JPRS ID: 8845 TRANSLATION COSMODROME BY PROFESSOR A.P. VOL'SKIY

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BY PROFESSOR R. P.C ~JOL' SK I Y ~ ~

7 JANUARY 1980 CFOUO) ~ 1 OF 4
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JPR~ L/8845
7 January 1980
- Transla~ion
- Cosmodrome
By Professor A. P. Vol'skiy
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JPRS L/8845
7 January 1980
COSMODROME
Moscow KOSMODROM in Russian 1977 signed to press 23 Mar 77 pp
1- 312
Book edited by Professor A. P. Vol'skiy, Voyenixdat Publishing
House 10,000 copies
~
CONTENTS PAGE
Annotation � 1 ~
From the Authors 2 _
Introduction 3
Chapter 1. General Information About the Space Rocket Complex 6
1.1. Cosmodrome 6
1.2. Space Rocket System 24 -
1.3. Main Cosmodromes of the World 36
Chapter 2. Engineering Complex 49
2.1. Purpose, Structure and Composition 49
2.2. Testing Booster Rockets and Space Vehicles 59 ~
2.3.. Means of Assembling Space Rocket Systems 65 -
2.4. Engineering Complex for the "Saturn-V-Apollo" Space
Racket System 69
Chapter 3. Launch Complex 75
3.1. Purpose, Structure and Composition 75
3.2. Operations Performed at the Launch Comple~~ g(
3.3. American Launch Complexes 89
- a - [I - USSR - A FOUO]
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Chapter 4. Special Technological Equipment 99
4.1. Means of Transporting the Space Rocket Systems to -
the Cosmodrome 99
4.2. Transport Equ~pment 107
4.3. Lifting-Erecting Equipment 109
~ 4.4. Launch Systems 112
~ 4.5. Service Means 118
4.6. Electr:Lcal Equipment 126
Chapter 5. Fueling Systems 133
5.1. General Information 133
5.2. Grounc~ Fueling Systems 141
5:3. Fueling Systems for Cryogenic Fuel Components 148
5.4. Systems for Filling with High-Boiling Fuel Components 173
5.5. Gas Supply Systems 183
Chapter 6. Thermostating Systems 195
6.1. Purpose, Structure and Composition 195
6.2. Classification of the Systems 197
6.3. Sources of Cold and Heat 198
- 6.4. Structure of the Thermostating Systems 201
Chapter 7. Communications of the Ground Systems with the
On-Board Systems ("Ground-On-Board" Communications) 213
7.1. Nature of the "Ground-0n-Board" C;immunications 213
7.2. Standard "Ground-On-Board" Comm.unications Layouts 218 '
7.3. "Ground-On-Board" Communications of the "Saturn-V-
Apollo" Space Rocket~System 222
Chapter 8. Guidance Systems of the Space Rocket System 226
8.1. General Information 226
8.2. Basic Devices of the Guidance System 229
8.3. Nonautomated Guidance Systems 233
8.4. Automated Guidance Systems 236
Chapter 9. Monitoring and Control Systems for Technological
Process Operations 240
9.1. General Information 240
9.2. Purpose of the System 245
9.3. Classification of Systems 248
9.4. EstimatiQn of the Eff iciency of the ASPA 253
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Chapter 10. Space Center Monitoriz~g and Control Systems 258
10.1. Basic Characteristics of the Ob~ect of Monitoring and
Control 25g
10.2. Automatic Systems 264
10.3. All-Purpose Systems 269
10.4. Functional Monitoring Systems Q 273
10.5. Interaction of the Monitoring and Control Systems 277
10.6. Telemetering Systems 281
10.7. Cable Communications 284
Chapter 11. Control of the Space Rocket Complex 287
~ 11.1. Organization of Control 287
` ~ 11.2. Human Operator in the Control Process 289
11.3. Information Display and Communications During Launch
Control 294
. Bibliography 29$
Books Which Will be Publ~shed by Voyenizdat on Rocket
Engineering and Radioelectronics in 1978 300
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PUBLICATION DATA
English title : COSMODROME
Russian title . KOSMODROM
Author (s) �
Editor (s) : A. P. Vol'skiy
Publishing House ; Voyenizdat
Place of Publication : Moscow
Date of Publication : 1977
Si~ned to press ' ; 23 Mar 77
Copies ; 10,000
COPYRIGHT : Voyenizdat, 1977
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ANNOTATION
[Text] General information is presented on the space rocket launching
complexes. The classification and designations of the cosmodromes, their
composition and structure are presented. Primary attention is given to
- the engineering complexes and launching pads, buildings and structures,
transport, lifting-po~sitioning and launching equipment, service systems
and thermostating. A study is ma~le of the communications between the
_ ground systems and the on-board systems of the booster rockets. General
characteristics, the organizational and structural principles of the
monitorin~ and control systems for the technological process operations
and the space rocket complex are presented.
The book is designed for engineering and technical workers, the students
' at the higher institutions of learning and people interested in space
rocket engineering.
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FROM THE AUTHORS
The book presented to the reader is a first effort to use foreign and
Soviet sources in a systematic discussion of the basic organizational -
principles of the engineering complexes and launching pads of the cosmo-
dromes and the requirements imposed on them, to familiarize the reader
with the buildi.ngs and structures at the cosmodrome, the structure of the
ground units and systems and to demonstrate the variety and complexity of
equipment required to assemble, prepare for launch and launch from space
rocket systems. General infarmation is also presented about the cosmodromes
of the world and br~ef characterizations of them are given.
Since space rocket engineering is s"till a relatively new field,up to now there
is no standardized terminology either in the Soviet Union or abroad; there-
_ fore the terminology of the M.ALEN'KAYA ENTSIKLOPEDIY KOSMONAVTIKA
[Small Encyclopedia of Cosmonautics] (Moscow, Sovetskaya Entsiklopediya
[Soviet Encyclopedia], 1970) has been adopted.
The bouk was written by a collective of authors as follows: A. P. Vol'skiy -
(the introduction and Chapter 1), A. V. Khaldeyev (Chapters 2 and 6),
N. I. Prigozhin (Chapters 3 and 4), I. A. Shuyskiy (Sections 4.6 and 7.2),
V. N. Nikolayev (Cha.pter 5~and Sections 7.1 and 7.3), V. M. Karin
(Chapters 8-11). -
The authors assume responsibility for the fact that the book is not free
- of deficiencies, and tihey will be grateful to the readers for critical ~
comments and suggestions.
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INTRODUCTION
On 4 October 1957 with insertion af the first Soviet artif icial earth
satellite in history into space orbit, the era of cosmonautics was born.
As a science, cosmonautics was born far before this date, and by rights
its author is considered to be the Russian scientist Konstantin Eduardovich
Tsiolkovskiy.
The history of development of cosmodromes is closely connected with the
development of cosmonautics. As a compositional part of a united space
rocket complex, the cosmodromes must, in accordance with their purpose,
_ meet the requirements imposed on them by the booster rockets and the space
vehicles,l for the performance of ground preparations, launching and flight
control. The structure and compositior~ :,f t;~e cosmodromes and the structural
design of the equipment depend entirely on t?ie structure of the ~pace
rocket systems and the goals which have been set for them.
A characteristic feature of the first foreign cosmodromes was the fact that
the greater part of them were built on the basis of test areas for combat
missiles. The geophysical and meteorological rockets which can be con-
sidered as the f irst generation of space rockets were launched from mobile
ground complexes. In 1946, the United States began a program of launching
the captured German V-2 rockets to ~nvestigate the upger layer of the _
a~tmosphere from the White Sands Proving Grounds (New M~exico, United States),
which included a gun mount type erector, mobile fueling units, a diesel
electric power plant and monitoring and testing equigment. The launches
took place from a pad installed on a concrete foundation.
Ir~ 1949 the two-stage Bumper-VAK rocket (V-2 and VAK-Corporal) launched
fiom White Sands Proving Grounds reached an altitude of 303 lan. The
ground units making up the launch complex for this rocket were also mobile.
1By space velzicles here and here~..;:ter we mean both manned spacecraft
and various satellit,es of the earth and other planets.
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Minimum number of service personnel;
Preparation for launch and launch at any time of year or day under defined
meteorological conditions.
The high reliability of the launches is insured by fail-safe operation of
- the space rocket system and the ground launching complex.
Reliability is the property of the equipment to maintai.n its output
characteristics (parameters) within defined limits under given operating
conditions. From determining the reliability it follows that not only is
the system considered unreliable in iJhich mechanical or electrical damage
is manifested leading to unfitness of it, but also the system for which
the characteristics go beyond the admissible limits.
The reliability of a unit or a system is built in w!~en the system is
designed; the most effective methods of improving reliability are selection
of the elements of increased reli.ability, simplification of the system,
creation of systems with limited consequences of failures of the elements,
redund~ncy (redundancy of the assemb]:ies and systems), built-in monitoring,
automation of checks, and so on. The reliability of the equipment is
increased by improving the productioii technology, automation of the produc-
tion processes, strict monitoring of production quality, the introduction
of special tests with simulation of the operating conditions (usually
extreme values of the loads, pressures, vibrations, temperatures, and so
on are used).
Reliability is closely connected with various aspects of the operating
process: observation of the operating rules excluding the possibility of
breakage of the equipment; periodic checks; performance of preventive
repair work; maintenance of equipment: in technically good working order,
and so on.
- The preservation of the equipment the property of the equipment to remain
in working order in storage is an importatct technical concept. Inasmuch
as storage is an inseparable part of operation and maintenance, the fitness
of the units and systems depends on it to a very high degree.
The cha.racteristics of the possibility of repairing failed systems and
units or individual elemer.ts of them repairability, that is, adaptibil-
ity of the equipment to the detection and elimination of failures and
also prevention of failures has great signif icance. Frequently when
preparing the space rocket systems for launch, it is not the fact of failure
of the unit or system itself that causes alarm, but the impossibility of
quickly finding the failure and quick elimination of it.
As applied to a cosmodrome it is expedient first of all to consider only
the systems and units which have a direct influence on the preparation
of the rocket for launch and the launch itself and secondly, to investi~ate
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their reliability only from the point of view that they are either in
working order or not in working order. Inasmuch as ~n appearance of the
f~ilures in individual elements the systems and units of the cosmodrome
as a whole can continue to perform its function, instead of reliability it
is more appropriate to talk about efficiency.
By the efficiency of a complicated technical complex we mean the degree
of its correspondence to the solution of the stated problems. With this
approach the most important criterion is estimation of the completeness
of the fulfillment of the mission. However, such factors as awkwardness
and complexity of the equipment, Che application of expensive deficit
components and materials, the requirement of high qualif ication of the _
service personnel, high cost of operation and maintenance, and so on, have
a high influence on efficiency. At the present time these factors are more
and more being taken into account in the development of equipment and
organization of operations at the modern cosmodromes.
The insurance of operating saf ety at the cosmodrome is an important require-
ment. It is possible to consider that the cosmodrome is an increased danger
- zone, and in a number of cases, f iguratively speaking, a"powder keg":
explosives and current sources, fuels and spontaneously combustible
components, high pressure lines and toxic working of fluids are side by
_ side k~ere. Therefore inappropriate technical solutions or insignificant
violations of safety measures in operation and maintenance can lead to
emerg~nc~es and even to a disaster.
' The measures to provide for operating safety at the cosmodrome can be
divided into two groups: the f irst group includes the measures provided
for when designing the structures, the systems in the units of ground equip-
ment and the cosmodrome as a whole; the second group includes the organiza-
tional measures providing for observation of safety measures and fulfillment
of the behavioral roles of the service personnel.
The f irst group includes the placement of the buildings and structures of
the cosmodrome at a safe distance fro:n each other, the corresponding organ-
ization of the technological cycle of pre--launch preparations and l~unching
of the space rocket systems, reliable protection o� the structures from
fire and the effects of a blast wave, and the presence of ineans of protect-
ing the service personnel and means of evacuating them in an emergency,
exclusion of improper action taken by operators, and so on.
The buildings and ~tructures of the cosmodrome ar~e grouped in zones depend-
ing on their functional purpose, the degree of danger involved in the
operations and in accoraance with the technological process sequence for
preparation of the space rocket systems. The launch facility is usually
placed at a distance from the other zunes and facilities of the cosmodrome
in order to protect them from damage in case of the explosion of a rocket
during launch or in the initial phase of the tra3ectory. The service
station, the powdered charge storage, the zones for production and storage
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of fuel components, and so on are located at a safe distance. The
launch facility structures are designed for dynamic forces, excess
pressure and sound effects. The dynamic forces occur in the case of
emergency postponement of the launch (shutdown of the engines) and with
respect to ma,gnitude exceed by 1.8 to 2 times the launch mass of the space
rocket system. The excess pressure is created in case of an emergency
exnlosion of the rocket system on the launching paid and it is expressed
in "TNT equivalents" the amount of TNT equivalent to the blast energy.
The sound effects arise from operation of the rocket engines of the booster
rocket during launch, and they are measured by the magnitude of the sound
pressure.
Thus, the launch complex No 39 for the "Saturn-V-Apollo" space rocket system,
in accordance ~~ith the admissible critical values of the excess pressure -
and acoustic effects, is broken down into four functional zones: launches,
launch support, general purpose and industrial.
The launch zone:~s delimited by an excess pressure line of a possible explo-
sion of 0.0028 MPal and a sound level of 135 decibels. The launching pads,
the direct launch ~upport equipment, automatic and remote control optical
and electrical equipment are located in this zone. The distance between
launching paids (2670 meters) was selected so that in case of explosion
the service personnel and space rocke�t system on an ad~acent pad will not
be subjected to above admissible pressures.
The launch support zone is located between the sound eff ect lines of
135 and 120 decibels. The vertical assembly building, the launch control
center, the facility for storin~ chemicals, the storage battery charging
station, and so on are located in this zone. The vertical assembly build-
ing is located beyond the reach of large fragments in case a rocket explodes
during launch.
The general purpose zone beings with the sound effect line of 120 decibels
and reaches the boundaries of the large complex. This zone is relatively
safe and is designed for the general engineering equipment structures.
The industrial zone is located within the limits of the general purpose
zone and includes the installation and test facility for the space vehicles,
the administrative buildings, the pyrotechnical buildings, laboratories,
and so on. '
The structures are protected, as a rule, by being partially underground,
the use of high-strength structural elements, embanl~ents, shielding slabs,
and so on. The structures in which .there are people during the final .
operations and launch are especially reliably shielded: these include
11 Pa=10'S kg-force/cm2; 1 kg-force/cm2=9.80665�104 Pa (exactly)~105 Pa= -
0.1 MPa.
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the facilities for the central preparation and servicing panels and the
launch control center.
The launch structures are also shielded from the gas jet of the rocket
engine on launch. In cases where it is inexpedient to shield the gas
deflector of the launching system, bhe elements for fastening the rocket
during launch and some of the ground cables reliably from the gas jet,
they are made to be used on~y once or for partial:.replacement (repair)
after each launch.
_ The design saf ety measures include so-called classification of the facil-
ities, that is, division of the buildings and struatures into explosion
hazardous, fire hazardous, and so on. For example, the liquid oxygen
storage facility of the launch facility is a f ire-hazardous structure,
and the service tower is explosion-hazardous and fire-hazar.dous, for
the service lines, drainage lines, high pressure lines and electric cables
are run along it. Depending on the category, the equipment of these
structures has also been developed in the corresponding execution.~
For safety, the service personnel are provided with means of collective
and individual protection from the effects of toxic vapor, heat and
shielding in case of fire. These means are varied and include the equi:p- '
ment and attachments from the stationary shielding structures (bunkers,
heat shields, fire-f ighting systems,~ ventilation units, and so on) to the
- simplest fire extinguishers and individual gas masks.
