Separate from S/C, 

attain proper attitude, 

fire retro-rockets, 

maintain re-entry 

attitude, deploy recovery 

equipment, land 

9-10 

CL SSJF 

• 

• 

Time in Phase 

Phase Events (Hours) 

Table 9-1. Mission Profile 

tt 

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9.2.3 Mission; Success Definition 

Reliability is defined as the probability that a device will give- satisfactory 

performance without failure for a given period of time when used under 

specified conditions. This study is concerned with a total mission time of 

approximately 725 }tours, and satisfactory performance is assumed to be 

successful launch-, 30-day orbital operation, and safe landing .of the re-entry 

vehicle. This is an extremely simplified definition of mission-success because certain failures may occur during the mission and be corrected by 

proper maintenance procedures. A failure which is corrected does not constitute system failure. However, this simplified definition of mission 

success does consider any failures as a deviation from satisfactory perf6rmance. Because oftheobvibus.paradox, the definition for'mission success 

used is only a firstapproximation of the over-all effectiveness of a MOL 

mission. Howeve-r, it is extremely useful for a preliminary reliability 

analysis and when relative values of reliability are more important than absolute. The reliability figures for each subsystem provide a comparison of 

the complexities of the various subsystems and indicate possible problem 

areas. The -same -comparison can be made between the two systems under 

consideration. These preliminary reliability predictions which are 'presented 

here may be used in reliability/maintenance trade-off studies to determine 

the spare parts' requirements and the subsystems which are expected to 

require the most maintenance. From these trade-off studies, the effectiveness of the competing designs may be compared to determine which system 

has the higher probability of completing the mission, regardless of the state 

of reliability of the system. 

9.3 LAUNCH VEHICLES 

Studies (Reference 9-4 and A Study By The Aerospace Reliability. Department) 

have been performed which compare the predicted reliability growth of the 

Saturn IB and the Titan IIIC launch vehicles. Reference 9-4 predicts a reliability of 0.82 for the Saturn IB and 0.77 for the Titan. IIIC by 1967, a difference 

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of approximately five percent. The earlier Aerospace study - predicts a difference of approximately five percent, with the Titan IIIC as the higher. Reference 9-5 presents a NASA estimated reliability growth curve'for the Saturn IB 

which shows a 0.90 reliability by 1968. Another Aerospace study predicts a 

reliability of at least 0.84 for theTitan IIIC by the start of 1967. 

9.4 GEMINI B/MOL RELIABILITY ANALYSIS. 

The reliability analysis of the Gemini B/MOL was performed using information from three studies: the MAC 706 Study; a report from McDonnell 

Aircraft Co. describing the use of the Gemini B spacecraft for a MOL 

mission; and a typical MOL Laboratory configuration described in a report 

from the MOL Program Office. Each of the reports was analyzed, and 

sufficient information was available to make a reliability-prediction for each 

of the subsystems which were expected to comprise the Gemini B/MOL 

System. The subsystem reliability predictions were divided into separate 

predictions for the laboratory, described by the McDonnell reports. The 

subsystem and equipment failure rates which were used in the reliability 

models of the various subsystems were provided by McDonnell. The MOL 

laboratory configuration described by the Aerospace MOL Program Office, 

utilized typical McDonnell subsystems and failure rates. 'The failure rates 

assigned by McDonnell are representative of the state-of-the-art equipment 

and provide a reasonable estimate of the reliability of the particular equipments and subsystems. The reliability block diagrams and assumptions 

which we're used by McDonnell, and the Aerospace MOL Program Office, 

are consistent with good reliability evaluation techniques and present a conservative prediction of system reliability. In some cases, the mission time 

used for the reliability calculations provided by Aerospace was longer than 

the assumed: 30-day mission. When this occurred, it was necessary to 

recompute the reliability for a 30-day mission assuming an-exponential 

reliability model. The result of this reliability analysis is shown in 

Table 9-2; the over-all Gemini B/MOL spacecraft reliability is approximately 0.88 for a 30-day mission. 

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Table 9-2. Gemini B/MOL Reliability Analysis 

Reliability 

Gemini B 

S/C 

Communications 0. 99996 

Telemetry 0.9989 

Power 0.999314 

Environmental Controls 0.9999 

Re-entry Control 0.999 

Earth Landing 0.9963 

Sequentials 0.9991 

Guidance and Control 0.999 

Sub Total: (0. 9915) 

Laboratory Power 

Environmental Control 

Instrumentation and Telemetry 

Reaction Control 

Attitude Control Electronics 

Communications 

0.9844 

0.9675 

0.9975 

0.991 

0.985 

0.960 

Sub Total: (0. 890) 

Gemini B/MOL Space Vehicle Reliability= 0.882. 

9- 13 

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9.5 APOLLO/MOL RELIABILITY ANALYSIS 

The reliability analysis of the Apollo/MOL was performed using information 

from North American studies supplied to Aerospace by NASA. The information, consisting of three separate Apollo configurations, was from the MODAP 

study (a logistics vehicle), the XMAS study (an extended orbital-mission), 

and the standard Apollo Lunar configuration. The baseline Apollo/MOL configuration chosen for the Aerospace study is described as Configuration II 

in the North American XMAS study. This configuration consists• of the 

standard Apollo command module, the service module, and an orbital laboratory utilizing the lunar excursion module housing. 

In Configuration II, the orbital laboratory utilizes several of the Apollo subs ysterns to provide housekeeping equipment. These subsystems include 

communications, telemetry and data, environmental control, service module 

reaction control, stabilization and control, and power. Because this configuration shares subsystems between the laboratory and the command module, 

a single analysis was performed which includes both the orbiting laboratory 

and the command module re-entry vehicle. The subsystem and equipment 

failure rates which were used in the reliability models of the various subsystems were obtained from the North American reports. 

Some of the failure rates and failure probabilities given in the North American 

reports appear to be more optimistic than comparable failure rates provided 

by McDonnell. However, they represent comparable state-of-the-art equipment. The -subsystems were not always described completely by reliability 

block diagrams, and in some cases it was necessary to assume a subsystem 

reliability configuration and model. Because of this lack of complete documentation, the reliability calculations which went into the Apollo/ MOL 

reliability analysis lack the degree of confidence which may be placed in 

the Gemini B/MOL prediction. In some cases. the mission times used for 

the reliability calculations given by North. American were 14 days for the 

Apollo lunar mission and 90 days for the XMAS mission. When this occurred, 

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it was necessary to recompute the reliability for a 30-day mission assuming 

an exponential reliability model. The results of this reliability analysis is 

shown in Table 9-3 and provide an over-all Apollo/MOL spacecraft reliability 

of approximately 0.87 for a 30-day mission. 