Special attention has been given to the problems of evacuation of service
personnel on occurrence of an emergency, for which provision is made for
emergency exits in the facilities, fire escapes, emergency hatches, and
the structures of the launch facility have tunnels or underground passages,
sometimes running for quiCe long distances. The greatest diff iculty arises
in evacuating people from the service tower, the platforms of which are
located at a great height. The elevators cannot provide fully for the
solution of this problem, for the possibility of their failure as a result
of an emergency is not excluded, and descent by ladders is too slow.
Therefore, special cable devices, rescue cradles and chutes are used in _
emergencies. Designs have been developed for these catapults, individual ,
jet packs and even helicopters and dirigibles.
The measures to insure safety of performinQ operations at the launch
facility include the measures to rescue the spacecraft crews. If an
- emergency develops before the crew boards, high-speed elevators, rescue
devices, evacuation systems and other means of leaving the ser~ice tower
are used (sometimes the same as for the ~erv~ice personnel). ~~BunkPrs and
other shielding structures have been provided to shelter Che crew. In'
case of an emergency with the booster rocket during launch, the emergency
rescue systems of the spacecraft are used which have various structural
executions, but one goal removal of the compartment of the spacecraft
in which the crew is located to a safe distance from the launch site.
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Thus, the compartments with crews o~i the "Soyuz" and "Apollo" space
vehicles are separated by solid-propellant engines with subsequent descent
- of these compartments on parachutes, and the "Vostok" space vehicle had a
catapult seat with the cosmonaut in it.
Incorrect action by the operators during the pre-launch preparation is
excluded by maximum automation of the preparat~on process, modularization
of the units and systems, sound and light signals, warning inscriptions
and all possible forms of monitoring.
The secand group the organizational measures performed at the cosmodrome
primarily.i.ncludes observation of safety measures. For each type of
operation there are specific safety engineering rules (for example,
inadmissibility of an open flame or the occurrence of an electric spark
in the facility where gaseous or liquid oxygen is to be found; forbidding
repairs of tanks and lines under pressure, and so on). In addition,
there are the general standards and rules of behavior of the ser.vice
personnel working at the cosmodrome: the only peopl.e allowed to perform
the various operations are those who have studied the corresponding system
or unit and have the necessary training; people not involved in performing
these operations must be removed from the area where they are performed.
_ Inasmuch as the operations with respect to prepa~ing the space rocket
systems are performed in a strict technological sequence, the violation of
this sequence without the permission of the launch director is categorically
forbidden.
~ The preparation time of the space rocket systems for launch is an important
operating and technical index of the cosmodrome. For modern space rocket
_ complexes the pre-launch preparation cycle involves the time from several
days to 2 or 3 months and depends on the work schedule, the class of
rocket and also the f:1ow chart for preparing the space rocket system for ~
launch. In certain cases the preparation time is not limited, for it has
no significant effect on the fulfillment of the stated mission. In other
cases this time is strictly limited. This arises from the need to launch
at given astronomical times (for example, for a flight to the moon or
- other planets) or after def ined time intervals (when docking space vehicles
in orbit), or when it is necessary to have the space rocket:system ready
to launch in the launching system for emergency aid to a manned spacecraft
which has gotten into trouble
When planning the preparation time it is considered that on the one hand -
shortening the length of the c,ycle can lsad to complicatiun uF the equip- -
ment, the construction of additional structures, the expansion of the
working areas and an increase in the number of service personnel and on -
the other hand, to a decrease in the number of launches from each launch
conplex and reduction of the equipment loading coefficient. Therefore
~ in the general case, when striving to reduce the launch preparation time
of the space rocket system, it is necessary to consider all of these
factors.
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The number of service personnel at the modern cosmodromes, in spite of
the high level Qf automation and mechanization, will reach several thousands
and even tens of thousands of people. This is explained by the great
complexity and variety of the ground systems and the rocket systems requir-
ing special.ists of diff erent prof iles for thelr servicing. In addition,
the buildings, structures and services of the cosmodrome, as a rule, are
split up territorially and are sometimes at significant distances from
each other, which excludes the use of the same specialists. In striving
to reduce the number of service personnel engaged in preparing the space
rocket systems, everything begins with the fact that the machine cannot
completely replace the human operator, who plays the primary role in the
performed operations.
The cosmodrome must provide for launch preparations and launches of
space rocket systems at any time of year and any time of day. This arises
from the necessity for launching at srrictly given astronomical times and
the rocket preparation schedule which must not depend on the capricious-
' ness of the weather. Considering that the cliuaatic conditions at the
locations of the cosmodromes are frequently severe, this requirement cannot
- always easily be met.
1.2. Space Rocket System
General Information
The space rocket system (RKS) includes the booster rocket and the space
vehicle.
The booster rocket is used to obtain the first and second cosmic velocitiesl
and insert the space vehicle into the given orbit. In space engineering
only multistage rockets are used, that is, rockets made up of several
stages in which the spent stage is separated after using up all of its
fuel, and its speed becomes the initial speed for the subsequent stages
and the space vehicle (the payload).
1The first cosmic velocity is the least initial velocity which must be
communicated to a body at the surface of the earth in order for it to
~ become an artificial earth satellite. It is equal to the angular velocity,
and in the absence of an atmosphere it is 7.91 lan/sec. The second cosmic
velocity is the least initial velocity which must be communicated to the
body for it to overcome the earth's gravity on beginning to move near the
earth it varies with altitude and on being reduced to the surface it
is 11.19 lan/ s ec .
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4 .
. 3 '
s � '
' ~cinynen6 ~ ~
.~3~. . a
Y . ~ 3
~ n~ 1
a ~ ~3~
A~ '
~2) ~ ~ D`~'"e~" ' a II~"Y'~'~� ~
? ~2~ ~2~
B 1rmy~xno
1~ Ja~ (1)
- .(1) ~ ~1~ ~ ~
~ . ~ e =
_ . .q -
_ Q ~b _ ~ ~
Figure 1.8. Schematic diagrams of a multistage rocket:
_ a-- with transverse division of the stages (the "tandem" system);
b-- with longitudinal division of the stages (the package sys~tem);
c-- a combination system; 1-- fuel compartments; 2-- rocket
engines; 3-- payload; 4-- nose cone; 5-- control equipment
compartment; 6-- power plants of the stages
Key: 1. lst stage; 2. 2d stage; 3. 3d stage
Structurally the multistage rocket can be executed with transverse division
of the stages (the tandem system), with longitudinal division (the package
division) or a combination of these two systems (Fig 1.8).
In the system with transverse division of the stages their engines operate
successively; in the system with longitudinal division the engines of
the subsequent stage can operate simultaneously with the engines of the
preceding stage; in the combined system, both simultaneously and successively.
However, in any of these systems when the fuel is used up the spent stage
is discarded.
The space vehicle is equipped with a nose cone to protect it from aero-
dynamic loads occurring when the rocket passes through the dense layers
_ of the atmosphere. Structurally the nose cone, the space vehicle, the
engine of the emergency rescue system (if the vehicle is manned) and the
last stage of the booster rocket or its connecting element (adapter)
constitute a single last stage or top module.
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II
1
r-'
~
~ 3
3
2
~~.i
_
:i
'ii
r~
,
~
~
,
~ ~
~
3. I
~
I~ I?
~ ,
I h
~ i
~
Figure 1.9. Booster rocket and the "Vostok" spacecraft:
1-- top module with last (third) stage; 2-- central module
(second stage); 3-- peripheral module (first stage)
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The booster rocket of the "Vostok" spacecraft (Fig 1.9) serves as an
example of the combination system of joining the stages. It is a three-
stage rocket made up of six modules: central, four peripheral and the
third stage module. The first and second stages (the peripheral and
central modules respectively) are made in accordance with the system with
longitudinal division, and the third stage which is installed on the
central module, in aacordance with transverse division.
In space engineering primarily liquid-propellant rocket engines are
applied which use liquid fuel components to create.the jet thrust. The
solid-propellant engines find application only as individual stages or
boosters, and in the space vehicles, for emergency residue systems, soft
landings, and so on.
- The space booster rockets are distinguished by relatively light construc-
tion, the ma.ss of which does not exceed 10 to 12y of the mass of the
entirely filled rocket. When creating the structural design of rockets
having high strength and rigidity, along with using high-strength light
alloys, other solutions are used (maintenance of a defined inside pressure
in the rocket tanks using ground pre-launch blowing systems, supporting
elements for "suspending" the rocket on the launch system, wind fastenings,
and so on).
With respect to launch mass, the space booster rockets are divided into
superlight, light, medium, hea.vy and superheavy. This classificatiun is
somewhat provisional; it has no clear bounds and nevertheless has found
broad application in the technical literature, especially foreign litera-
ture. In the United States the following classification of rockets is
used:
Superlight wi.th a launch mass up to 50 tons ("Scout");
Light with a launch mass to 100 tons ("Thor-Alter," "Thor-Werner");
Medium with a launch mass to 300 tons ("Thor-Delta," "Thorad-Delta,"
"Thor-Agena," "Thorad-Agena," "Atlas-Agena," "Atlas-Centaur," "Titan-1B");
- Heavy with a launch mass to 1000 tons ("Titan-IIIC," "Saturn-IB");
Superheavy with a launch mass of more than 1000 tons ("Saturn-V").
The purpose of the space booster system is determined by the space vehicle.
In automatic space vehicles all the operations are performed without
the participation of man using equipment aid instruments. The manned
vehicles are controlled by cosmonauts on board; some of the manned vehicles
can also operate in the automaCic mode.
With respect to orbit the space vehicles are divided into artificial earth
satellites and interplanetary s~ations. Depending on the purpose of the
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satellites, they are divided into scientific,meteorological, communications,
navigational, and ao on (Fig 1.10, 1.11). The interplanetary atations
(Fig 1.12) are designed for flight to other planets. Some of them can have
the artificial satellites of these planets also in their position.
The structural line of spacecraft has a number of peculiarities connected
with the specific factors of outer space weightlessness, deep vacuum,
the presence of ineteoritic particles, intense radiation for which the
nature of the friction process changes, tihe phenomena of so-called "cold
, welding" occur, meteor erosion takes place, and so on.
Space vehicles which must operate for a long period of time under space
conditions have systems that insure a defined thermal regiune, power supplies
for the instruments and equipment and radio communications with the earth.
On manned spacecraft, the required atmospheric composition is maintained
in the compartments, and conditions required for life support of the crew
are created.
~
; ~ ~ C..
.~.i . C.~�...._ . .
~w `S ~~DSS40~~
w~~ r~~~~~~~~~e~~~~~~~
~~~'~~~~w~~~~a~~~~~~~~~ I
~~~~~~~~~~~~~~~~~ra~~~ ~ ~rii~
iiii~i ~i iiiii~~i~iii~~ ; I M_~~~ ~
~
� �~~.r~~~~~ I ~i
~
~i =i~d~~~~~~~~.~~=iti~.s~i
~ .y~e
~~~~~~rii~~~w~~~~~ ~`=ie
~iiiii=ii~~u~~iii~i~~ I iji
~~i~ii~iu~�i~MS~
2
1 ~
3
Figure 1.10. "Kosmos" meteorological satellite:
1-- actinometric equipment; 2-- infrared equipment; 3--
television equipment
Usually the entire space vehicle does not descend to the earth, but only
part of it the descent veliicle which contains the crew and some of
the on-board systems; the remaining compartments with equipment providing
for orbital flighC of the vehicle are separated from the descent vehicle
at the beginning of the descent trajectory. At the end of descent the speed
of the vehicle is reduced, and a further decrease in speed before landing
is accomplished usually by a parachute system. On some of the spacecraft
("Soyuz," "Apollo"~) a soft landing system is included which makes use of
powder propulsion units permitting the landing speed (dry land or water)
to be reduced in practice to zero. .
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5 6
4
3
2 ~ 7
~ ~i
~ 8
10 9
Figure 1.11. "Syncom-2" communications satellite:
- 1-- telemetric and command rod antenna; 2-- jet nozzle of the
orientation system; 3-- nickel-cadmium storage battery;
4-- radio receiver; 5-- hydrogen peroxide bottle; 6-- coaxial
communications slotted antenna; 7-- radio transmitter with
traveling wave tube; 8-- command radio receiver; 9-- nitrogen
bottle of the position control syst~; 10 solar indicator
w ~ .
fi; - ~~^v . r 7 r~.. p~~
i.:. . .;p~ ~crr ~F a r ~ . �.c.. _p i; -'a
. .~~C { , ~
j , u:` ''.~V:~'~"~,~;,
- ~ 14.. ~
<
~ . ~ ~
~ ~
-Y ' ~M`y~
; j?~, ,
; ~
S r;`*:a~
I~ II II i~ I~ ~ r
~?~i ~ r~
k;~: ~ :t "A
~ 9:, r ;
~ a~
A g .+T'.� d r .t,~s
~ . _ . _ Tr , h .....t.Rr"
Figure 1.12. "Venera-7" ["Venus-7"J interplanetary automatic
station:
1- solar cell panel; 2- astronavigational sensor; 3-- shield-
ing panel; 4- correcting engine; 5-- collectors of the
pneumatic system with control nozzle; 6-- cosmic particle
counter; 7-- permanent solar orientation sensor; 8-- orbital
compartment; 9-- radiator-cooler; 10 low-directional
antenna; 11 high-directional antenna; 12 automation module
for the pneumatic system; 13 compressed gas bottle; 14
descent vehicle
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The variety of structural designs of spacecraft, in addition to their
functional purpose, is also connected with their national origin, with
diff erent approaches to the solution of many engineering problems. This
has been manif ested in preparation for the joint experimental flight in .
docking in space of the Soviet spacecraft "Soyuz" and the American,
"Apollo," in July 1975.
For rendezvous, docking and joint f light, the "Soyuz" and "Apollo" space-
craft (Fig 1.13) were developed considering their compatibility. Instead
of a"rod-cone" docking unit, androgynous docking units were installed
: with peripheral location of the locks.
The problem of compatibility of the atmosphere was also solved. Inasmuch
as an "earth" atmosphere is used on the "Soyuz" spacecraft, and pure oxygen
is used for breathing on the "Apollo" in order to provide for transfer of
the cosmonauts from one ship to the other a special chamber was built for
pressure equalization (this transf er module was built inCo the "Apollo").
' ~ ~ nMOd~nena~7 ~ 2 ) ' :
~ . ~C~~ ~
~ yt
__-~O � _ / - - ~
0
~ .Q ~
V i ?
0 .1
1 '
,A(1Q~1AON" ! i �
Figure 1.13. "Soyuz" and "Apollo" spacecraft
Key:
1. "Apollo"
2. Transf er module
3. "Soyuz"
Orbital stations play a special role in cosmonautics. The f irst ea~peri.mental
space station in the world was created by docking the "Soyuz-4" and ~
"Soyuz-5" spacecraft in orbit. The next important step in their develop-
_ ment was insertion of the long-term "Salyut" orbital:.station into artificial
earth satellite orbit (Fig 1.14).
The further development of space flights is continuously connected with
- the creation of large orbital complexes in terrestrial space. The basis
for such complexes will be multipurpose orbital stations made up of various
purpose modules which will be inserted into orbit by multiple-use rockets
and spacecraft and tiiey will be replaced by new ones as they complete their
missions. Crew to service the space stations will be delivered and changed
analogously.
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Recently specialists of many countries have worked on building multiple
use space transport systems. The solution to the problem of saving the
space rocket systems is being sou~ht in different directions. The possi-
bility of building a multiple use spacecraft with a nonreturnable booster
has been established. The efforts to decrease the amount of nonreturnable
equipment have led, for example, to the investigation of a transport
space system with the multiple use last stage which simultaneously serves
as part of the booster rocket and the space vehicle. The possibility of
saving and multiple use of the most expensive equipment of the booster
rockets has been discussed: the instrumentation of the control and tele-
communications systems, the liquid-propellant rocket engines, the mounted
solid-fuel modules, and so on.
Figure 1.14. Long-term "Salyut" orbital station _
American specialists have developed a design for the "Rombus" space trans-
port system which is recovered by parachutes (the recovered vehicle weighs
252 tons); in this case the landing site of the vehicle is planned to be -
near the launching pads and waterways. After landing, the vehicle will
be delivered on a self-propelled caterpillar unit to a barge and transported
to the installation and testing facility of the cosmodrome.
- The multiple-use space transport systems can be considered as representa-
tives of space rocket engineering of the next generation.
Interrelation of Space Rocket Systems with Ground Complexes
The ground equipment complexes provide for the preparation of space rocket
systems in all stages, beginning with transportation from the manufactur-
ing plant to launching the booster rocket. -
During the initial period of development of space rocket engineering, the
~ goal was not set of insuring (perhaps, even at the expense of some compli-
cation of the space rocket systems) simplicity of operation, convenience
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of servicing, efficient construction of the units and systems ~f the
ground equipment, or reduction of pre-launch preparation time. This led
to ~erious complications with respect to automating the preparation pro-
cess and, consequently, required the presence of a large number of units
and service personnel.
The.experience in developing space rocket engineering has led to the fact
that the space rocket system has begun to be developed as a unit whole, _
which has made it possible to find more efficient solutions to the stated
problems.
The requirements and the possibilities of the ground complex are constantly
considered from the first structural elements of the space rocket system.
Thus, the dimensions of the booster rocket are selected beginning with
the optimal ratio between its length and diameter. However, if only this
condition is adhered to, the ro~ket can turn out to have dimensions such
that it will be i.mpossible to deliver it to the launch site by the exist-
ing means of transportation, and the creation of special transport means
will lead to increased cost of the entir2 complex.
If we begin with an effort to decrease the mass of.the rocket structures,
it is expedient to make the on-board filling lines and cable networks as
short as possible. However, this is ~ot always advantageous for the space
rocket complex as a whole, for in this case it is necessary to have access
to the f illing heads and plugs located at significant height during the
pre-launch preparations which complicates operation and inaintenance,
records a large number of service personnel and complicates automation
of the operations. Consequently, it is sometimes;more expedient to allow
some increase in weight of the structure of the space rocket system and
as a result, to insure convei:ient arrangement in operational respects of
the booster rocket elements cc~upled to the ground equipment.
An analogous situation arises alsc,when selecting the fuel components when
it is necessary to consider not oni'y their energy but also their operating
characteristics. The choice of fue,l components, method of f illing and -
batching has great influence on the structural design of thE space rocket
systems and its pneumohydraulic system. Thus, when using loc~-.temperature
cryogenic components the rocket tanks ust�~ily are lined with thermal
insulation; although this increases the mass, it makes it possible to use
the component in the supercooled form which significantly decreases its
evaporation and also prevents air condensation on the tank walls. The
f illing and servicing conditions have a signif icant influence on the
- strength characteristics of the tanks, the structure and dimensions of the
drai:~age anrl safety valves.
In order to increase the reliability of the launch process it is desirable
at launch time to have a min3mum number of couplings of the space rocket
system to the ground systems. Therefore the majority of ground-on-board
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couplings are broken in advance, just as before launch, and only those
which are required up to liftoff time of the rocket from the launching
system are broken directly during launch.
The interrelation of the space rocket systems and ground equipment is
complicated and va.ried, and the mutual effect is large. The proper con-
siderations of all factors determines how effectively the problem of
building the space rocket complex with optimal parameters will be solved.
Systems For Preparing the Space Rocket Sysr_ems for Launch
The preparation of the space rocket ~j~stems for launch includes the follow-
ing basic steps:
Transport of the elements of the space rocket system to the cosmodrome;
Assembly and testing of the booster rocket and the space vehicle at the
engineering complex;
Transport of the booster rocket to the launch complex and installation of
the launch system;
Pre-launch preparation of the space rocket system and launch.
- The method of assembling the space rocket system, as a function of which
it is possible to isolate three process flow charts, has the most signifi-
cant intluence on the entire preparation cycle:
The first flow chart includes the horizontal assembly of the space rocket
system and complex testing on the installation and testing setup at the
engineering complex; the transportation of the space rocket system in the
horizontal position to the launch complex and erection of it to the verti-
cal position on the launch system;
The second flow chart includes horizontal or vertical as~embly of the
individual stages of the booster rocket in the installation and test equip-
ment, transporting it to the launch complex, assembly of the space rocket
system in the vertical position on the launch system and subsequent
performance of comp~.ex tests;
The third flow chart includes vertical assembly of the space rocket system
and performance of complex tests on the installation and test units (the
vertical assembly building) at the engineering complex; the transportation
of the space rocket system to the launch complex and installation of it
on the launcli pad (the stationary part of the launch system).
_ Each of the systems has its advantages and disadvantages, and the applica-
tion of one system or another is determined by many factors.
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In the first system the individua.l stages of the booster rocket are
delivered to the installation and test facility (i~IIK), where autonomous
checkout, assembly in the horizontal position using a coupling unit or
installation and coupling dollies, complex testing and coupling of the
space vehicle (the las*_ stage) are carried out. The completely assembled
space rocket system is transported to the launch complex on the erector. At
the launch complex it is put in the vertical position and installed in the
launch system.
This arrangement is applicable for space rocket systems, the structural
design of which permits transportation of them in the horizontal position
- (which is determined by the strength capabilities of the rocket and fre-
quently is connected with some increase in its weight). In accordance
with this system the assembly and testing of the space rocket system is
accomplished in the facility under favorable conditions.. which wlll permit
convenience of performance of the operations and the quality of them.
At the same time there is no necessity for building a high-rise installa-
tion and test facility, the creation of a carrier for vertical transf er
of the space rocket system and special tracks which is connected with great
technical difficulties (in particular, with subjection to significant
wind loads). The deficiencies of the system include assembly of the space
rocket system in the nonoperating (horizontal) position; the necessity for
repeated complex testing in the launch position, for the transfer of the
booster rocket from~the horizontal position to the vertical position and
installation of it on the launch system can be the cause of the occurrence
- of failures; the coupling of the service, pneumatic and electric lines to
the rocket at the launch position, which is connected with deficiencies
and operating difficulties, in particular, under unfavorable climatic
conditions.
The first preparation scheme is used for the heavy class "Soyuz" Soviet
- rockets and the American "Scout" rockets.
The second system is used (in Americar_ termonology ~alled the "joint
preparation method") when the individual stages of the booster rocket and
the space vehicle are delivered in a defined sequeiice from the installa-
tion and test facility to the launch site where it is assembled on the
launch system using the service tower, lifts or cranes. During assembly,
the individual systems are tested and checked out, and on completion,
the space rocket system as a whole is subjected to complex testing.
According to this system, only individual stages of the booster rocket are
assembled in the estimation and testing facility, which essentially reduces
the size and cost of construction of the installation and test facility
and excludes the necessity for special transport means to be used for the
completely assembled space rocket system. The deficiencies of this system
are unimproved test process in connection with the performance of opera-
tions in the open air, which lowers the reliability of the preparation
of the space rocket system and the fact that the assembly of the space .
rocket system on the launch system occupies the launch complex for a
- prolonged period of ti.me, reducing its carrying capacity.
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This preparation system is used primarily for rockets with long intervals
between launches (for example, the "Europa-II" booster rocket), the
medium class rockets (American booster rockets "Chori Delta," "Atlas-Agena,"
"Atlas-Centaur" and so on), and it is admissible �or cosmodromes located
_ in areas with mild climate.
The second system for preparing the sp2ce rocket systems for launch became
widespread in the United States during the birth of space engineering,
and it was inherited from the process used in preparing combat rockets.
American specialists consider it expedient to build complexes for space
rocket systems by adapting the related launch complexes for strategic
rockets with established rules for preparing them and not by adopting new
structural designs taking into account the sp~cific nature of space engineer-
ing. This approach which was advantageous from the point of view of rapid
~ introduction of the space rocket complex into operation, did not ~ustify
itself when the necessity arose for launching various versions of rockets.
By the third system (according to American terminology, "the mobile prepara-
- tion method") the space rocket system is assembled in the vertical position -
on the launch platform (the upper part of the launching system) which is ~
transported tagether with the space rocket system to the launching site;
the launch takes place from it subsequently (after installation of the
~aunching pad). This system permits all of the numerous filling, pneumatic `
and electrical lines located at various leuels to be coupled to the rocket
at the installation and test facility (the vertical assembly building).
In addition, the rocket coupling lines can be led out through the service
- cable tower installed usually on the launch platform to a convenient service
zone which facilitates coupling of them to the ground systems at the launch-
ing site. The deficiencies of this system are the construction of an
eapensive vertical assembly building, the creation of the carrier with
complex configuration or transportiiig the space rocket system in the verti-
cal position from the engineering complex to the launching site~and laying
a special track which, as already been stated, is a technically'difficult
problem. The third preparation system is used for the heavy and superheavy
class American booster rockets.
In the American literature on space rocket engineering it is possible to
encounter a description of the "f ixed preparation method" which is a version
' of the second system and is used for the medium-class booster rockets. Its
essence consists in the fact that the individual stages of the booster
rocket, bypassing the engineering complex, are delivered to the launch site
where the vertical assembly of the booster rocket takes place, it is
coupled to the service tower coupling lines, undergoes complex checking
and launch.
For super-heavy space rocket systems it is probably necessary to have other
assembly and transport systems, for the TNT equivalent and sound effect
increase significantly and, consequently, it becomes necessary to place
the launch system at a greater distance from the engineering complex. As
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a result of signif icant size and weight of such booster rockets, difficul-
- ties occur in the application of traditional means for delivering the
~ space rocket systems to the launching site.
Thus, there is an American plan for the use of a marine floating assembly
and launching system where the basic element is large, the greater part
of which is occupied by the compartment for assembling the space rocket
system. During assembly the barge is kept on a special dock, and its decks
are kept open. The launch will take place at sea. The bow of the barge
- where all of the control and launch equipment are located remains in the
horizontal position before thr; launch, the stern is disconnected and put
in the vertical position by filling the aft tanks with w~ter, and after
launch returns to the initial position.
1.3. Main Cosmodromes of the World
The Baykonur Cosmodrome one of the largest cosmodromes in the world
(Fig 1.15) is located in Ka.zakh SSR, in a semiarid zone with sharply
continental climate (hot, dry summer and cold winter with high winds and
insignif icant precipitation); it was founded in 1955.
The basis for selecting the construction site for the cosmodrome was its
sufficient remoteness from large populated areas, the possibility of insur-
ing safety of the rocket launches, the crea.tion of alienation zones, zones
for landing the returnable space vehicles and also the presence of a large
number of cloudless days during the year.
The cosmodrome routes extend thousands of kilometers over the territory
of the Soviet Union ~3nd end in the Pacif ic Ocean where the last stages of
the boosrer rockets a.re dropped. Along the routes there are measuring
stations and especially equipped ships. The space vehicles are inserted
into orbits with an inclination to the plane of the equator from 48� to 81�
with easterly direction of the launch. The space vehicles and manned space-
craft usually land in the northeastern regions of the Kazakh SSR.
Launches have been made from the Baykonur Cosmodrome in accordance with the
Nationa.l Program of the USSR for the Study and Use of Outer Space, within
the framework of cooperation with socialist countries by the "Interkosmos"
program and also in accordance with agreements for joint efforts to explore :
outer space concluded between the USSR, the United States, France and other
countries.
TY?e first artif icial earth satellite in the world was launched from the
Baykonur Cosmodrome. The flights by cosmonaut Yu. A. Gagarin ancl the first
female cosmonaut V. V. Tereshkova into outer space were made from this
same location. It also launched the automatic interplanetary stations
"Luna," "Venera," "Maris," "Zond," the space stations and artificial earth
satellites of various types ("Kosmos," "Elektron," "Polet"),.the
satellites of the "Molniya" s~ries for relaying television programs and
long distance telephone and telegraph communications.
36 -
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Figure 1.15. Baykonur Cosmodrome (launch site of the "Soyuz"
space rocket system
Launches of manned spacecraft "Soyuz" and the orbital stations "Salyut"
are made regularly from the Baykonur Cosmodrome.
Space vehicles with French equipment have been launched within the frame-
~ work of a cooperative program with France.
- 37
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In accordance with the Soviet-American EPAS program, the "Soyuz" spacecraft
was launched from Baykonur Cosmodrome on 15 July 1975, for docking in orbit
;~ith the "Apollo" spacecraft.
The "Soyuz" spacecraft equipped with a camera made 3n the German Democratic
, Republic was inserted into orbit in September 1976.
The cosmodrome has built a number of launching sites and engineering
complexes. One of the most important is the complex from which the three-
stage booster ro~kets with the "Vostok" and "Voskhod" manned spacecraft
were launched, and at the present time the "Soyuz" spac.e vehicles are being
launched.
~
The launch structure for this booster ~ocket is the semiburied type. It
has a launch system with ejectable supporting beams. The rocket is
"suspended" in the launch sysCem behind the power packs. The space rocket
system is delivered to the launching site from the installation and tr~sting
unit of the engineering complex where it is assembled in the horizontsi
position.
In addition to the ins*_allation ann testing unit, the MIK KO building, the ~
service station for the space vehicles, the storage battery charging stat.ion
and a number af other buildi.ngs and structures are located in the engine;:r-
ing complex. The measur.ing stations are located here which are equipped
with telemetric equipment, television set, antennas, radio receiving and
transmitting units.
The living quarters of the cosmodrome in which there is a complex for train-
ing the cosmonauts (cla~sreoms for exercises of the crew in accordance with
:.he tec:hnical and scientific training program, a sports complex with a
swimming pool, laboratory for preparing the ca~monauts for flight, a medical
- compl.ex) and also an institute, technical high school, schools, club, ~
stadium, television broadcast center, and so on are located.
The cosmodrame is connected with other plants in the country by air, high-
way and railroad transportation. The cosmodrome territory also has a
branch network of highways and railways.
The eastern test area (before 1965, the Atlantic Missile Range) is the
largest American cosmodrome. It is located at Cape Canaveral and Merrit
Island (inthe state of Florida), and it has a territory of about 400 km2
(Fig 1.16). The basis for the selection of this site was its adequate
isolation, which guaranteed launch safety and offered the possibility of
further expansi:on of the territory. In addition, the convenient location
of the islands of the West Indies and the South Atlantic made it possible
to install monitoring and measuring complexes on them to observe the flight
of the rockets.
The range track about 20,000 lan long runs above the Atlantic and Indian
Oceans to the Prince Edward Islands, and has 15 measuring stations
38
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equipped with optical, telemetric and radar units. The rockets are
tracked also from ships and aicraft and from more than 100 individual
ground observation stations.
The syr~tem of. tracking stations existing in the eastern test area permit
l~unche~; wilh an aztmuth from 44 to 110� and insertlon of the artificial
earth satellites in orbits with an inclination to the plane of the equator
from 28�30' to 52�24' with easterly direction of the launch. Launches of
artificial earth satellites both into equatorial and polar orbits are
possible from the test area, but insertion into polar orbits is connected
with the performance of the heading maneuver in the active section of the
booster rocket flight. Much greater expenditures~.of the energy reserves
of the space rocket system are required to achieve polar orbits than
equat~rial orbits.