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Table 9-3. koollo/MOL Reliability Analysis 

Subsystem Reliability 

Command Module Reaction Control 0. 99915 

Service Module Reaction Control 0. 991 

Stabilization and Control 0. 94463 

Environmental Control 0. 974 

Guidance and Navigation 0. 98 

Power (Assumes 2 Spare Fuel Cells) 0. 99925 

Cryogenic Storage 0. 9999 

Earth Landing (Assumed) 0. 999 

Data and Telemetry 0. 9852 

Communications 0. 99977 

Ascent Survival (Assumed) 0. 995 

Apollo/MOL Space Vehicle Reliability = O. 87333 

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9.6 CONCLUSIONS 

A comparison of the reliabilities of the competing systems is shown. in 

Table 9-4, which includes the probability of successful booster operation. 

Each system has a probability of completing a 30-day mission of approximately 0. 7 with less than 1 percent difference in the two estimates. It is 

concluded from the results of this analysis that neither system provides an 

advantage when only the reliability of the systems are compared. This is 

true when the probability of failure-free operation is the only consideration. 

However, as was previously explained, it is realistic to consider the effects 

of maintenance upon the over-all probability of continued system operation. 

The reliability figures presented here can be assumed to be the first 

approximation of the relative reliabilities of the individual subsystems and 

may be used as a basis for further reliability/maintainability trade-off 

studies. 

The reliability analysis of the Gemini B/MOL is considered firmer because 

of the greater credibility of the reliability information provided by McDonnell. 

This does not mean that the information provided by North American is in 

error; but is not so complete as the information provided by McDonnell, and 

consequently we have less confideb.ce in. the North American figures. It is 

believed that the failure rate information used as an input to the Gemini B 

analysis is more conservative, and it is concluded that there is a greater 

expectation of achieving the Gemini B/MOL reliability prediction. 

9-17 

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Gemini B/MOL 

Table 9-4- System Reliability 

Apollo/MOL 

Titan II1C 0.80 Saturn IB - .0.80 

Lab 0.89 Apollo S/C & 

Lab (Concept II) - 0.87 

Gemini B S/C 0.99 

Mission Rel. : 0.705 Mission Rel. : 0.696 

Assumptions 

1. Mission success requirements are for 30-day orbital life and safe 

landing of the re-entry vehicle. 

a. Booster reliability is assumed to be at least 0.80 for both launch 

vehicles by 1967. 

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REFERENCES 

9-1. Informal Report on Gemini B Applications to MOL MisSion, compiled 

from McDonnell. Aircraft Corporation Internal Technical Memorandums issued during First Six Months of 1964 (Confidential). 

9-2. "Final Report Program 706, Phase 0, Volume III: Analysis of Performance and Requirements," McDonnell Aircraft Corporation, 

15 November 1963 (Secret). 

9-3. "Extended-Mission Apollo Study (XMAS;),, Final Report, Volume I: 

Summary," North American Aviation, 24 November 1963 

(Confidential). 

9-4: "Report on a System Comparison and Selection Study of a Manned Orbital 

Research Laboratory, Volume I: Technical Summary, Appendix 

19, Safety, Reliability and Maintainability,' Douglas Missile and 

Space Systems Division, September 1963 (Secret). 

9-5. "Modified Apollo Logistic Spacecraft Study (MODAP), Final Report, 

Volume VII: Operational Implementation," North American 

Aviation, 30 December 1963 (Confidential).. 

9-19 

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a • C 

SECTION 10 

FLUID MECHANICS 

N. R. O'Brien 

M. J. Adams 

SUMMARY 

The effect of the retropack on re-entry stability and trim 

characteristics is computed to allow for the case of retropack hang-up. 

Considering the possibility of utilizing the Apollo command 

module for other missions tha.n the moon flights, the following questions 

were raised: 

1. What are the capabilities and limitations of the Apollo 

earth landing system? 

2. What are the possible ways of increasing the allowable 

landing weights? 

From available information, an increase of the allovable 

landing weight of the command module is possible if (1) a new earth 

landing system design is allowed,, and (2) increased space and weight 

is provided for the new earth landing system. The new earth landing 

system would require some combination of new drogue chutes, main 

parachutes, and/or impact attenuation system. The degree of change 

would depend on the extent of the weight increase and the capabilities 

of the final design for the Apollo earth landing system. 

A new impact attenuation system would most likely be the 

first new item. Adequate impact attenuation could be provided by a 

system composed of retro rockets (velocity attenuation) and shock 

struts (possibly with an extended heat shield); however, other techniqUes 

and combinations are possible. 

Aerodynamic characteristics of Titan III C, with an 

appropriate bulbous payload, and of the Apollo re-entry vehicle are 

presented for use in performance calculations. 

10-1 

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CONTENTS 

Eau 

10. FLUID MECHANICS 10-7 

10.1 EFFECT OF RETROPACK ON APOLLO HYPERSONIC 

TRIM CHARACTERISTICS 10-7 

10.2 CAPABILITIES OF THE APOLLO EARTH LANDING 

SYSTEM FOR HANDLING INCREASED WEIGHT 10-9 

10.2.1 Description of Apollo Earth Landing System 10-9 

10.2.2 Effects of Increasing Allowable Landing Weights on 

Terminal Velocity Parachute Parameters 10-10 

10.2. 3 Maintaining the Descent Velocity of 24 FPS. by 

Changing Parachute Characteristics 10-10 

10.2.4 Capabilities of the Apollo Earth Landing System 10-12 

10.2.5 Major Problem Areas with Apollo Earth Landing System . . 10-12 

10.3 MISCELLANEOUS.FLUID MECHANICS DATA 10-16 

10.4 CONCLUSIONS 10-18 

REFERENCES 10-19 

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TABLES 

Page 

10-1 'Earth .Landing System Design Criteria 10-13 

ILLUSTRA TIONS 

Page 

10-1 Effect of Retropack. on Static Stability and Trim 

Characteristics• of Apollo Re-entry Configuration 10-8 

10-2 Descent Velocity Versus Command Module Gross 

,Weight for Apollo• Earth Landing Main Parachute 

Clusters 10-11 

10-3 Parachute Diameter Versus Command. Module 

Gross Weight for Different Main. Parachute 

Clusters 10-11 

10-4 Normal-Mission g Limits 10-15 

10-5 Titan III - Booster. Plus 156 Inch Diameter 

Bulbous. Payload 10-17 

10-5 

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10. FLUID MECHANICS 

10.1 EFFECT OF RETROPACK ON APOLLO HYPERSONIC TRIM 

CHARACTERISTICS 

This section presents the pitching moment coefficient versus angle-of-attack 

in response to a request for the effect of a possible hang-up of the Apollo retropack at re-entry. 