The test area is located in a highly swampy, flat area with rock occurring
at a depth of about 50 meters; the air temperature fluctuates from 0 to +50�C
during the year; powerful hurricanes and typhoons are possible with
a wind speed of up to 55 m/sec.
This test area has all forms of communications (air, sea, railroad, motor
vehicle). The booster rockets are transported predominantly by air and
water; the light class booster rockets and their elements are transported
on aircraft, and the heavy class booster rocket stages are transported on
barges and ships.
Along the coast line of Cape Canaveral ard:the soutihern part of Merrit
Island there are 20 launch complexes, of which 12 belong to the Eastern
Test Area and 8 belong to the Kennedy Space Center. The launch complexes
of the Eastern Test Area are designed for launchiilg various space vehicles
using the "Atlas," "Titan" and other booster rockets and the Kennedy Space
Center, using the "Atlas-Agena," "Saturn-IB," and "Saturn-V" booster rockets.
In addition to forming space research, the Eastern..Test Area is used wide3.y
to test American combat missiles: more than 200 flight tests and several
thousand bench tests are run on the rockets yearly at the test area. The
service personnel, including the tracking stations, number more than 20,000.
The Kennedy Space Center (Fig 1.17) is the main NASA test area and is
designed for launching space vehicles and testing booster rockets in
accordance with the American National Space Research Program.
The mission of the center includes the following:
Planning launches of NASA space vehicles;
Assembly, testing, check out and launching of space vehicles;
Coordination of operations performed by the joint programs with the Eastern
Test Area;
39
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40
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Performance of scientific research and experimental design work in the
f ield of developing methods of launching space vehicles, the design of
new launching complexes and also modificat3on of the existing launching
equipment;
The study of the possibility of launching rockets from orbit and servicing
them in orbit;
The planning and design of storage and operational equipment for new fuel
components;
The training of scientific and technical person~el, and so on.
The ballistic characteristics of the center (the'directions of the booster
rocket flight paths, the range of launch azimuths;, the inclination of the
orbits) are analogous to the characteristics of th~~ Eastern Test Area,;
and the rocket flights are observed by common monitoring and measuring
complexes.
For communications with Che plants of the space rocket industry, the same
means are used as at the Eastern Test Area.; The staff of the center
numbers about 2800 people.
Although the Eastern Test~Area and the Kennedy Space Center are territorially
joined and interact with respect to certain problems, th~y are two
administratively independent organizations which have different equipment
and solve independent problems in the interests of the U.S. Air Force and
NASA, respectively.
The Western Test Area (until 1965, the Pacific Ocean Missile Range) is
located on the Pacific.:oast of the United States, north of Los Angeles
(Fig 1.18), and it includes the Vandenberg Air Force Base, the Point Mugu
Marine Test Area, the Point Arguelo Test Area and an inland test area,
of which only the Vandenberg Air Force Base and the Point Arguelo Test
_ Area are used to launch space rocket systems. The Vandenberg Base (the
missile testing range) works on the development and testing of ground
equipment for the air force missiles, training of launch crews to service
_ the booster rockets, the creation and testing of antimissiles and launch-
ing of military satellites into polar orbits ("Discoverer," "Midas,"
"Samos" and so on). The Vandenberg Base has three launch complexes for
the "Atlas" rockets, two for "Titan" rockets, one for "Scout" rockets and
14 pads for launching "Minuteman" missiles. The Point Arguelo Test Area
is used for launching artificial earth satellites into polar orbits.
About 140 launches are made annually from the Western Test Area.
The track of the test area which is more than 16000 km runs over the
Pacific Ocean and is divided into three test areas: the Hawaiian Islands,
Kwagelein Atoll and Eniwetok Atoll. The uionitoring and measuring means
are located in these areas, including 10 measuring stations equipped
with optical, telemetric and radar equipment. Ships and aircraf* are also
used for rocket flight tracking.
41
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42
FOR OFFICIAL USE ONLY
:
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The admissible launch sector is bounded by azimuths of 301� (the upper
bound) and 170� (lower bound). The central track runs along the 261�
azi.muth. The orbits of the launched satellites have inclinations from
34�2' to 90� (when moving west) and from 81�48' to 90� (when moving east).
The test area occupies a territory of about 400 km2 (the continental
section). It is located near large enterprises which build rockets and is
connected with industrial areas by waterways, railroads and air service.
The total number of personnel working at the test area exceeds 17,000.
The Western Test Area has a number of advantages over the Eastern Test Area.
For example, artificial satellites passing over the poles of the earth can
be launched from it, which makes it possible to study almost the complete
surface of the earth, including the northern regions. The trajectories of
the booster rockets, the active section, run over the ocean; no pieces of
dry land are encountered until Antarctica itself. This makes it possible
to insert artificial earth satellites into polar orbit without risk that
the spent stages or failing rocket will fall on populated areas and also
to use coastal waters for separation of the launch boosters.
The test area on Wallops Island (United States) which is part of the Wallops
Test Station is one o~ the principal NASA bases for launching research
rockets and artificial earth satellites (Fig 1.19. The test area was built
in 1945 by the Lang].y Scientific Research Center (the National Consultation
Committee on Aviation the predecessor of NASA) to test unmanned vehicles
and study the aerodynamic problems of flight. With the formation of NASA
the test.area was reorganized into an independent center.
The Wallops Test Station is located on the eastern coast of the United
States 260 km southeast of Washington and is made up of three zones: the
basic zone which was the former air force base, the zones on Wallops
Island (8 lan long and 0.8 lan wide) and the continental zone which is
located 3.2 kn west of Wallops Island.
In the basic zone are the administrative and functional branches; the
experimental design office, laboratories, the launch control center, the
communications center and telemetric data reception center, one of the
stations for transmitting commands and receiving data from the "Tiros"
- meteorological satellites and airports. The zone on Wallops Island includes
six launch complexes equipped with equipment for assembly, preparation
and launching of the rockets and also for observation of the flight. The
tracking stations, the radar complex and the experimental flight base of
the 7.incoln Laboratory are located in the continental zone.
The vehicles can be inserted into orbits with inclination from 37� to 54�
basically by the "Scout" booster rocket.
The track of the test area runs over the Bermuda Islands where tracking
stations are located which are equipped with measurement means and receiving
radio telemetric stations. The launch sector is bounded by the azimuths
of 67� (upper bound) and 143� (lower bound).
43
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~
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s . ~a
�
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- /0 -
5.
0
~
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Figure 1.18. Western,Test Area of the United States (location
of the launch complexes at Vandenberg Base):
1-- flight control center for the "Atlas" rocket; 2-~ installa- -
tions for launching the "Atlas" rocket; 3-- flight control center
for the "Thor" rocket;.4 devices for launch~.ng the "Thor"
rocket; 5-- devices for launching the "Titan" rocket; 6--
telemetric station; 7-- tracking station; 8-- control center;
9-- station for sending signals regarding emergency elimination
of rocket; 10 liquid oxygen plant which produces 50 tons a day;
' 11 liquid oxygen plant which produces 25 tons a day; 12 test
area for teaching the techniques for recovering the nose cones.
The test area on Wallops Island is used for flight testing of i.ndividual
structural elements and equi:pment of the vehicles developed by NASA and
also for launching research rockets and launching certain artificial
- satellites, including those built by other countries.
~ 44
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There are more than 400 launches a year at this test area. During its
existence it has launched more than 6000 research and experimental;
rockets of various classes. Its service personnel number 500-600.
The Kuru Test Area (Fig 1.20) which is under the ~oint control of France
and tlie ~uropean Rocket Development Organization(ELDO) is located on the _
Atlantic coasC in French Guiana 32 km from Cayenne (almost on the equator
5� north latitude). The test area has three launch comp~exes: for
high altitude rockets, the French "Diamant-B" booster rocl.zet, and for the
"Europa-II" booster rocket (the ELDO organization).
The launching complex for the "Europa-II" booster rocket is oriented along
the "north-south" line, but launches are possible to orbits with a declina-
tion from 0 to 100�. The "Europa-II" booster rocket was built on the basis
of the "Europa-I" rocket launched from the English-Australian Test Area in -
Woomera. The move from the Woomera test area to the Kuru Test Area was
made because with latter is located closer to the equator and is more
favorable for inserting a payload into stationary orbit (the "Europa-II"
rocket is designed for launching commuiiications satellites into stationary
orbits). ~
Although the Kuru Test Area is located in a wet tropical climatic zone with
prolonged rainy periods, the rocket equipment has not been modified, for
the launches are undertaken only during the dry seasons. In addition, the
greater part of the time the stages of the rocket and the payload are in
- air conditioned facilities.
The first stage af the "Europa-II" booster rocket is delivered by wat&:r to
the test area port; the upper stages and the payload are delivered by air
to the airport at Cayenne, and then by motor transportation to the test
area. The port can take ships with deep draught only during high tide
which comes at 14-day intervals, which limits the capabilities of the test
area.
The tracking units (radar, telemetric data reception stations, movie
theodolites and other equipment) are located both in the test area itself
and at other locations.
- According to schedule two launches of the "Europa-II" b~oster rockets
must.take place each year.
The permanent personnel of the test area. number 600 to 700.
The English-Australian rocket test area in Woomera (Fig 1.21) is located ~
in the vicinity of Woomera (southern Australia). The dry land part of
the test area track runs 200 km over the lightly populated parts of
Australia and can be extended 4400 km into the Indian Ocean. Experimental
launches of the English "Blue Streak" booster rockets and the "Europa"
rockets and also launches of research rockets to the upper layers of the
atmosphere take place from the test area.
45
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v~ _
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~i' b
Figure 4.12. Gun mount type erector:
a-- general view; b--- diagram; 1-- lower logement; 2--
folding arm of the boom; 3-- arm lock; 4-- boom; 5-- calibrated
- support; 6- upper grapple; 7-- truck; 8-- cable coil;
9-- winding mechanism; 10 "wing" of the frame; 11 tension
bolt; 12 suspension shackle; 13 guy; 14 pump; 15 fasten-
ing rod; 16 frame; 17 hydraulic support; 18 control cab;
19 hydraulic cylinder for rais~ng the boom; 20 shackle; 21
boom tie rod; 22 suspension rod
Key: 1. A view
1],0
~ FOR OFFICIAL USE ONLY
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section of the boom which is fastened in hinges and the erector is taken
out of the launch system.
The gun mount type erectors have their origin in the birth of rocket
engineering; their structural design has been developed, and they are
reliable in operation. The relat~vely small dimensions (determined by the
diameter and length of the rocket ~ystem) and light weight (exceeding the
~ weight of the rocket system by a total of 2 or 3 times) of these erectors
has insured them quite broad application in rocket engineering.
The stationary ar~d semistationary erectors 3re placed near the launcli sys-
- tem.
The stationary erectors with lifting truss (Fig 4.13) are designed for
assembling the rac:kPt system on the truss in the horizontal position and
subsequent installation of the assembled rocket system together with the
truss in an inc].ined position. The "Scout" rocket system is installed
on the launch complexes flf the Western Test Area and the test area at
Wallops Island by this system.
The stationary erectors with lifting service tower are used to erect the
transport dolly together with the rocket stage to the vertical position.
The first stage of the "Titan-II" boostPr rocket is lifted to the vertical
position by this scheme on the launch complex No 19 of the Eastern Test
- Area.
The semistationary erectors with lifting frame of the transport unit
(Fig 4.14) are used to erect the rocket to the vertical position with the
help of the hydraulic lifter of the frame of the transport-erection dolly.
The Frame of the dolly is hinged with the launching pad.
The semistationary erectors with lifting platform provide for erection of
the space rocket system to the vertical position using a boom and the
railroad transport-erector dolly. The boom of the erector has the form
of a platform with rails, and it is located at the launching site so that
the railroad transport-erection do11y can be rolled on the platform. The
dolly is made up of the running gear, frame, supports and fasteners, a
remote mechanism for opening the fastening clamps to release the rocket
system after it is installed in the launch system.
The stationary erectors have the same advantages as the erectors with the
gun mount type lif ting boom. In addition, it is possi.ble to cansider
among their advantages the possibility of automating the process of
installing the rocket system on ~he launch system and short installation
time.
Some of the service towers (Fig 4.15) placed at the launch complex near
the launch system are equipped with cranes with a set of attachments for
assembling the rocket system by parts. The application of the cranes
111
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permits the rocke:t system or its stage to be moved from the hbutJittal
position to the v~ertical position by simple technical means,
significantly increases the duration and labor consumption of the operations
required to assemble the rocket system.
This system is used on tlinlatheh"S turn-IB"Nrocketnwith tierspacenvehicles.
Space Center when assemb g
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f \ '`q
a 9'�
rYY
'1~, * ~
..~t
/ .~r
~a'
t 3 ,
~ 7 .:s=' A- :.,.t
Figure 4.13. Stationar,~ erector with clamping of the rocket
at the top (for the "Scout" booster rocket)
4.4. Launch Systems
The launch system which provides for the acceptance, erection, verticaliza-
tion and launching of the rocket is also used to bring various lines to
the rocket system, service it, rotate it and provide azimuthal guidance,
and it is the base on which the service cable tower, the cable masts, the
supports of t~secaneC~�~v~de forhtransportation of thetspacesrocket systems
launch syste p
and erection of it to the launch position.
~112
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Therefore, ttie laun~h system not only .'epends on the structural design
of the other units coupled to it, but directly influences their structure.
The structural design of the launch system is determined by its ~rimary
function support o..f th~ launch and it is developed considering the
class of rocket, its power system and the gas dynamic characteristics of -
the engine.
Beginning with the power system of the space rocket the launch systems can
be built to support the space rocket on the end and with suspension on the
supporting elements; the most widespr.ead are the ?aunching systems of the
first type. In order to hold the rocket against wind loads, fastening
assemblies are provided on the launchii:g pad (levers, clamps and locks).
When the rocket is launched, the gas jet is deflected by the gas deflectors.
The distance rrom the engine nozzles to the gas deflector and the angle of
encounter of the jet with the deflector walls determine the structural
design of the deflector, the dimensions of the launch system with respect
to height and the depth of the gas removal channels for the semiburied type
of launch structure. The spacing and the angle are selected beginning
with the admissible temperatures and escape velocity of the gas jet which
can cause erosion of the deflector and also considering a decrease in the
possibility of the formation of reflected waves (the so-called bottom effect
phenomenon) which can destroy the tail section of the rocket.
2 \
4 _
0
0
0 o u
.y. .y., ~o..o~~ .
�1. ?�'1~ '~yp~;i,:pO ~
�,h �o~
~ 3 �
Figure 4.14. Semistationary erector with lifting frame of the
transport unit:
1-- tractor; 2-- frame of the transport dolly; 3-- hydraulic
lift; 4 launching pad
With respect to structural design the gas deflectors are pure metal, wedge
shaped and trough. The pyr~m3.da1 de~lectors usually have a number of
faces equal to or a multiple of the number of combustion chambers of the
_ rocket engines. In thi.s case the gas jet either freely flows over the
launch site or is removed along several gas removal channels. In the
case of the wedge gas deflector the jet is split into two parts and is
removed to the side; in the case o~ the trough deflector, it is removed
in one direction.
113
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_ ~
Figure 4.~5. Assembly of the booster.rocket stages using
a service tower crane
The launching pads (Fig 4.16) for launching the light class rockets are
made in the form of a frame installed on several supports (from 3 to 6)
in which the lifts for moving the frame during acceptance and verticaliza-
tion of the rocket are mounted. The lift mechanisms have hydraulic
(hydraulic jacks) or mechanical ~screw type 3acks) drive. The gas
deflector is located between the supports of the bench; sometimes the -
supports are protected by fairings.
On the upper section of the frame for erecting the rocket~there are support-
ing elements, the number and structural design of which depend on the
supporting elements ~~f the rocket, wind and storm fastenings and also the
114
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attachments for fastening the electric plugs, the pneumatic block~, ttie
fill and drain connections. The rotating section is made in the Form of a
frame on a ball race and a rotary mechanism (toothing mounted on the
rotating fr2me and reduction gear with drive). -
The launching pad is fastened to the foundation of the launch site us~.ng
anchor bolts or other elements. It can be dismantled and transferred to
another launch site.
The stationary launching pad (Fig 4.17) is a quadratic reinforced concrete
structure on supports with a hole in the upper plate and a roedge shapPd gas
deflector. The rocket is ~~nstalled on the pad using its support:;~ng elements.
The launch platform (the upper part of the launch system) fox the "Saturn-V-
Apollo" rocket system has a two-level structural design with a p1at.Eozm
7.6 meters high, 48.8 meters long and 41.1 meters wide with a hole in the
center (13.7x13.7 meters) for the gases to pass through. A ser.vice cable .
tower is mounted on the platform (Fig 4.18) along with four grapples which
hold the rocket system and three service cable masts. The platfor.m is
equipped with fastening mechanisms to the caterpillar carrier and to six
supports and four telescopic columns of the launch stand.
In the compartments of the platform and on the upper plate electrical and
mechanical plugs are installed which provide for connecting the booster
rocket systems to the corresponding equipment in the vertical assembly
building and on the launch stand and also the launch and testing electrical
equipment, the equipment for testing the hydraulic systems, the fueling
- and pneuma.tic lines, the ventilators, air conditioners, and so on. The
floor of the compartments is equipped with shock absorbers, and part of _
the equipment is mounted on springs. The compartments with electronic
equipment have sound insulation, which reduces the noise level when the
rocket engines are operating.
The launch system with removable trusses (Fig 4.19) is designed for the
booster rockets that do not have supporting elements on the end and are
auspended from the supporting assemblies on the central module at the
point of fastening the side modules (for example, the "Soyuz" booster
rocket). The launch system is iu the form of four supporting trusses on
which the booster rocket is hung and which are withdrawn under the effect
of counterweights after thrust is developed; in the lower section the
system has guides for the movement of the space rocket system in the
initial part of liftoff. .
The trusses and guides are fastened to the platform providing for vertical-
ization of the rocket system using the hydraulic syste~ iocated in the
base of the supporting trusses. In the upper part of the platform there
are service trusses, a service cable mast and cable mast for bringing
the fuel:ing and electrical lines to the rocket system.
115
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- , c~
b
~ ~
u
�n
~ ~
I
~ .
~ / tr'1 N
- ~
/ - - - - ~ O b~0
t~ I
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i
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w
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I
FOR OFFICIAL USE ONLY
~ ~ , ~i~ 2 =
~ 3 ,
a~ y ~
~n I 6
~ ~
M ` � i ~
~
~
''I
I I ~~~'M 8 .
I
4 ~ ? n
~ �
+i ~
~ C _ I
~i ~ ~~I~ , ~ _
9 ~ i
o ~
~ o ~ ~ ~ 19 : ' ~
_
~;;liy~.ii;ti;!~~�iii~a:~.+r:{l'.~.`~'~::u;::L.:':iw'~`:. ~i'j ~ --a .
~
Figure 4.17. Launching pad (stationary):
1,18 service masts; 2,4 cable mast; 3,7 heaters;
5,16 supporting structures of the rocket; 6-- valves;
8,9,10 fill lines; 11 water supply line; 12 hydraulic
system control panel; 13 dollies with equipment for ser-
vicing the engine; 14 service platform; 15 electric
cables; 17 shield with instruments for detecting leaks;
19 control panel of the engine service system; 20 tail
section of the rocket; 21 gas deflector.
The launch system is mounted in the launch structure of semiburied type.
The tail section of the booster rocket is below the "zero" level in this
case.
- A single-slope gas deflector and the trough type gas moving channel are
used to remove the gases.
117
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Figure 4.18. Launch platform (upper part of the launch system)
for the "SaturrrV-Apollo" rocket system with
service cable tower
4.5. Service Means ~
The service means include the tower trucks, trusses, towers and service
cabs, the service cable towers (masts) and the cable masts.
The tower trucks (Fig 4.20 and 4.21) are used to service the light and
medium class rockets; they are usually towers mounted on a truck chassis
and have a drive to life the service platform with power takeoff from the
truck engine or from an outside current source.
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~
r � t'~~
~ ; ~ .
T ~k 2,, y~~~,,
. ` ~ e ~a ~ M~I i
r , . ~ . ~ ; ~~r. e .i _
~
~ha'~.r i~^.'r..? ,A' ';'~{7 i~A ~,5
~-",y~' ~ i ~ A i A6 ~ ~ ,
. r
j p iM ~M~..W~ _
a ~r~ 1j ~ , ~ ,
s
, 2 }-~..y ~.~YM ~ ` n -t~ _'.'�e 7 'vz t ~,7! i
~Ti.
. r #~�~,N . ~ '~~~'~~h~~~~.._
~ $ r~ r i
~ ~.F.~ ~6 ~ ~ - ~7~.:4
: , ~1 y i4. ' '
k 3 �.u'"jfh'
P~.. r ~ s ' 4~ ^";F,~ re y i~''; ` i
y ~6 r };3 ~ $ ~t' T t . . I ~f'
~E} ~~n ~ ~ ~
1V a. ~t r:
~
x.i w
q':`.
~ y.~.,.
k ~ C` ~
~r~
~
t C
�c '
Figure 4.19. Launch system for the "Soyuz" booster rocket
The basic deficiencies of the tower trucks limited service height and -
low load capacity force them to be used primarily for auxiliary purposes.
The service trusses (Fig 4.22) are designed to service the heavy class
_ rocket systems. They encompass the rocket systems on both sides, they have
- telescopic, folding and stationary platforms with enclosures and ladders.
Each service truss is made up of a bearing structure, the supp~rting
- assembly, the hydraulic system and control panel. The hoist runs through
all of the service platforms.along one of the trusses. There are hinged "
lever mechanisms �or folding the service platforms on the bearing structure
of the trusses when they are lowered to the horizontal posita.on. The
trusses can be mounted on the rotary of the launch system and rotated in
the azimuthal plane together with the rocket system.
Before launching the rocket system the trusses are moved from vertical to ~
' horizontal position.
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. _ _ .
-.r . ~ . . . . _ . . _ . . . . _ . . .
' ~ ~ ~ . ~ . ~ . . ~
. ' ~ ~ ~~~I
t
r
. ~ . t ' . , ~ ~ ,
`t 1
~
t i
~ , ~ . . 6 A }j'+' j
. . Y . �~y~.. I
A r ~
~ z r r ~
' ' ~rt'r'
t
, ~ z ~i ~F~ m
~:i ~
~ :r ~ ~ ` 1=k -
1 ~ ~ . . ~ jiJ ~ .
~ ?~~4... A.M
? . s\
~,~p~ . f 1 ~ - .:�Gb~
{"''p k 3 I, k
;
~ i~~ ~r ~.y~ ~~~j, '
. T
K,: . . ~,y
' : . . F~!.
at" i~
.A+~
.:r
.y": 4 ~ (
..2..:'. ff. _ .
~ -
~rS7+' .
io~~~ Y Y
Figure 4.20. Arm type tower truck
The service towers are used for the same purposes as the service trusses,
and they can be both movable and stationary.
- The movable towers can be moved on railroads (up to 30 meters wide) a
distance insuring their safety during launch or during an emergency with
- the rocket system.
The rotating towers (Fig 4.23) are used to service the rocket system in
one plane. Before launch these towers are rotated along a ring rail
, around the central support at an angle insura.ng their safety.
The stationary servi.ce towers are autonomous units with an eZectric power
plant, an air conditioning system, ventilation, lighting network, heating
and communications. Their height reaches 100 meters, and they weigh up
to 3,500 tons. Electrical, pneumatic, fill and drain lines with filling
connections and also the thermostat:~ng system lines are laid on the
service towers.
~ 120
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~
; ti;:
n~ .,y:.;:y.
it'%
~ i:w~r i . ~ . ~
p
L.i
r4i!~"
~~i
i Ff
~i t ' ~ ~ . f * A`,i. ~
. ~f y . .M .
~ z
. ~~x
r
4 s~~
~
. ~~a
t
Figure 4.21. Telescopic type tower truck
The movable tower for the "Saturn-V-Apollo" space rocket system (the so-
called mobile service tower) of the launch complex No 39 (Fig 4.24) is -
designed to service the compartments of the booster rocket and the `
"Apollo" spacecraft and install explosive hazardous equipment (the solid- `
propellant braking rocket engine, the engines of the emergency rescue
system, pyrotechnical devices, and so on). The tower is a welded metal
structure 122 meters high and it has a square base 41 meters on a side. '
~ A rotating 4-ton crane is installed on the upper platform of the tower.
The tower has five trusses used as service platforms and supports for the
- pneuma.tic, hydraulic, electrical and other lines. The two lower platforms
move freely along the vertical; the three upper platforms are rigidly
attached, but the entire truss-platform unit can be installed at different
levels depend:ing on the service zones of the booster rocket and the space
_ vehicle; the upper and the two lowe: platforms are open, the two middle
platforms are covered on all sides, and they have an air conditioning
system. -
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The tower is delivered to the launch stand by a caterpillar carrier; then
it is lowered to the supports of the launch site prepared for it, and before
long it is withdrawn to the parking area (approximately 2 km from the
launch stand).
7 ~
Q y :
3 .
~ . _
, b : � _
~ 3 . . . . , �
� ,f ~ . .
Y � ~r -
/4 ~`7~ ~
~`A
! .
. ~ ' '
'
� . Q!/ ~ .
Figure 4.22. Service truss:
1-- truss support; 2-- hydraulic cylinder for raising Che truss;
3-- middle platform; 4-- intermediate platform; 5-- bearing
structure of the truss; 6-- upper platform; 7, 10 wind
shield; 8-- cable of the cargo hoist; 9-- transfer platform;
11 hinged arm mechanism for folding the lower platform on the
truss; 12 ladders and gangways; 13 lower platform;
14 pull xod o~ the hinged arm mechanism for folding the
~ middle platform on the truss.
The service cabs (k'ig 4.25) are designed to service the lower buried part
of the rocket installed on the launch system and also the necks of the
fill collector of the launch structure.
122~
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_ '
~ ~ -
2 ' ~
3 4 �
~ / .
S
7 6
Figure 4.23. Rotary type tower:
1-- crane; 2-- tcwer trunk; 3-- telescopic support; 4--
braces; 5-- central support; 6-- supporting frame;
7 rollers.
Before launching the rocket system, the cab is brought to the bottom of
the launch structure along the suspension rail and it is protected by a
heat shield in the enga.ne gas jet.
The service cable towers (masts) and cable ma.sts are used to bring the
electrical, fill, drainage and pneumatic lines to the rocket system. They
have different di.mensions (height to 100 meters and weight to several
hundreds o~ tons), and they can be stationary or removable (withdrawable).
The stationary towers are mounted on the launch system or beside it; such
towers have trusses (platforms) which are withdrawn to a safe distance
before or at the time of launch.
123� .
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. ~
~ .Y J,. . ~
~;F G?.
~,~e\.,''~r,'~
`~a . ~ �
. :`1,'K. . l~~ Y ~�Y
9
~ " :'s: ~
, ~ ~ "..`~tz ;r r~*~ - rt
r, '~*~n '
e ~ _
~~`j
+y+ . .
.
~ K~ ( F r1~ ~ . .
4:~
~ .:a;,1;.5':f -
~J;
a~.+,s t
'
' ~ p'
J;~" ~:.~;~:;.th
. ' ?*,i.
~A~j .n'.
~ ~
i
~`J~
FY{~ -r
, ..i�~' L~L
1
.
- Figure 4.24. Service tower of the "Saturn-V-Apollo" space -
rocket system
The withdrawable service cable mast (Fig 4.26) or cable masts usually are
. installed on hinges on the launch system, and at launch time they are
withdrawn to the required augle, using a counterweight or pneumatic
(spring) drive;. the kinetic energy during withdrawal is extinguished by a
hydraulic shock absorber.
~ If the communications lines with the rocket are coupled to the installation
and test facility (the vertical assembly building), the service cable towers
or cable masts are transported to the launch position ~ointly with the
rocket system.
- The stationary cable mast (Fig 4.27) is a structure through which the cables
are run to the upper stage of the rocket. At launch time, the plugs are
unplugged, and their ground sections together with the cables are dropped
from the rocket under their own weight.
The service tower cable for the "Saturn-V-Apollo" space rocket system
(Fig 4.28) is mounted on the upper part of the launch platform and together
with the rocket system is delivered to the launch complex by the caterpillar
carrier.
124
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- 1
I
:
r-. a
~ 1 I' y -
~
--r-- - ~ ~5'
~ ,
_ 6
.
~ -
--e----F --e---i-~--a-
~ � '
~ O i o ~ 8
. ~ JI~-9
- 14 i - \ l~~v
O ~ ~l
~ " - O � h
' ~ ~
. ~ p
- � i ~-�--s-~
. . . .
~
Figure 4.25. Service cab:
1-- drive for the displacement mechanism; 2-- central pads;
3-- carriages of the displacement mechanism; 4-- heat shield; ~
5-- telescopic bisectional columns; 6-- telescopic trisectional
columns; 7-- hydraulic cylinder; 8-- platform; 9-- telescopic
bridge; 10 ring platform; 11 rotary disc; 12 control
panel; 13 mechanism for turning the disc; 14 emergency
ladder; 15 gangway; 16 chain drive
The tower is a steel truss 116 meters high through which the fuel and
pneumatic lines, electrical and television cables, telephone lines, water
lines and other lines are run. The tower has nine folding arms; eight
fueling units are connected to five of them. A"clean chamber" with con-
ditioned air is mounted on the upper arm. Tt is coupled to the hatch of
the command compartment of the spacecraft and provides for entry and exit
of the astronauts. The feed lines made up of rigid or flexible lines are ~
_ joined through the crossovers to the lines laid in the tower, and they
have plugs connected to the fueling units of the booster rocket. The
service tower is equipped with 17 work platforms; all of the platforms are
are connected by two high-speed lifts which were used for emergency exit
from the spacecraft in case of an emergency and delivery of the crew to
the fast-exit chute of the launch structure which begins at the launch
platform.
125
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~g'~ ~C~f,~ 4'"y ~y,~g,, } rf
~ lR:,' ' ,~�k ,~1
~w. . : t~�~~y: `u~y.l ~3~.,I:
rt. ~-,y5'' �fw ~ ~ ~ ,s ~ T
~'~f y~;5~ t L~r
e+~~'~H'"~yt ~ e ~ x
r~~,.'
. .~-A w~~~' .\~r'~.,�
,y..~ x
~ ~ t ^'95 } y'E ~ ? 11"i'b ~:'t~, hro. .
s ~.'.r 4` ~ 4 ~ ;rt c . i t ~i
~ y r.~
Y - %~R� F C ~ ~
~ ~rr~`- 'a.'I~ ~ ~ a8 zdt5 ry, ,~k e Z''~{'.~jt' ; . ~ i`
~~~tf%~,h`+G.~e)~ -Zj tF~~~`.~
~41'r s u~~ . i~ r,i s~
~ . ' x ' d'~ .
. n ~4
o
'f't
a -
i
-
t
Figure 4.26. Service cable mast for the "Soyuz" space
rocket system -
A rotary arm drane with a capacity of 22.5 tons is installed on the tower. -
It can be controlled ~rom any platform using a portable control panel.