The solid curve shows the experimental pitching moment coefficient of the 

basic Apollo re-entry configuration (without retropack) as obtained from 

Reference 10-1. The dashed curve represents the pitching moment coefficient 

curve fOr.the Apollo with the retropack, and was obtained by adding a 

theoretiCal (Newtonian) increment for the retropack (Reference 10-2) to the 

experimental data (solid curve) for the basic configuration. The results 

shown were referenced to the offset center of gravity indicated in Figure 10-1. 

The change in trim angle of attack (i. e., at Cm CG = 0) due to the presence of 

the retropack is of the order of two degrees for the indicated configuration. 

The retropack tends to reduce the required trim angle of attack. Without the 

retropack, (LID)trim = 0.45 at &trim= 29 and with the retropack, 

(LID)trim = 0.43 at Atri.

m 

= 27o. 

The net change in pitching moment and trim is due. to the two components: 

(a) The retropack cylinder with front face excluded, and 

(b) The part of the Apollo spherical surface shielded by the retropack. 

The contribution of the retropack cylinder excluding the front face reduces the 

pitching moment (i. e., nose up), but this loss is compensated by the stabilizing 

moment of the shadowed segment of the Apollo face. The effect of the retropack is sensitive to location of the center of gravity offset since this moment 

is largely a function of the induced CN of the retropack cylinder and the loss in 

C D due to shadowing of part of the face of the Apollo by the cylinder at angle 

of attack. In other. words, the change in trim angle shown on the, figure is valid 

for the indicated center of giavity location and would vary somewhat with varying 

center of gravity location. The effect of the front lace of the cylinder was neglected 

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140.11111111111111ft 

10 20 30 40 50 60 

RETROPACK 

\ 

IN. 

7 IN. 

E1 CG 

37 

\ 

\ 

WITHOUT 

\ 

\ 

RETROPACK 

WITH RETROPACK 

____) 

\ 

\ 

s 

..... 

so' 

se 

se 

ANGLE OF ATTACK, DEG 

Figure 10-1. Effect of Retropack on Static Stability and Trim 

Characteristics of Apollo Re-entry Configuration. 

10-8 

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-0.04 

0 

• 

2 

C.) PITCHING MOMENT COEFFICIENT , 

0.10 

0.08 

0.06 

0.04 

0.02 

0 

-0.02 
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• 

since it was assumed to have about the same contribution as that part of the 

Apollo face to which. it is attached and the effect of that part of the face of the 

Apollo.has been reflected in the. basic..data. (solid. curve). 

10.2 CAPABILITIES OF THE APOLLO EARTH LANDING SYSTEM 

FOR IIA.NDLING INCREASFTIZWEIGHT 

10.2. 1 Description. of Apollo Earth .Landing,.S.ystem.. 

The Apollo' earth landing system consists of a stabilization and deceleration 

subsystem, recovery subsystem, impact attenuation subsystem, 'and three subordinate subsystems. Together these subsystems provide for recovery of the 

Apollo_c.omm_and_module.any time.from.pad abort through mission completion. 

10.2.1. 1 Stabilization and Deceleration Subsystem 

At 25,000 feet altitude a baroswitch closes, jettisoning the apex cover.

1 Two 

seconds later two ribbon-type parachutes are deployed in a reefed condition by 

mortar action. After six seconds, the drogue chutes are disreefed to their full • 13.7-foot diameter. At 15,000 feet altitude, a baroswitch closes causing the 

disconnection of the two drogue chutes and the firing of the mainechute pilot 

mortars. The drogues can be deployed at higher altitudes by crew command for 

stabilization of descent if the return flight becomes unstable. 

10.2.1.2 Recovery System 

The three main parachutes are 88-foot diameter ringsails each deployed by its 

own 7.2-foot diameter flat ringslot pilot chute. The three pilot chutes, each 

deployed by a mortar, inflate and deploy the three main chutes in a reefed condition. The main chutes disreef six seconds later by pyrotechnic cutter. The 

spacecraft descends at 25 fps, and at landing the main chutes are disconnected by 

crew command or inertia switch in unmanned flights. For abort conditions, the 

sequencing is different depending on the altitude, and an escape rocket, tower, 

and canards are utilized. 

This operation may no longer be necessary; the apex cover may have been 

jettisoned with the escape tower. 

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10. 2. 1. 3 Impact Attenuation Subsystem 

The Apollo impact attenuation system consists of a crushable toe and crushable 

seat struts for the command module, thus requiring a 30° pitch angle at impact. 

10.2.1.4 Supplementary Subsystems 

These are the sequence controller, ordnance and pyrotechnic devices, and 

search and location aids. 

10.2.2 Effects of Increasing Allowable Landing Weights on Terminal 

Velocity Parachute Parameters 

Figure 10-2 illustrates the change in descent velocity with increasing weight of 

the Apollo command module using the present main parachute system. However, 

not all of the weight range may be available. Once a parachute has been 

designed for a given recovery weight, the strength of the parachute defines the 

allowable weight increase because opening loads will increase with increasing 

descent velocities. For instance, if the design strength of the present main 

parachute is utilized during deployment, then only by allowing the factor of 

safety to be reduced could the gross weight be increased. For such a case the 

elimination of a factor of safety of 1. 5 would allow the weight to increase from 

9000 to 13,500 pounds. For an increase in weight of 5000 pounds an increase of 

5.6 fps in descent velocity occurs. This may seem to be a small penalty; however, 

a higher velocity at a higher weight loading means a much larger increase in 

energy to be absorbed by the impact attenuation system. Since there now is some 

indication that the present impact attenuation system for the Apollo command 

module is marginal, any increase in weight would most likely require a new 

design for the impact attenuation system. 

10.2.3 Maintaining the Descent Velocity of 24 FPS by Changing 

Parachute Characteristics 

Figure 10-3 illustrates for three different main parachute clusters the parachute 

diameters required for various weights maintaining a descent velocity of 24 fps. 

While this figure shows how to maintain a specific descent velocity, any increase 

in weight increases the energy to be absorbed by the impact attenuation system. 

Thus, redesign of the impact attenuation system may be inevitable with any 

increase in command module weight if the system is marginal. 

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10 II 12 13 14 15 16 

GROSS WEIGHT, 1000 LB 

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28 

a_ 

lL 

>: 

3 

g 26 ta

to La 

24 

3 —88.1 FT DIAMETER RINGSAIL PARACHUTES 

30 

Figure 10-2. Descent Velocity Vs Command Module Gross Weight 

for Apollo Earth Landing Main Parachute Clusters 

3— RINGSAIL 

PARACHUTES 

3 — 25° CONCI 

FULLY 

EXTENDED 

AL 

— 

I 

SKIRT 

PARACHUTES 

4 — RINGSAIL 

PARACHUTES 

DESCENT VELOCITY OF 24 FPS 

16 

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22 

9 

• 

PARACHUTE DIAMETER , FT 

709 

10 11 12 13 

GROSS WEIGHT, 1000 LB 

Figure 10-3. Parachute Diameter Vs Command Module Gross 

Weight for Different Main Parachute Clusters 

14 15 

110 

100 

90 

80 
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UNCLASSIFIED 

While the 25° conical fully extended skirt parachute shows better drag performance 

than a ringsail parachute, the stability of a ringsail parachute is better than that of 

a 25° conical fully extended skirt. 

10.2.4 Capabilities of the Apollo Earth Landing System 

Table 10-1 and Figure 10-4 from Reference 10-3, indicate the earth landing system 

design criteria for the command module of the Apollo Program. This system meets 

the following gross requirements after normal re-entry or during abort: 

a. Stabilization 

b. Velocity control 

c. Impact attenuation 

d. Land or water touchdown 

To meet similar gross requirements for increased command module weights of a 

couple of thousand pounds is within the state of the art, but would require another 

comprehensive design and development program. Both the drogue chutes and main 

parachutes would require re-sizing and possibly a change to a different canopy 

design to maintain the same descent velocities. The impact attenuation system 

would require a new design to meet similar g-load requirements. There is a tradeoff between the main parachute system (rate of descent) and the impact attenuation 

system; however, the consideration of minimum weight would most likely make the 

utilization of the combination of velocity and impact attenuation system more 

desirable. For instance, a velocity attenuation consisting of retro rockets could be 

activated prior to ground impact reducing the command modules final rate of 

descent. If the Apollo impact attenuation system of crushable toe and crushable 

seat is not adequate, then the heat shield could be designed to extend and shock 

struts used to absorb the impact. 

10.2.5 Major. Problem Areas with Apollo Earth Landing System 

Major problem areas associated with the design and operation of the deceleration, 

stabilization and recovery systems are: 

a. Recovery system weight and its effect on command module stability. 

b. The inter-relation of command module dynamic stability and drogue 

parachute design. 

c. The problem of main parachute non-uniform cluster operation. 

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Ta

ble 10- 1. Earth Lan

ding SystemDesignCrite

ria 

• 

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20 G/1000 G/SECOND 

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THAT ILLUSTRATED IN THIS SECTION 

Figure 10-4. Normal-Mission g Limits. 

10- 15 

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d. The design loads for the parachute attachment points on the command module. 

e. The marginality of the present impact attenuation system. 

10.3 MISCELLANEOUS FLUID MECHANICS DATA 

Figure 10-5 shows the estimated drag of the Titan IIIC at A = 0 for Mach Numbers 

from zero to four. This estimate pertains to the configuration shown in the 

figure and is based on wind tunnel tests of a bulbous payload fairing. It is 

recommended that, in order to minimize adverse transonic aerodynamic 

effects due to the hammerhead payload, the command module should be followed 

by a cylindrical section terminated by an abrupt ( A 30°) reduction in diameter to 

the 120 inch core dimension. The drag characteristics shown in Figure 10-5 are 

adequate for performance calculations for all such configurations with or without 

the Apollo escape tower. 

Hypersonic aerodynamics of the command module at trim angle of attack 

(Reference 10-1) were supplied for the re-entry analysis presented in Section 5. 

The center of gravity of the capsule is assumed to lie 8 to 10 inches off the axis 

of symmetry and no further than 49 inches aft of the blunt face of the heat shield. 

This location produces acceptable vehicle dynamics (according to'NAA) and hypersonic L/D between 0.4 and 0.5. It is not yet known if the MOL version of the 

capsule can be balanced in this range. 

The question of aerodynamic control of the re-entry trajectory was discussed 

with NAA during inspection of the mock-up. The current plan is to use the 

lifting capability of the capsule during return from the lunar mission only as 

a means of staying within the entry corridor defined by overshoot, heating and 

deceleration limits. However, in return from earth orbit, this capability 

could be used to steer out impact dispersions due to deboost and atmospheric 

uncertainties with a resulting decrease in the required recovery force. 

Investigation of this possibility might be desirable in the Aerospace study if 

cost or availability of the recovery force becomes critical. 

10-16 

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10.4 CONCLUSIONS 

For the specified location of the center of gravity, the retropack would have 

a very small effect on the static stability and trim characteristics of the 

re-entry vehicle. 

The Apollo Earth Landing. System cannot handle significantly increased 

command module weights without a comprehensive additional design and 

development program. 

10-18 

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REFERENCES 

10-1 "Static Longitudinal Aerodynamic Characteristics of a 0.07 Scale 

Model of a Proposed Apollo Spacecraft at Mach Numbers of 

1. 57 to 4.65," by J.R. Morgan and R.H. Fournier, NASA 

TM X-603, dated September 1961. 

10-2 "Equations and Charts for the Evaluation of the Hypersonic Aerodynamic Characteristics of Lifting Configurations by the 

Newtonian Theory," by E. L. Clark and L. L. Trimmer, 

AEDC TDR 64-25, dated March 1964. 

10-3 "Modified Apollo Logistics Spacecraft Study," North American 

Aviation SID 64-14614, Final Report, Vol. IV, Earth 

Landing System, 30 December 1963. 

• 

10-19 

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SECTION 11 

PROPULSION 

M. J. Russi 

SUMMARY 

The propulsion requirements are studied for abort, de-orbit, crew 

module reaction control, laboratory reaction control, and laboratory main 

propulsion. Appropriate propulsion systems which will meet the various 

requirements are identified as tentative choices. Specific propulsion devices 

currently under development are chosen in each case, but none of them (with 

the possible exception of the escape tower) can be considered "off the shelf" 

for the MOL mission. In each case, additional development and qualification 

will be required to meet the MOL requirements. As the mission duration is 

extended, the confidence in these choices decreases. 