- 4.6. Electrical Equipment
With respect to amount of intake electric power it is possible to compare
the cosmadrome with a large modern plant equipped with complex highly
automated systems and units. The electric power users at the cosmodrome
- are the electric drives of the trusses and service platforms, the electric
motors of the elevators and hoists, the electric pumps of the servicing
126
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systems, the pneumatic pumps and ejectors, electric heating and air condi-
tioning systems. All of them require 380/220 volt, 50-cycle 3-phase _
alternating current to power them. The sources of this elect~ric power are
~ called primary current sources. At tYce cosmodrome there are other electric
power users which require special forms of current and voltage for their _
power supply. Thes~ include the technological operations control systems,
the measurements and functional monitoring systems, the guidance system,
' and so on. The special forms of currents are required also by the on-board
equipment which is powered from ground power supplies during preparation
of the booster and spacecraft for launch. The power supply for the on-board
equipment comes from special currents, and their sources are called
secondary current sources. The secondary current sources are combined into
,
~ f~~ -
e.t : H..
* , { ~r~ K�~~
y ~
~{:~i(
.w
~
~~t _
'.Y '
!
J
7.t~,Rr; ~ .,A
d'- 9'~ i
. ~ ~y~ ~~~4 f
~y *
i ~
a~ t . . ..~r~ . ~ I ~ ~ I .
~
, , ~ . c~ I I ~~T s _
;i~ ~
r ~~~r:
,r r
-
.i"
Figure 4.27. Cable mast at launch time
the so-called special currents ground electrical supply system (SNEST)
which generates alternating and direct current of di~ferent voltages and
frequency. It must be noted that in the measurement and control systems;
secondary electric power supplies can also be used as individual feed
modules.
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Since the launch time of the space rocket system usually is strictly regu-
lated, interruptions in the electric power supply can lead to postponement
- of the launch to a given time and sometimes to more serious consequences,
emergencies. This is why special attention is given to the problems of
reliability and quality of the electric power supply.
The high reliability of supplying electric power is achieved by the follow-
ing:
The application of highly reliable elements and assemblies of the systems;
Duplication of the primary and secondary electric power sources and cable
couplings, the application of ring systems, and so on;
Clear organization of the work and high qualifications of the service
personnel.
The sour.ces of ground supply with electric power are divided into the primary
ones which include the industrial power systems, diesel electric power
plants and chemical current sources and the secondary ones, that is, the
devices that convert the electric power of the primary 380/200 volt, 50
hertz current sources to currents and voltages necessary to power�~the
ground and on-board automatic control and measurem~nt systems.
The direct current converters (unstabilized and stabilized), the mechanical
AC converters and static AC converters with increased frequency are used
as secondary electric power sources.
The system for supplying the cosmodrome and the rocket system with electric
power is presented in Fig 4.29.
The industrial power systems are the basic source of the electric power
supply on which the requirements of increased reliability and maintenance
of high voltage stability are imposed.
The electric power is fed to the launch complex through a step-transformer
(see Fig 4.29), after which it goes through special entrance shields to
the users and to the secondary power supply sources in the form of a three--
phase, 380/200 volt, 50 hertz alternating current.
The diesel electric power plants (DES) are, as a rule, a reserve (redundant)
source which provides electric power when the basic electric power trans-
mission lines fa3.1. A diesel generator is used as the source of electric
power in the diesel electric power plants, the power of which also
detexmines the power of the plant itself. The continuous operating time
for the 3iesel electric power plants is not regulated, and it depends on
the amount of fuel, that is, the capacity of the fuel tanks.
Usually automated diesel electric power plants are used which are auto-
matically started when the power is shut down from the industrial sources
and are capable of operating for a prolonged period of time without service
personnel.
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' ve�
� +1 I �f Z i
. . , ~ .
1 .
12 .
. . ,
, . .
. 3
� ;~ti,~ ;
. t ,
11 . ~ ~Y~.
' 3 4 ,i m'
_ 4 y
. 5 .
6 .
.~i 10 _ 7
~ g 5 i
~o J
_w;.-
.k
12
~;~~r 9 .
' ,
. ~r~ .
. L
7 ~ ~
_ , Iw
$ i ! '~j .
b 00.:
: : _ , . ' : �
,
Figure 4.28. Cab1e service tower for the Saturn-V Apollo Rocket
System
1--control module for the pneumatic system of the equipment module;
2--module switch; 3--cooling module for the equipment module; 4--control
units for the pneumatic systems of the main engine of the third stage;
5--cooled gas (helium and hydrogen) feed module for blowing the third
stage tanks, cooling the main engine jacket and filling the tank for
whirling the turbine; 6--control module; 7--control module for the pneu-
matic systems o~ the auxiliary third stage engines; 8,9,10--control module
for the pneumatic system of the second stage engine; 11--gas (hydrogen)
feed module for blowing the second stage tank and cooling the engine jacket;
12--control module for the pneumatic system of the first stage enginee
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The diesel electric plants are equipped with units to maintain the normal
operating conditions, for protection, starting and stopping, emergency
warning signals and also a system for automatic monitoring and control of -
the current frequency and voltage, the water and oil temperature and
charging the storage batteries. In addition, manual remote contral is
provided which provides for starting and shutting down the diesel electric
power plants, switching the generator on and off, reclosure after the
response of the protection and elimination of the failure.
The diesel electric power piants can be portable, placed in the beds of ~
trucks or on railroad cars and stationary, located in the cosmodrome
structures.
.
' - . 2 ~/lICQ6flVHb1C UCIlIOYM!!R([
~3C ,~eKmpuaNepuar
~ ~1~ A3/1 ~ ~3~ ~ .
I ~
8 nt ~ nT
) 4) B~ncRmpv~ne~ptt~rtt
mrrt (5) -
~ .
. j(6) nc ncc 7) (s) ~Mr ~Mr ~js~ ~
t ~
L- -
Cunoeera 1. OOusemernuvecnre
n a ~ r cucme+va
1~ ~,0 ~1 ~rl 1 ) 2 oceeu~exue (12) -
- Figure 4.29. System for supplying the cosmodrome and the
space rocket complex with electric power
Key :
- 1. electric power transmissicn lines mechanical type AC converter
2, diesel electric power plant 9. power drive
- 3. primary sources of electric power 10. technological process operations
4. trans�ormer substation control system
5. secondary electric power sources 11. measurement system
6. unstabilized DC converter 12. 1. general engineering systems -
7. stabilized DC converter 2. lighting
The chemical current sources provide direct current with a voltage of 30
and 6 volts to the measur ing control systems and also the emergency lighting
~ system, and they are used in the stationary automatic monitoring control
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units,in the portable lighting units and as a current source in "ready
reyerve"1 for the case of �ailure of the basic source, f.or which the so-
culled "buff:ered" .inclusion i~ practiced.
Inasmuch as the requirements of fire and explosion safety are imposed on
the measurement and control systems in a number of cases, it is possible -
to use dry chemical current sources (dry cells) which have high intern.al
resistance and low power. This permits insurance of operating safety of
the systems, for in the case of the appearance of a failure, the magnitude
of the current is limited by the internal resistanc e of the source. However,
the small continuous operating time, the dependence of the parameters an the
ambient temperature, the comparatively high cost complicate and limit the
use of the chemical current sources. In recent times the DC rectifiers _
based on semiconductor elements have been most widely used.
The ground electric power supply system for special current (SNEST) inrludes
the unstabilized DC converters (PS), stabilized DC converters (PSS),
mechanical type AC converters (AMG), the current distribution unit (TRU),
remote control panels (PDU), power distribution boxes (RSK) and a cable
network.
The unstabilized static DC converters are designed to supply direct current
to the systems and individual measurement and control instruments when
highly stable voltage and increased reliability are not required. The
voltage at the output of such a converter depends an the fluctuations and
variations of the input AC voltage and the mode current.
The stabilized static DC converters will permit us to obtain a DC voltage
at the output with deviatio ns from the rated by no more than +3%, and
in the best cases, less than +1,�;.
The mechanical type converters or electromechanical converters are electri-
~ cal machines which convert one type of current to another, with different
voltage, frequency, and so on. Depending on the purpose, they are
divided into DC-AC converters (w~.ic :~~nvert alternating current to
direct current or vice versa), DC converters (which convert the DC voltage),
frequency converters, and so on.
The current distributing devices in the ground electric power supply systems
to supply specialized currents play the role of the power commutators of
~ 1"Ready reserve" is the method of reserv3.ng or redundancy in which failure
of the basic source of power does not lead to interruption of the power
to the users, whereas in other reserve techniques time is required to
connect the reserve power supply.
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the output buses, static stabilized converters. These devices are made
in the form of individual bays of usually unitized construction in which
contactors and current protection elements are placed (automatic network
protection systems, fuses), terminals and oil seal entries for cables
used to couple the current distributing devices (TRU) to the output f eeders
of the static stabilized converters and load and also intermediate relays
for remote control of the contactors and the current and voltage quality
control instruments on the output buses.
The remote control panels are designed for remote inclusion of static -
stab~lized converters or connection and disconnection of their output
feeders in the current distributing devices directly from the point of
connection of the instruments the power users.
The distributing power boxes in the ground electric power systems for
- special currents are used as terminal units designed to connect users
remote from the static converters.
For the SNEST cable network, usually two types of cables are used: flexible
and stationary. The flexible cables with plugs are used to connect the
individual instruments (panels, bays, converters) entering into the func-
tionally independent equipment complexes located inside one facility.
The stationary cables which are soldered to the terminals inside the
instruments have armor protection and are designed for operation in un-
heated facilities.
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~
CHAPTER 5. FUELING SYST.EMS
5.1. General Information
The fueling systems are designed to fill the rocket systems with f_uel
~ components and compressed gases.
The fuels can be divided into high-boiling and low-boiling, two-componen~
and single-component fuels in accordance with their basic physical--chemical
properties.
The liigh-boiling fuels are liquids with a boiling point above 298�K at
atmospheric or somewhat increased pressure under operating conditi.ons which,
being put an the rocket tanks or in the ground system storage fac~lities
can be stored for a long time under the indicated conditions in practice
- without losses.
The low-boiling fuels are liquids with a boiling point below 298�K under
- operating conditions; they include liquef ied gases (liquid oxygen, fluorine, -
nitrogen, hydrogen, and so on) having so-called "cryogenic" (below 120�K)
the boiling point, from which they have received the name cryogenic.
Under ordinary operating conditions the cryogenic fluids put in the tanks
without thermal insulation are intensely evaporated at the expense of the
influx of heat from the environment and require the application of effective
- thermal insulation to decrease losses from evaporation. _
The two-component fuels are the fuels, the thermal energy of which is formed
as a result of oxidation of one component (the combustible) by the other
(the oxidizing agent) in the combustion process in the engine chamber.
The two-component liquid fuels can be self-igniting (if combustien begins
when they mix) and nonself-igniting (if additional means are needed for
combustion of them).
The single-component fuels are complex compounds capable of decomposing
under definEd conditions into simpler and more stable materials with the
release of thermal energy. The single-component �uels are used for
auxiliary purposes: light thrust engines (orientatiun and stabilization)
of the space vehicles and the upper stages of the booster rockets, for
turning the pump turbines, and so on.
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The eff iciency index of rocket fuels is the specif ic thrust the ratio
of the thrust created by the engine to th~ fuel consumption per second;
- the fuel having greater heat value and lower molecular mass of the combus- -
tion products has ~reater specific thrust.
Another important characteristic of fuel Is its density: the greater the
density, the smaller the volume and, consequently, the mass of the rucket
tanks. An increase in density of the added liquid fuel components is
achieved by cooling them or introducing a heavy inert admixture. This
procedure was proposed for the first time by Academician V. P. Glushko in
~ 1933. Among the assimilated fuels, the two-component fuels are more wide-
spread than single-component fuels, which is explained by their higher
specific thrust.
The fuels must have high specific thrust, high density, safety in handling,
the possibility of long-term Gtorage both under ground and space conditions,
low cost, and so on.
None of the existing liquid fuels fully satisf ies all of the enumerated
requirements; therefore in each specif ic case certain of their advantages
and disadvantages are taken into account, which is one of the causes of
the great variety of them.
� In world practice, among the high-boiling oxidizing agents the most ~aide-
spread is nitric acid, nitrogen tetroxide and nitric acid solutions with
nitrogen tetroxide; of the low-boiling ones, liquid oxygen. Studies are
being made with respect to the use of liquid fluorine, ozone and mixtures
of them with oxygen, for they have better oxidizing properties than oxygen. _
Amon; the high-boiling fuels broad use is made of kerosene, hydrazine _
and its derivatives (monomethylhydrazine and asymmetric dimethylhydrazine
NDMG); among the low-boiling ones, liquid hydrogen. Hydrazine is usually -
used in mixtures with other materials (thus, the fuel "aerozin-50" which
is widespread in the United States is made up of 50% by mass hydrazine and -
50% by mass di.methylhydrazine). At the present time studies are being made
of the use of pentaborane.
The bicomponent fuels (combustible and oxidant) usually are characterized
by the oxidizing agents, for they are the basic part of the fuel, and their
number is comparatively small. Thus, the fuels used at the present time
with the oxidizing agent liquid oxygen insure the greatest specific
- thrust; such fuels as kerosene, dimethylhydrazine and liquid.hydrogen are
used with them. The "oxygen-kerosene" fuel is best assimilated in rocket
engineering, it is cheap to produce and convenient in operation. This
explains its application in the first stages of the American "Thor-Delta"
booster rockets, the "Thor-Agena," "Atlas-Centaur," "Saturn-I," "Saturn-IB,"
"Saturn-V" and also the Soviet booster rockets. The oxygen and asymmetric =
dimethylhydrazine fuel has the greatest specif ic thrust for liquid-
propellant rocket engines of the oxygen class operating on high-boiling
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fue1; it is used on the second stage of the Soviet "Kosmos" booszer
rocket.
The "oxygen-hydrogen" fuel has a specif ic thrust that is 30 to 40% titgher
than the other assimilated rocket fuels and, in addition, is ideal from the
point.of view of environmental protection, f.or the products of its
combustion are water vapor. This fuel has been used in the uppe.r. stages
of such booster rockets as "Atlas-Centaur" (third stage) and "Saturn-V"
(second and third stages). With further development of rocket en~ineering
hydrogen can find application also in the first stages of the large space
rocket systems.
The f uels based on nitric acid and nitrogen tetroxide are significane7_y
- inferior with respect to specif ic thrust to the fuels based on axygen.
They are capable of prolonged storage, the duration of which depends to a
- significant degree on their corrosive activ3t;~; 'out they are to:sic. Such
fuels as kerosene, asytnmetric dimethylhydrazine, and tiydrazine are used _
with nitric acid. The last two form self-igniting fuels, and with nttrogen
tetroxide the fuels "aerozin-50," dimethylh_~drazine and monomethylhydrazine
are used (all self-igniting). The p~ssibility of prolonged storage without
losses, high density and self-ignitability explain the broad application
of these fuels in the multiple-action engines and in low-thrust engines
(for stabilization, orientation, braking, taken from other planets, and
- so on) and also in the booster rockets created on the basis of the combat
rockets where the basic requirem~nt is insurance of prolonged storage of
the booster rocket in the filled stage. I`or example, in the engines of the
"Apollo" spacecraft fuel bas~d in nitrogen tetroxide is used: the "nitric"
acid and asymmetric dimethylhydrazine fuel in the last stage of the
"Atlas-Agena" booster and the "Thor-A~ena" booster, and the nitxogen
tetroxide and aerozine-50 fuel is used for the "Titan-III" booster rocketse
Hydrogen peroxide in various concentrations, hydrazine and asymmetric
dimethylhydrazine are used as single-component fuels. Thus, ir_ the "Vostok"
booster rocket the products of deco~position of hydrogen peroxide were used
as the wor�king medium for the turbopumps, in the upper stage of the -
"Atlas-Centaur" rocket it was used to operate the auxiliary engiuze and
turn thP booster fuel pump drives.