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CONTENTS 

Page 

i i , i INTRODUCTION 11-7 

11.2 ABORT PROPULSION 11-7 

11.3 DE-ORBIT PROPULSION 11-9 

11.3.1 Ground Systems 11-12 

11.3.2 Orbital Storage 11-12 

11.3.3 Firing Temperature 11-12 

11.3.4 Qualification 11-12 

11.4 REACTION CONTROL SYSTEMS 11-13 

11.4.1 Crew Module 11-13 

11.4.1.1 System Capacity 11-13 

11.4.1.2 Module Control 11-13 

11.4.1.3 Space Storability 11-15 

11.4.2 Laboratory Module - (MOL) 11-15 

11.4.2.1 Design Basis 11-15 

11.4.2.2 Thruster Lifetime 11-17, 

11.4.2.3 Space Storability 11-17, 

11.5 MAIN PROPULSION 11-17, 

11.5.1 Requirements 11-18 

11.5.2 Candidate Propulsion Systems. 11-18 

11.5.3 Main Propulsion Selection 11-20 

11.6 AREAS REQUIRING FUlit.THER STUDY 11-20 

11.7 CONCLUSIONS 11-21 

11.7.1 Escape Tower. 11-21, 

11.7.2 De-Orbit 11-21. 

11.7.3 Crew Module Reaction Control System 11-21, 

11.7.4 Laboratory Reaction Control System 11-21 

11.7.5 Main Propulsion 11-22 

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ILLUSTRATIONS 

Page 

11-1. Abort Propulsion 11-8 

11-2. Retrograde Propulsion Requirements 11-10 

11-3. Retrograde Propulsion Configuration 11- 11 

11-4. Apollo Command Module Reaction Control System 

Description 11-14 

11-5. Apollo Service Module Reaction Control System 11-16 

11-5 

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11. PROPULSION 

11.1 INTRODUCTION 

In choosing propulsion systems for Apollo/MOL, the NAA extended Apollo 

Configuration II was used as a‘baseline. This configuration makes every 

effort to retain current propulsion subsystems. The functions to be performed by propulsion include,abort, de-orbit, crew module reaction control, 

laboratory reaction control, and laboratory main propulsion for final orbit 

injection and maneuvering. 

11. 2 ABORT PROPULSION 

The various abort propulsion possibilities for the Apollo/MOL are depicted 

.schematically in Figure 11-1. For abort within the atmosphere, the present 

. launch escape tower is. used and appears to be satisfactory. It can be used 

..up to an altitude of about 300, 000:feet and is. then jettisoned.. 

For:abort outside. the atmosphere,..the present technique- is, to use .the 21, 900-

pound thrust service module engine. In .the NAA studies, the LEM descent 

engine (10, 500-pound thrust), was chosen to replace.the 21., 900-pound thrust 

motor for main propulsion and abort. The .main consideration was a minimum 

acceleration of 0.3:to 0. 4 g established.by NAA for abort, which eliminated 

lower thrust,motors. However, NAA did. not consider the possibility of using 

. the solid retro motors ,as posigrade .abort. propulsion. In this case, a 

minimum thrust of 3500 pounds would be required since only :.the crew module 

would. have, to be accelerated. The NAA retropack configuration ,(described 

_below) has a thrust level of 6600 pounds and burning -time of 34 seconds in 

salvo which .appears. to be:more than adequate for the abort requirements: 

A trade-off study is .required to determine the relative advantages of using 

either or both the retro motors and the main propulsion engine. for posigrade 

abort outside the atmosphere. 

11-7 

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LAUNCH ESCAPE MOTOR 

TOWER JETTISON MOTOR 

PITCH MOTOR 

UNCLAbvt-altu 

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0 © 8 

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AbortProp

ulsion. 

LEM DESCENT ENGINE 

SOLID RETROS (PROPOSED FOR DEBOOST PHASE) 

EXO

ATMOSPHERIC ABORT - USEMAIN ENGINE OR SOLID RETRO ENGINES (TOWER JE

TTIS

ONED) 

ATMOSPHERICABORT- USE ESCAPE .TOWER 
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• 

11.3 DE-ORBIT PROPULSION 

The two alternatives considered for providing retrograde propulsion are 

(1) the use of the main liquid propulsion engine (service module) and 

(2) the use of a newly designed solid propellant system (as in Mercury and 

Gemini). The latter approach is preferred for several reasons; lower 

vehicle weight, higher reliability and its capability .to also perform the 

abort function. 

Lower vehicle weight results- from the fact that a solid motor system would 

have to de-orbit only the crew capsule whereas the main propulsion system 

would have to de-orbit the service module as well. Furthermore, the 

relatively small .6V required for de-orbit and the relatively small difference 

in specific impulse between the liquid propellant system and the solid propellant system tend .to favor the latter. The retrograde velocity increment 

requirement involves a trade-off between system weight and landing accuracy. 

Figure 11-2 shows NAA data, for estimated heat shield weight and retro 

rocket weight versus retro velocity increment. The current Apollo heat 

shield weight is about 1300 pounds and is overdesigned for re-entry from 

low orbital altitude.. Several hundred.pounds weight could be shaved from 

the heat shield.,but it would reqUire:Teq_ualincatibn and additional flight tests. 

From Figure 11-2, for ,the current heat shield, the combined weight is a 

.minimum for a retro velocity increment of about 300 fps. However, from 

Figure 11-2 which also shows touchdown range sensitivity vs. retro velocity 

increment, it can be seen that a minimum retro velocity increment of. 

500 fps. is needed to avoid large landing inaccuracies (NAA). For this 

reason 500 fps was selected as a design. point. 

The NAA design approach for the retro-pack is shown in Figure 11-3 

together with data for the selected solid motor and an overall weight estimate. 

Six spherical'inotors (XM-85 flight proven) are used in a. cluster. Five 

motors are required and an additional one is included for redundancy. 

The use-of a cluster of solid motors for de-orbit is recommended rather 

than the liquid propellant main propulsion system. In addition to a lower 

weight system, the former approach offers greater reliability. Although 

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3000 

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Figure i-Z. Retrograde Propulsion Requirements. 

it-to 

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6 SPHERICAL SOLID 

MOTORS 

(5 REQUIRED PLUS 

1 REDUNDANT) 

BATTERIES , 

UNCLASSIFIED 

MOTOR DATA 

DESIGNATION XM- 85 

THRUST 1100 LB 

BURN TIME 34 SEC 

SPECIFIC. IMPULSE 270 SEC 

PROPELLANT WEIGHT 136 LB 

TOTAL- WEIGHT 155.6 LB 

LENGTH 17.3 IN. 

DIAMETER 17.2 IN. 

WEIGHTS (NAA) 

STRUCTURE . 102 LB 

MOTORS 936 LB 

BATTERIES 232 LB 

TOTAL 1,270 LB 

Figure 11-3. Retrograde Propulsion Configuration. 

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41111111111111.1111 

neither liquid propellant engines nor solid propellant motors have been 

stored in space for 30 days or more, it is likely that the solid motors will 

be easier to store. If need be, the solid motors. can be completely sealed 

until needed for re-entry. On the other hand, the liquid propellant main 

propulsion would be used for orbit injection and on-orbit maneuvering and 

would be subject to prior malfunction and leakage. Furthermore, the 

multiple solid motor concept offers some redundancy and can accept a 

single motore failure. It has been used successfully on Mercury, will be 

used on Gemini and appears to be a logical choice. There are; however, 

several areas regarding the selection of solid motors for de-orbit which 

require further study: 

11.3.1 Ground Systems 

The solid motors will have to be installed during vehicle buildup. Usual 

procedures prefer to leave the propellant loading until final checkout to keep 

the vehicle in an inert state: The motor selected by NAA (XM-85) has a 

beryllium loaded propellant which may not be compatible with safety regulatiOns. An aluminized version of the motor exists (NOTS-100A) but would 

require modification to reach the total impulse of the XM-85. 

11.3.2 Orbital Storage 

The effects of long term space storability (30 days +) on solid propellant 

motors are not fully understood. Temperature cycling, vacuum, and 

radiation effects must be explored. 

11.3. 3 Firing Temperature 

The total impulse produced by a solid motor is affected by its firing temperature and landing accuracies are sensitive to motor total impulse. An 

active temperature control will probably be required to keep the propellant 

within acceptable temperature limits. Emergency return to earth may 

require continuous temperature control. 

11.3.4 Qualification 

The selected solid motor (XM-85 or equivalent) will probably require requalification as a result of any design changes and confirmation complete retro system 

compatibility with launch vehicle, operational environment and performance. 

11-12 

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11.4 REACTION CONTROL SYSTEMS 

11.4.1 Crew Module 

The Apollo command module reaction control system (RCS) is shown 

schematically in Figure 11-4, together with pertinent characteristics.. The 

system consists of two independent equally capable, identical systems, each 

with six thrusters. Both systems function in normal operation. The design 

is such that failure of either one of .the systems.. is acceptable and the mission 

can be completed with the remaining system. The system, which is sealed 

until time for re-entry, provides 3-axis re-entry orientation of the command 

module after separation from the service module. 

The Apollo command module RCS is basically suitable for the Apollo/MOL 

application. However, there are several considerations which may require 

design changes that must be investigated. 

11.4.1.1 System Capacity 

The current Apollo command module RCS is designed for use only during 

re-entry. Use in the Apollo/MOL mission will require control before, 

during, and after retro firing prior to re-entry. Additional propellant will 

be required of .the RCS for offsetting thrust misalignment during retro firing. 

Compensating for this, to some extent, is the shorter re-entry time from 

earth orbit than from lunar return. In any case, the total propellant requirement must be re-examined for the Apollo/MOL command module. If the 

propellant quantity is increased, the thrust chamber rated lifetime must be 

increased also which will require re-qualification to the new duty cycle. 

11.4.1. 2 Module Control 

The current Apollo command module RCS was designed for use without a 

retrorocket attached. The .thrust level and location of the reaction jets will 

have to be re-examined to determine adequacy of the system during firing 

of the retrorockets (thrust misalignment) and after retropack jettison 

(center of gravity shift). Re-sizing the thrust chambers would require 

additional development and re-,qualification. 

ii-i3 

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ABLATIVETHRUST CHAMBERS 

RATED DURATION-130 SEC 

THRUST - 91 LB EACH PRESSURE FED WITHHELIUM 

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"1111111111111.11MPI 

• 

11.4.1.3 Space Storability 

The current Apollo systems are designed for 14 day storage in space. 

Propellant pressurization system leakage rates must be examined for the 

extended duration. Temperature control of propellant and valves is 

necessary to keep the system in a ready condition. An active thermal 

control system is probably required (electrical heaters have been suggested 

by NAA). The problem of storability increases with time and if growth 

beyond the 30 day mission is anticipated, initial design review studies 

must allow for it. 

11.4.2 Laboratory Module - (MOL) 

The Apollo service module RCS is the basis for a minimum modification 

approach to a MOL RCS.and is shown schematically in Figure 11-5. Also 

listed are the system weight and characteristics. Basically, the service 

module RCS consists of four independent, identical clusters of four 

thrusters each located as shown. The hypergolic propellants, N204/AZ-50, 

are pressure fed by helium and expelled by Teflon bladders. The .thrust 

chambers are fabricated from coated molybdenum and are cooled through 

radiation. 

The above system appears to most nearly meet the MOL requirements. 

However, there are several factors which need additional study before it 

can be concluded that the system will be suitable for the MOL mission: 

11.4.2.1 Design Basis 

The current Apollo service module RCS design is based on a very small 

percentage of limit cycle operation. The design was influenced strongly 

by rendezvous and translation requirements. As a result, the thrusters 

were sized (100 pound thrust) for the latter requirements and the minimum 

impulse bit which is important for limit cycle propellant consumption was 

not optimized. The main function of the MOL RCS will be to provide long 

term limit cycle operation and the non-optimized minimum impulse bit 

may be too costly in terms of propellant consumption, particularly for 

11-15 

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Ap

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ntrol System. 

a) 

1• 

- 100 LB ( EACH) 

- 838 LB ( TOTAL4 SYSTEMS) 

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U 

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RATED DURATION - 1000 SEC 

INERTWEIGHT 

PROPELLANT 

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WT 
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missions longer than 30 days. A smaller thruster (s,maller impulse bit) 

may be required: for limit cycle operation which would require further 

development and qualification. 

11.4. 2. 2 Thruster Lifetime 

The current service module RCS is designed for 10, 000 on-off cycles and 

1000 seconds firing time. The MOL requirements will exceed the lifetime 

requirements ,for the current Apollo service ,module RCS because of the 

longer mission time. The radiation cooled thrust chamber =is potentially 

capable of longer lifetime than required by the current Apollo mission but 

considerable development and qualification will be required. 