In space rockets for individual systems, cryogenic tluids are used (oxygen,
nitrogen, hydrogen and helium). Electrochemical processes between _
gasified oxygen and hydrogen in the fu,l Plements of the electric power -
supply system provide electric power for the on-board equipment of the
space vehicle; oxygen and nitrogen in the life support systems of the
spacecra~t compensate for losses of oxygen during breathing and restore the
atmosphere of the spacecraft when performing operations in space or docking
with other vehicles; liqui.d nitrogen, hydrogen and helium, and in some
cases, solid hydrogen and nitrogen f ind application in cooling the infra- ~
red radiation receivers, quantum amplifiers, and so on.
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The application of cryogenic liquid fuel components with reduced tempera-
ture with respect to saturation temperature and atmospheric pressure
(the so-called supercooled or underheated liquid) has great advantages. _
Such a reduction in temperature in practice is possible in a wide range,
to the formation of a mixture of solid and liquid phases (the so-called
slush). ~
Slight supercooling of the fueled liquid permits a significant decrease in
the vapor formation in the booster rocket tanks during fueling process
and makes it gossible to have on-board drain valves of smaller size.
Filling the tanks with deeply supercooled liquid increases the time fhe
fueled booster rocket is reac~y to launch without losses, for the supercooled
liquid must first be heated to the boiling point and only then does it
begin to evaporate. As a result of an increase in density, such a liquid
permits a decrease in volume and mass of the rocket tanks and an increase
in the drainless storage time under space conditions. Here the maximum
effect is achieved by conversion of the cryogenic liquid to the solid state
although the cryogenic components of the fuel in the solid state are in
practice not used. For use in the engines, the deeply supercooled liquids
(to the ternary point) and also slush in the form of finely disperse flow-
able mix are convenient.
In addition to the liquid fuel components, the space rocket systems are
filled with compressed gases which ha.ve such properties as simplicity of _
accumulation and capacity of energy conservation for a prolonged time in
- the state of being compressed to high pressure, safety (the nitrogen and
helium) for operation in various media as a result of the inertness. These
properties make it possible to use compressed gases as a source of energy
for on-board equipment with pneumatic drive or a solid state in the tank
blowing systems, various types of purging, and so on.
In the space rocket systems . nitrogen and helium have come to be widely
used as a result of f ire and explosion saf ety with respect to the working
environm~nt and absence of condensation at low temperatures. In this
respect compressed helium is the most a11-purpose gase, then nitrogen
which has limitations with respect to condensation at low temperatures.
Compressed air usually is used in systems that are fire and explosion safe
with respect to gaseous oxygen.
Compressed gases are put in the open bottles (banks of bottles) of the
corresponding on-board pneumatic systems. Of ten one bank of bottles
provides the operation of several pneutnatic systems (blowing the tanks
and purging various engine elements). A�ter filling and before launching
the rocket, a constant makeup of the bottles and supply of the gases for
operation of the on-board pneumatic system takes place in the pre-launch
period; the on-board pneumatic systems are converted to compressed gas
feed from the on-board bottles only dirPctly before launch.
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In some cases, in order to increase the density of the compressed heliun?
tliat is put in tlie bottles and in order to decrease tl~e weight of tlir
pneumatic systems, the banks of bott_les are placed in a container fj.lled
with cryogenic liquid. Thus, on the "Atlas-Centaur" rocket, the heliuni
for blowing the fuel tank and purging the engine in the first stage is
in the bank of bottles placed in a jacket with liquid nitragen which is
poured into the ground system directly before launching the rocket, and in
the f irst stage of the "Saturn-V" rocket the helium for blowing the fuel
tank is in four bottles (volume 0.88 m3 each) installed inside the oxygen
tank.
Depending on the type of on-board pneumatic systems, the on-board bottles
are filled both at the filling station and at the launch complex. Befoxe
f illing with compressed helium and nitrogen used for operation in fire and
explosion-hazardous environments with respect to air or in working environ- -
ments with low temperature, the on-board bottles are purged with the
gas they are being filled with in order to remove the air to the admissible
concentrations.
The schematics of the ground filling systems depend significantly on the
structural design of the pneumatic hydraulic system of the rocket. From
the point of view of filling, both systems on board and ground are
parts of a united system capable of functioning normally only under the
condition of close interrelation of its component element~.
Basically the structure and the pneumohydraulic system of the tanks in the
filling section are determined by the physical-chemical properties of the
components with which they are filled, the amount (batch) and method of
obtaining the given batch in the tank.
- The structural design of the tanks is essentially influenced by the tempera-
ture at which the fuel components are put in them. Thus, tanks for cryogenic
fuel components have thermal insulation to protect the liquid with which
they are being filled from evaporation at the launch complex an3 in outer
space and also to protect against aerodynamic heating when flying through
the atmosphere.
The presence of thermal insulation on the tanks filled with liquid fuel ~
components (supercooled oxygen, hydrogen and so on) with a temperature
equal to or less than the condensation point of air prevents condensation
of the air on the walls of the tank and significantly decreases the
evaporability.
Thus, for hydrogen tanks of the American booster rocket "Saturn-I," E~'
"Saturn-V" and "Atlas-Centaur" plastic �oam insulation is used which is
blown with uncondensed gas (helium) to prevent condensation of the air
_ on the cold walls of the tanks. Thi.s blowing increases the thermal conduc~-
tivity of the insulation, but under space conditions the helium volatilizes,
and the ~aff ectiveness of the insulation increases sharply. The sampling
of the helium for analysis after blowing permits monitoring of the seal
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of the tank and the li.ne and the adoption of necessary measures to prevent
emergencies. Thus, fo:: example, the seal of the hydrogen tanks of the
"Saturn-V" is control].ed while they are being filled. _
The hydrogen and oxygen tanks of the life support systems and power supply
systems of the "Apollo" spacecraf t have insulation with several reflecting
shields placed in the vacuum space between the inside and outside walls
- of the tank. This insulation called the vacuum shielding is most effective
for prolonged storage of cryogenic liquids under ground and space conditions
although it causes some increase in mass of the tank as a result of the
outside wall.
In the space rocket systems a multilayer shielded insulation with outside
soft thermal cover of insignificant mass is used. In order to protect the
insulation from the condensation of air on the cold walls and crushing of
the insulzting shields (which are made of thin f ilm) by the outs3.de pressure,
the thermal cover is blown with uncondensed gas during the filling process.
After the vehicle (the stage of the rocket) goes into space the cavity of
the thermal seal is connected by opening special valves and diaphragms
to outer space, and the insulatian begins to function as a multilayer
vacuum shield.
The tanks of the booster rockets for cryogenic liquids with a temperature
above the condensation point of air usually are not insulated, although
individual sections of a tank have local insulation to prevent the effect
of low temperatures on ths instruments and elements of the engines. Losses
from evaporation in the tank after filling are compensated for by additional
f illing (ma.keup) of the component. During makeup, such tanks are covered.
with a layer of frost formed as a result of freezing of the moisture from
the surrounding air; the layer of frost to some degree lowers the heat
influx, playing the role of a type of insulation.
The filling with supercooled cryogenic liquids has its characteristic
- features consisting primarily in a decrease in pressure of the saturated
vapor of the cryogenic liquid as it cools and secondly, in the formation
of a signif icant thermal layer of f illed cryogenic liquid along the height
of the tank as a result of the heat accumulated by the structural elements
of the tank and f illing through the filling valve which is usually located
below the tank. Therefore, in order to avoid implosion of the tank as a
resuZt of the rarefaction created in it and to keep atmospheric air out
which is capable of destroying the r_oanposition of the liquid with which
the tanks are filled, the tanks az�e f illed with the saf ety drain valves
closed at constant excess pre~~ure of blowing by uncondensed gas. The
temperature stratificat~~r~ which is undesirable for operation of the engine
is eli.minated by m~~j,ng the liquid during the filling process or after it.
The liquid is mixed by passing uncondensed compressed gas through it
(so-called bubbl3.ng) or feedi,ng the liquid with which the tanks are filled
from the top through special manifolds.
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During filling with cryogenic liquids and in the subsequent period before
_ launching the rocket, the danger of heating up and boiling the liqua.ct
_ arises i.n the intake lines of the engine which is usually located
appreciably below the tank as a result of inflow of heat from the environ-
ment. This danger can lead to rupture of the lines. In order to prevent
this, a circulating loop is used (tank-intake line-circulating line-tank)
which operates as a result of natural convection, for elimination o.f. wh~_ch
- gaseous helium is fed to the circulating line.
When filling with toxic and self-igniting fuel components, their vapor is
removed through split drain connections to a special ground neutraJ.izin.g
_ system. For saf ety the exit arc.as of the drain lines in the tanks arP
selected, as a rule, at diametrically opposite locations. Before filling
with liquids (liquid hydrogen) that is fire and explosion hazardous with
respect to air, the air environmei!.t of the tank is replaced by a neu~ral
environment, and during the f illin~ process, these vapors are removed
through the pre-ignition or diluting systems to a safe concentrati~n.
The fuel tanks of the space rocket systems, depending on the supplied dose,
have a volume from tenths to several hundredths and thousandths of cubic
meters. The small volume tanks which are used for auxiliary engines and
are designed for quite high inside pressure permit evacuation of their
inside cavity, whi~h permits application of the simplest filling system
(the fill line and tank) called drainless. In this case the filling pro-
cess consists in preliminary evacuation of the tank in order to remove
air from it and subsequent filling of it with a given amount of liquid.
- This fill system usually is applicable for high-boiling components.
When adding low-boilin~ components by the drainless system it is necessary
first t~ cool the structural elements of the tank to the temperature of
the adc:~d component and to maintain this temperature during the time the
rocket is on the launch system. This significantly complicates the struc-
tural design of the rocket and ground equipment; therefore the drainless
system for filling with cryogenic components, as a rule, is not used.
In order to decrease the mass of the structure the large tanks (5 m3 or
more) are not designed for evacuation; they can only withstand small
internal excess pressure basically determined by the operating requirements
of the engine. These tanks are filled with discharge of pressure from
- their gas cushion through drainage safety valves in accordance with more
complex f illing system (the fill line-tank-drain line). During the
process of filling and draining by this system the internal pressure o�
the gas volume (cushion) is constantly monitored by gauges or signal ele- �
ments; w:ith an increase in pressure above the admissible, ait instruction
is automatically put out to stop the ~illing process, which is reinitiated
only a~ter establishment of normal pressure.
In the majority of cases the fuel components are drained from the tanks
with the drain-safety valve closed, which insures the required cleanness
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of the inside volume of the tanks and, in addition, accelerates the
drainage process. In order to avoid implosion of the tank by atmospheric
pressure under these conditions when the inside pressure drops to a defined
amount the on-board blower is sw3tched on, and in some cases the safety
drain valve is opened.
The seal of the tanks of the booster rocket stages and the space vehicles
designed for operating under space conditions is insured by using a
f itting with small diameter of the passage cross sections and the applica-
tion of f ill systems with minimum number of f i11-drain and discharge
lines. In addition, the tanks of the space rocket systems are carefully
checked for seal during the process of preparation at the engineering
complex.
When filling the tanks, the precision with which they are filled to the
given amount (batching) has great significance, for the excess mass on
board the rocket, especially in the upper stages and on the space vehicle,
leads to a decrease in useful load. The required accuracy is insured by
selecting the filling cor,ditions and the batching method.
The filled dosage is measured by the ground f illing system means (the
so-called external dosage), the devices installed in the rocket tank (the
inside dosage) and a combination of ineasuring devices of the ground system
and the rocket.
With external batching the required dosage is automatically measured by
special ground units (mass or volumetric batching) entering into the
f illing system composition. The mass batcher measures the mass of the
f illed liquid directly, using high-precision devices of the balance type;
the volumetric batcher measures the volume of the dosage or the volumetric
- flow rate using the measuring calibration tanks or the volumetric flow
meters, as which liquid meters are used which determine the amount of
liquid f lowing by the number of displacements of the servoelement and
various devices (turbine, choke, ultrasonic) which measure the speed of the
liquid in the line. By monitoring the liquid temperature in the batcher
and knowing its chemical composition it is possible to establish the
density of the liquid at the time it is added and determine the mass dosage.
When batching comparatively small amounts (tens to several hundreds of
kilograms), the mass batchers are used which insure greater accuracy than
the volumetric ones and do not require the introduction of temperature
corrections for taking into account the variation in density of the liquid
or corrections which take into account saturation of it by its own vapor.
Increasing the filling batch leads to an increase in the size of the
batchers which becomes commensurate with respect to volume with the ground
storage and to complication of the �illing system. The measurement of
a large batch in partsis inexpedient, for the batching error increases.
140
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Therefore for batching large amounts of liquid, the method of internal
batching is used in which the functions of the batcher ~nd the tank are
matched. When determining the dosage in the tank, overflow cranes or a.
more complex monitoring system is used which in the presence of calibration
of the tank are used to determine the volume of the batched liquid, record-
ing it or several defined (discrete) positions of the level or continuously
measuring it in the process of filling the tank. At the present time,
capacitive, inductive, manometric and ultrasonic systems have become wide-
spread for monitoring the level. In these systems the accuracy of ineasuring
the volume of the added component depends on the accuracy of installing the _
sensitive elements and the filling flow rate on completion of batching
selected in such a way that during the time of closure of the cutoff valve
by the signal from the system which monitors the level, the error with _
respect to ama.unt of batched component will not exceed the admissible
error. -
In some cases, combined batching is usen, which consists in filling the
tank to a def ined level measured by the rocket means with subsequent drain-
age of a pr~cisely measured excess batching to the mass batcher. This
method makes it possible not to ad~ust the level monitoring system in the
tanks for various flight programs.
5.2. Ground Fueling Systems
The space rocket systems are fueled with liquid components using the
corresponding systems of the space center making up a significant part of
the ground equipment and playing an important role in the process of pre-
launch preparations. It is sufficient to state that the mass of fueled
components is up to 90% of the launch mass of the modern booster rocket
using liquid-propellant rocket engines. The filling systems to a high
degree determine the structure of the space center and essentially influence
the outcome of the space experiment itself. Two versions of servicing
space rocket systems are possible:
Servicing the rocket systems with all fuel components at the launch complex
after installation on the launch system;
Filling the tanks of the space vehicle with high-boiling component at the
filling station of the engineering complex and the booster rocket tanks
(also the space vehicle if cryogenic fluids are used) at the launch complex.
The first version insures greater operating safety with the space rocket
system at the engineering complex, but it complicates the process of pre-
paring it �or launch and increases the number of pneumohydraulic "ground-
on-board" connections. The second version permits the number of fueling
systems and "ground-on-board" connections at the launch complex to be
reduced and also the flow chart for the pre-launch preparations of the
rocket to be reduced with respect to time, but it requires the performance
of a number of ineasures ai.med at insuri^g the required thermal conditions
and safety when transporting the fueled vehicle and during prolonged stays
at the launch position.
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The classification of ground fueling systems is presented in Fig 5.1.
The number of fueling systems depends, as a rule, on the number of fuel
components put in the space rocket, although for fueling essentially
- different on-board systems with identical components, various fueling means
are used. Thus, the stages of the "Saturn-V" booster rockets and tanks of
- the power system for the "Apollo" spacecraft are filled with hydrogen from
various ground systems.