11.4. 2. 3 Space Storability 

The Apollo service module RCS is designed for a 14-day mission. The 

effects of pressurization gas leakage rates will have to be determined for 

longer missions. In addition, the current propellants must be kept within 

rather narrow temperature limits .to remain in the liquid phase and within 

the design pressure limits. Electrical heaters have been considered by 

NAA for this purpose and a. 150 watt requirement has been estimated. The 

Teflon bladders used for positive expulsion of the N204/AZ-50 propellants 

may be marginal for a 30-day mission. Other techniques such as metallic 

bellows should be examined.. 

11. 5 MAIN PROPULSION 

For the concept under consideration, a propulsion system larger than the 

RCS is required to place the Apollo/MOL on station and to provide on-orbit 

maneuvering such as rendezvous .or as dictated by on-board. experiments. 

In the NAA Extended Apollo approach, the service module,main propulsion 

system (21, 900 pound thrust) is replaced by the smaller, variable thrust 

LEM descent engine which has a maximum thrust of. 10,000 pounds. Iri 

this study the requirements are first examined, candidate propulsion systems 

are considered, selection of the most promising, system is made and problems 

associated.with it.are discussed. 

11-17 

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11. 5. 1 Requirements 

The assumption is made that the Apollo/MOL is boosted into an elliptical 

orbit. The MOL main propulsion system is then used to circularize the 

orbit at apogee. This requires a AV of about 200 - 250 fps. No restart 

is required and the thrust level should be several thousand pounds to 

minimize engine burn time (lifetime) and gravity losses (propellant). For 

rendezvous, the NAA study indicates additional AV of 1300 - 1400 fps is 

required and multiple restart capability is necessary. 

In summary, preliminary estimates for main propulsion indicate a 

desired AV capability of about 2200 fps. A thrust of several thousand 

pounds is desired and multiple start capability with accurate control 

appears mandatory. 

11. 5. 2 Candidate Propulsion Systems 

.The desire for multiple starts and mission flexibility indicates a liquid 

propellant rocket for main propulsion. The basic ground rule of minimum 

modification and high confidence level limits the number of engines to be 

evaluated to a relative few. Cryogenic propellants offer the highest specific 

impulse and the RL-10 engine (LOX/LH2-. 15,000 pounds thrust) is a 

potential candidate. However, little performance gain would be realized 

at the low AV level required and extensive redesign of the Apollo service 

module would be required to use it. The agena engine (IRFNA/UDMH - 

16,000 pounds thrust-pump-fed) would also require extensive service 

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module re-design. The remaining candidates are the current Apollo service 

module engine, LEM descent engine, LEM ascent engine and the Titan III 

transta,ge engine. The characteristics of the above engines are listed below: 

CANDIDATE ENGINES 

APOLLO LEM LEM T -III 

SM , DESCENT ASCENT TRANSTAGE 

Propellants 

Thrust, Lb. ' 

N20

4

/AZ

-50 

21,900 

N 0 ./A 2 4 

10,500 * 

-50 N 0 /A -50 2 4 Z 

3500 

N20

4

/AZ

-50 

8000 

Rated Burn Time, Sec. 730 730 385 440 

Specific Impulse, Sec. 307 305 305 300 

Restart YES YES YES YES 

Engine. Weight, Lb. 692 360 145 200 

Length, In. 160 75 -- 80 

Diameter, In. 100 54 31 49 

10 to 1 throttling available 

The service module engine size andweight make it unattractive to this application 

since high thrust is not required. Its use would obviously require the least 

modification to the service module but would severely restrict the volume and 

weight available for MOL related equipment. 

The LEM descent engine is about 1/2 .the size and weight of the service module 

engine. It has adequate thrust and its throttling capability is very attractive for 

rendezvous and other maneuvers requiring accurate control. Relatively minor 

modification to the service module will be required to incorporate the LEM 

descent engine. 

The LEM ascent engine is .the most compact of the candidate engines. However, 

its thrust level is marginally small for orbit changes and it would exceed its 

rated burn time for AV's greater than 1500 fps. 

The. Titan III transtage engine is adequate as far as thrust level, size and weight 

are concerned. It is not as easily integrated with. other NASA systems. 

11-19 

UNCLASSIFIED 
11. 5. 3 Main Propulsion Selection 

From the above discussion it appears that the NAA selection of the LEM 

descent engine is a reasonable choice. Its reduced size and weight allows 

room in the service module for MOL equipment. It offers the greatest 

flexibility (throttling) for the MOL mission and will be limitedly qualified 

in the current Apollo program. 

11.6 AREAS REQUIRING FURTHER STUDY 

The full growth potential offered by the liquid propellant engine depends on 

its capability to be stored in space for long periods of time while providing 

many restarts. The current Apollo program will not qualify or demonstrate 

space storability of the LEM descent engine beyond 14 days. Furthermore, 

this storability will be in an unfired condition, i. e. , the engine is not fired 

until near the end of its storage time. The MOL mission requires the main 

propulsion system to put the MOL on station and then to be stored for 30 days 

or more during which time it will be fired several times. It is possible that 

initial MOL missions could be programmed to perform all main propulsion 

functions immediately following final orbit injection thereby avoiding long 

time system storage problems. However, this approach would limit the 

flexibility and usefulness of MOL. 

The problems to be solved for the LEM descent engine long space storage 

are somewhat common to the reaction control system. Since the propellants 

are the same, the thermal environment must be controlled to present freezing 

or tank overpressure. Propellant orientation control will be necessary under 

zero "g" conditions to minimize center of gravity shift and to aid in propellant temperature control. Present techniques (bladders) have shortcomings 

for long term storage systems. Passive systems such as surface tension 

control are promising but need further development. 

The lifetime of thrust chamber ablative materials under extreme vacuum 

conditions at high temperature is not well known. This condition occurs on 

engine shut down in _vacuum.. LikewiSe, the lifetime-Capability of coated 

refractory metals used for radiation cooled thrusters is not well known for 

extreme vacuum conditions. 

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11.7 CONCLUSIONS 

While specific propulsion systems currently under development have been 

identified for each MOL requirement, none of these (with the possible 

exception of the escape tower) can be considered "off-the-shelf" systems 

for the MOL mission. The degree of confidence with which each selected 

propulsion subsystem can be made to meet MOL requirements, with additional 

development and qualification, decreases with the amount of time it is 

required to function under space environment. The following conclusions 

hold for the 30-day mission. 

11.7.1 Escape Dower 

Since the escape tower for providing abort within the atmosphere is only 

required to operate for a very short time, the current Apollo system is satisfactory for MOL. 

11.7.2 De-Orbit 

The XM-85 flight proven solid motor appears to be the best choice for the 

retro pack. Some development is required to incorporate it into the necessary 

cluster of six motors and to provide necessary thermal control for the 30-day 

mission. 

11.7.3 Crew Module Reaction Control System 

The current Apollo command module RCS appears to be basically satisfactory 

for MOL. The system is not required to operate until just prior to re-entry 

and can be stored in a passive state. The thrust level and location of the 

thrusters will have to be examined to insure adequate control during the retro 

firing. 

11.7.4 Laboratory Reaction Control System 

The current Apollo service module RCS appears to be the best choice. However, 

additional development and requalification of the system will be required to 

provide the longer life and space storage required in the MOL mission. 

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11. 7. 5 Main Propulsion 

The LEM descent engine represents a reasonable choice for MOL main 

propulsion. Additional development and requalification will be required to 

insure adequate space storability for the MOL missions. 

11-22 

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SECTION 12 

SOLID MECHANICS 

J. Hook 

D. E. Hargis 

G. B. Fox 

A. J. Victor 

SUMMARY 

The following comments on the structural feasibility of the proposed 

Apollo/MOL configurations are submitted. Based on the results of the NAA 

XMAS study, the Saturn LB configurations appear to be structurally feasible 

in all respects. In the Titan IIIC configuration, the spacecraft structures 

appear to be feasible with the weights that were considered. The launch 

configuration appears to be structurally feasible as far as most loads are 

concerned, but wind tunnel tests would be required to determine the effects 

of wind-induced oscillations and transonic buffeting. 

L

12 

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CONTENTS 

Page 

12. SOLID MECHANICS 12-3 

12.1 TITAN IIIC CONFIGURATION 12-3 

12.1.1 Spacecraft 12-3 

12.1.2 Booster 12-3 

12.1.2.1 Ground Winds and Wind-Induced Oscillations 12-4 

12.1.2.2 Transonic Buffeting 12-4 

12.2 SATURN IB CONFIGURATION 12-5 

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12.1 TITAN IIIC CONFIGURATION 

12.1.1 Spacecraft 

A preliminary design study has been performed to ascertain the structural feasibility of the proposed Apollo/MOL spacecraft configuration. Critical loads were 

estimated and the structure of the "service" module between the transtage and 

the command module was sized accordingly. A conservative analysis indicates 

that the allowance, provided in Section 2 of this report, of 1400 pounds for the 

structural portions of this module should be entirely adequate. In fact, this 

weight might be significantly reduced by a structural optimization procedure. 

A comparison of the critical loads used in this study with those given in the NAA 

XMAS Study indicates that the command module also will be structurally adequate 

for this mission. 

12.1.2 • Booster 

Several loads studies have been conducted by Martin-Marietta Corporation 

(Denver) during the Task #2 MOL Payload Constraints effort which involved a 

matrix of 35 configurations with payload ranges between 15K and 25K, laboratory 

diameters between 120 inches and 156 inches, and payload lengths between 

30 feet and 50 feet. In addition, Aerospace Corporation has performed a loads 

study for a 156-inch diameter, 40-foot payload length, and 21,000 pound payload 

on.Titan IIIC. Based on the results of these studies, it is felt generally that 

loads imposed on the Titan IIIC by the 154-inch diameter/21K Apollo/MOL 

configuration will be well within the Titan IIIC design allowable loads. 

With regard to dynamic loads, however, there are two important areas of 

uncertainty that will require testing before their effects can be known with 

confidence. These are (1) ground winds and wind-induced oscillations and 

(2) transonic buffeting. 

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12. 1. 2. 1 Ground Winds and Wind-Induced Oscillations 

Results from a 7. 5 per cent 624A Ground Wind and Wind-Induced Oscillation 

Wind .Tunnel Test indicated a critical load condition and large dynamic response 

of the vehicle for one of the bulbous configurations tested in the presence of the 

ITL electrical masts. Since the ground winds problem is extremely sensitive 

to changes in the upper body (payload) configuration, and the interference 

effects of the ITL masts on the vehicle response are not known, it will be 

necessary to conduct additional ground winds tests on the Apollo/MOL/ Titan III 

configuration. The Titan III 7. 5 per cent booster model, .ITL -electrical masts, 

ITL umbilical tower, transporter stand, and turntable should already be available for .this test. Construction of a 7. 5 per cent Apollo/MOL capsule will be 

required. In order to utilize the existing model and turntable, the test must 

be conducted in the NASA Langley 16-foot transonic dynamic tunnel. 

12. 1. 2. 2 Transonic Buffeting 

Transonic buffeting can generally be divided into two areas of concern. Local 

effects are usually related to buffet excitation frequencies above 20 cps and 

apply to skin panel response and component reliability. Gross vehicle effects 

are related to overall vehicle response to buffet excitation in the frequency 

range below 20 cps. It is now known, based on wind tunnel tests, that bulbous 

payload shapes introduce large gross vehicle buffet loads. In the Task #2 

MOL Payload Constraint loads studies noted above, the portion of the total 

load due to buffet was based on -T -III wind tunnel buffet tests and T-II Gemini 

wind tunnel buffet tests; however, considerable interpretation and modification 

of the test data was required to provide buffet loads for these studies. Due to 

the possible severe buffet loads induced by the bulbous Apollo/MOL configuration, verification of the buffet loads by wind tunnel tests should be included 

in the Apollo/MOL program. 

The above comments apply to a configuration without stabilizing fins. We do 

not have enough information to enable us to assess the loads situation if fins are 

to be used. The necessity for fins and their effect on structural loads can be 

determined only through a further study program. 

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It has been possible to obtain very little information on the loads -analyses performed by. NAA in connection with the -T-IIIC booster. The data we do have seem 

to agree generally with Aerospace results generated for similar configurations 

(e. g. , bending moment diagram PS 7061.2 of NAA document PS 64-117, "Apollo 

Applicability:to MOL"). 

12.2 SATURN IB CONFIGURATION 

The configurations studied for, Apollo/MOL are within the-envelope of those 

investigated extensively by. NAA in the•. XMAS. Study, •in which feasibility was 

established for a mission having requirements generally similar to those of 

Apollo/MOL. For the purposes of the present study, it can therefore be 

concluded that the Apollo/MOL/SIB configuration is feasible from a structural 

standpoint. 

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