The space center fueling systems are distinguished with respect to fueled
components, the magnitudes of the batches, the number of fueled tanks, the
method of supplying the fuel, the peculiarities of the schematic diagram
and structure of the equipment. The basis for all the systems, in spite
of the indicated differences, is the common schematic diagram: a storage
with means of feeding the component pipelines with fitting user (the
tanks of the booster rocket or space vehicle). -
The storage is designed to store a component and is made up of one or several
tanks. The tank usually has several outputs for connection to the lines
from the portable transport means, discharge of the component to the fuel
lines, draina.ge of gas from the tank cushion, supplying of gas to the t.ink
blowing system, and so on. In order to simplify the structural design and
layout, the number of ou~puts is decreased as much as possible, combining
some of them. In order to maintain a defined composition (condition) of the
stored fuel, storage under excess pressu~e, sealed connections and cutoff
fittings, chemical analysis, periodic drainage of the liquid, cleaning o� -
inside spaces, and so on are used. For each component, a special method of
performing chemical analysis is developed in order to detect the micro-
impurities, including determination of the sample-taki.ng process. _
High-boiling liquids are stored without losses; cryogenic liquids are stored
with small losses as a result of the application of highly effective thermal
insulation or without losses using the devices for return condensation of
the vapor. _
The storages are filled from the portable transport means delivering liquids _
from the fuel storage houses of the space center, or directly from the plants
as they are produced. The volume of storage must be designed to meet all
the requirements for the given component during the technological cycle of
preparing the rocket system for launch, considering single or double repeti-
_ tion of the launch in case of postponement or early launch. When performing
the calculations, the possible irrecoverable losses (drainage, evaporation,
and so on), the unreachable remains in the storage tanks and the tanks of
- the booster rocket, the quantity of the cowponent it takes to �ill the
service system lines and also the possible increase in the fueled batch
for varioLS versions of the booster rocket are taken into account.
In the service systems two metho~s of supplying the components are the most
widespread: forced and pumped.
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143
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The first method insures the given pressures and flow rate by creating a
blown pressure in the tank and the second, by the operation of a pump.
The application of ttie forced method of feed is limited by the admissible
pressures in the tanks, for the achievement of the required strength
characteristics (specially for the high-volume tanks) leads to a signif i-
cant increase in mass and complication of the structure.
The pump method of feed in practice insures any pressures and flow rates,
although it requires continuous flow of the liquid without bubbles for
stable operation of the pump. The bubbles formed in the low-pressure zone
(the so-called cavitation phenomenon) and filled with liquid vapor can
lead to interruption of the operation of the pump. In order to avoid this,
the total pressure at the input to the pump is increased by blowing the tank.
In this case, the combination of forced and pump feed methods are obtained,
that is, the combined method of feeding the fuel components.
The fuel flow rate determines the filling time. For modern space rocket
systems, insurance of large flow rates does not cause any theoretical
difficulties. Considering the property of high-boiling liquids to be
stored in practice without losses, the tanks are f illed with them in advance,
several days before launch. Here the filling is done in a two-step operation:
the basic flow (to 90-95% of the given batch) and the small flow (to the
given batched amount). In contrast to the high-boiling liquids, in order to
decrease the time of the low temperature effects on the rocket elements and
the losses from evaporation, t?~e fast-evaporating (cryogenic) liquids are
put in the tanks several hours before launch. In this case in order to
meet the requirements connected with the structural peculiarities of the
rocket and for greater accuracy of batching, multistage filling conditions
are used. Thus, the filling of the SI stage of the "Saturn-V" booster -
rocket with oxygen is accomplished in the following mode: 1135 liters/min
to cool the tank; 5680 liters/minute (to 5% of the fueled mass) to exclude
large loads on the structural elements of the rocket at the beginning of
fueling; 37850 liters/minute (to 95% of the fuel~d mass) the basic high-
speed fueling; 5680 liters/minute to the level gauges signal to stop batch-
- ing. Then makeup takes place at 1890 liters/minute to compensate for losses
to evaporation.
The simultaneous f eeding of the oxidant and the fuel makes it possible to
reduce the rocket fueling time. However, in practice, especially when
using self-igniting and fire ~.nd explosion hazardous components, the rocket
system is fueled in succession and in some cases, modulaYly.
The fuel component is fed ~rom storage and goes i.nto the rocket tanks
through lines made up of pipes and fittings for the liquid and its vapor.
The layout of the lines of the fueling system connecting the storage and
the user 3.s determined primarily by the pneumohydraulic layout of the rocket
tank and the requi.rements with respect to insuring flow regimes. For
liquid components it can be si.ngle or double line.
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The single line layout provides for filling, drainage and for cryogenic
components and topping off through one fuel line having branches in
accordance with the number of fueled tanks. This layout has become the
most widespread.
The double-line system provides for circulation of the fueled liquid in
the tank in case of thermostating by the method of simultaneous feed and
drainage. It includes the f eed and drain lines connected to the corr.espond-
ing tank and ground storage lines. The fueling line is made up of one or
several pipelines if it is structurally disadvantageous to have one large-
diameter line.
The fuel f ittings are devices which insure a given flow of the component,
adjustment of the f low and curtailment of f eed. These f ittings include
valves, gate valves, regulators, chokes and so on. The valves insure
seal (or with the given degree of seal) separation of two sections of the
pipeline and they are d ivided with respect to functional contribut2s into
the shutoff valve for stopping the feed and discrete regulation of the
flow; drain valves for discharge of gas, liquid or a mixture of liquid
and vapor from individual sections of the lines and tanks; safety valves
for automatic dis~harge of excess pressure and check valves for automatic
pass ing of liquid or gas in only ~ne direction.
The shutoff and drain valves are controlled remotely using various drives
(pneumatic, electric, electromechanical). The valves are deslgned for
complete shutoff of tlie lines, and they are devices with manual drive.
The chokes are used to regulate the degree of cover ing of the line and
_ they are with manual or electromechanical drive controlled remotely.
In some cases, in order to insure complete separation of one volume for
another in the tanks or lines, diaphragm assemblies are used in which the
seal ing diaphragm is cut off automatically with the given pressure gradient
or forced using pneumatic drive.
The c utoff valves and drain valves, the gate valves and chakes can have
signa.ls of the extreme (open-shut) and intermediate positions pro�:riding
for remote monitoring of their operation.
In the filling systems various types of automatic regulators are also used
(for example, liquid, which insures variation of the liquid flow depending
on the pressure variation in the gas cushion of the f illed tank).
In order to prevent mechanical i.mpurities from getting into the f.uel tanks
there are filters. The liquid or gas is purified to remove impura.ti.es
- both by passing it through porous or lattice materials of filtez elements
and by centrifugal effect. The puri�ication method d~epends on the speci.~'ic
operating conditions of the fueling systems.
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All of the operations performed using the fueling systems in the pre-launch
preparation process can be divided into preparation, basic, final, post-
launch and auxiliary.
The preparatory operations include checking out the equipment of the systems
for operation, the bringing of the stored component to the required
parameters (with respect to quantity, temperature, pres~ure); taking
samples for chemical analysis; connection (coupling) of the fuel, discharge
and drain lines to the corresponding valves or flanges of the on-board
connections of the rocket with subsequent checking of the connections for
seal; initializing all of the control system elements, and so on.
The basic operations include the preparation of the inside cavities of the
rocket tank (for e:{ample, replacement of the air atmosghere by a neutral
atmosphere for f ire and explosian hazardous components with respect to air),
filling the f ill lines.with~the components, filling the tanks, makeup,
thermostating the fuel, draining the components from the tanks in case of
postponement of the launch.
The final operations include correction of the level (topping off to the
given batch), relieving the on-board and ground fill lines of liquid and
gas (draining the lines), disconnection (uncoupling) of the fuel and -
drain lines for the booster rocket, and so on.
The postlaunch operations include draining the remains of the fuel
component from the fuel lines, replacement of the throwaway assemblies,
conversion of the system to storage conditions.
The auxiliary operations include f illing the storages from the transports,
technical servicino of equipment, and so on.
The participation of the controlled fueling system elements in the per-
formance of the operations is different (from 50 to 100% of the total -
- number); therefore for large and complex fueling systems with respect to
composition, automatic and semiautomatic technological operation control
systems are used.
The automatic control systems provide for controlling the system elements
and monitoring their operation in the automatic mode, and the semiautomatic
control systems provide for only part of the technol~ogicai operations in
the au~omati.c mode.
In addition, these systems insure the possibility of remote control of any
element during performance of the operations.
The operation of the fueling systems and their elements is controlled by
the lighting of lights and transparencies on the service control panels.
When performing the technologi,cal operations, in addition to information
about the order of respozse of the control system elements it is necessary
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to know and monitor the basic parameters of the fueling systesn (pressure,
temperature, number of components in the storages and the rocket tanl:s,
pressure, temperature and flow rate of the components in the lines arid
~he feed units). The monitoring ef these parameters ia provided for b.y
the corresponding measurement means with output of the readings r_o secondary
instruments with visual scdles. The information obtained makes it possible ~
to estimate the state of the fueling system at any point in its operat~.on~
The equipment of the fueling system basically is placed in the st~rage ar~as,
_ It consists of tanks, f eed and other types of equipment providing for
storage an3 preparation of the components for fueling: the rout~~s fuL ~ayinb
the lines (fill-discharge and ~rain lines); at the service units, the
service cable towers (mass), the service towers and trusses providiri~ f_or w
bringing the ground lines to the rocket connection.
At the launch complex the f~.ieling system equipment usually is protected
from the destructive effects of the shock wave in case of a possible e~;p1o--
_ sion of a rocket by reinforced concrete arched structures banked with dir~
and service passageways capable of withstanding a defined load in case o�
_ explosion. The construction of such structures requires large means,
especially for large spherical tanks.
The placement of the tanks designed in strength respects for 4 defined load
from the shock wave in an open area essentially reduces the expenditures.
Thus, at launch complex No 39 the large spherical tanks with liqu3d oxygen
and hydrogen are lef t open at a distance of about 450 meters from the center
of the launch structure; their structural design is for an excess pressure
of 41 kPa, and the stability of the foundation with regard to shifting and
_ tilting loads which can occur in case of an explosion of the "S~turn-9"
booster rocket.
_ The equipment of the fueling system of the service units removed before
launch to a safe distance does not require special s?~ielding.
The fuel components storage facilities with f eed means can be portable and
stationary, The portable storages(tankers) are used to store a small amount
of fuel; they do not require special structures except accesaes and covered
platforms and especially they are advantageous for modifying the launch
complexes for vehicles with other fuel components. The stationary storages
are used to store a large (several thousand cubic meters) quantity of
fuel and for fueling the booster rocket tanks and space vehicles of the
sp.^ce rocket complex of the heavy or superheavy class.
~ The equipment of the service station is located in the main building
_ (batchers, thermostating means, vacuum e~uipment), i.n the storage builc~ing -
removed from the main building (tanks with ~acilities for storing the
component and preparing it for fueling) and in the main channels connecting ~
the ma.in building and the storage.
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Ti?e fuel sy5tems Cor fire and explosion hazurdous components both at the
f illing station and at the launch complex are separated by a safe distance
and are placed in isolated channels and structures.
5.3. Fueling Systems for Cryogenic Fuel Components
In modern rocket engineering cryogenic liquids are used as the fuel
components for engines, the operating means for the fuel elements of power
supply systems, the life support systems and the systems for blowing the -
space vehicles and rocket modules; coolants for supercooling other cryogenic _
_ liquids and compressed gases and also for so-called cryogenic purif ication
of the compressed gases to remove admixtures in the ground gas supply systems;
for special cryogenic systems installed in the space vehicles, and so on.
The cryogenic liquids are also used in the ground gas supply systems of the
space rocket complexes to obtain compressed gases by the gasification method.
Some of ttie data on the application of cryogenic liquids in the exi:~ting
and prospective space rocket systems are presented in Table 5.1, and the -
basic physical constants of the liquid oxygen, nitrogen, fluorine, and -
hydrogen, in Table 5.2.
Table 5.1
Some Data on the Application of Cryogenic Liquids in Booster Rockets
- Booster rocket ~'uel
~rocket stage) Oxidant Combustible component
"Jupiter" Liquid oxygen Hydrocarbon (kerosene)
"Atlas" The same The same
"Titan-I" The same The same
"Centaur" (stage) The same Liquid hydrogen
"Saturn-S-I" (stage) The same Hydrocarbon (kerosene)
: "Saturn S-II" (stage) The same Liquid hydrogen
"Saturn S-IVB" (stage) The same The same
"Kosmos" type rocket The same Asymmetric dimethylhydrazine
(upper stage)
"Vostok" type rocket The same Hydrocarbon (kerosene)
Let us consider the physical-chemical properties of some of the cryogenic
liquids that are most used at the present time.
Liquid oxyg~n 02 is one of the most effective cryogenic oxidants (it is
inf erior only to �luorine and ozone), it is available and cheap, which is
" explained by its large reserves and nature and simplicity of procurement,
it is nontoxic and in pure form is not explosion-hazardous; in the gaseous
- stage it is colorless and odorless.
148 t-
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Considering that at low temperatures many materials (hydrocarbon steel,
rubber, certain plastics and so on) become brittle and lose their
ductile properties, construction mater:ials in contact with liquid oxygen
must be ductile, strong and resistant to combustion. Such materials include
aluminum and its alloys, high-alloy stainless steels, copper and brass.
Oxygen supports combustion intensely and forms explosion-hazardous
mixtures with liquid or gaseous hydrocarbons, and porous organic materials
(sawdust, cotton wadding, rag, felt) impregnated with liquid or gaseous
oxygen; they are explosion-hazardous under impact or on ignition. At
ambient temperature the o~rygen vapor is heavier than air, it spreads over
the floor and can fill all of the low spots. When working with oxygen
service personnel must protect clothing and hair from impregnation with
gaseous oxygen (ignition of them on occurrence of a spark is possible), and
after work it is mandatory to air out the clothing. In facilities where
people work with oxygen it is categorically forbidden to smoke, light a fire
or use uninsulated sources of current. All of tha equipment used w3th -
liquid and gaseous oxygen must be carefully degreased, and any tools that
are used must be copp.er plated to avoid sparks.
- Liquid fluorine F2 is the most effective of the modern cryogenic oxidizing
agents, for the fuels formed by it have the greatest specific thrust and
density which makes its application highly prospective. Fluorine has high
chemical activity, it reacts with all organic and inorganic substances, -
on contact with it the majority of substances ignite. Metals can ignite
from friction in case of high flow velocity and also in the presence of
contamination in liquid fluorine. Some of the metals (iron, copper, nickel,
aluminum and its alloys) are resistant to the eff ect of fluorine as a result
, of formation of a strong film of fluorides prote~ting against destruction
on their surface. Liquid fluorine and its vapors are toxic and have a
strong effect on the eyes, skin and respiratory tracts. Fluorine vapors,
reacting in the damp atmosphere of air, form hydrof"luoric acid. The products
of combustion of fluorine-containing fuels are also toxic as a result of the
formation of corrosion-active hydrogen fluoride in them. The equipment for
liquid or gaseous fluorine must be carefully purified and passivated. For
operating safety tanks and lines for liquid and gaseous fluorine ~re made ~
with a nitrogen "jacket" which, as a result of the lower boiling point of
liquid nitrogen filling it provides for storage of the fluorine in transport
or filling tanks in the supercool state, which excludes losses to evapora-
tion and condensation of its vapor formed during the filling of the tanks.
The low excess helium pressure above the surface of liqua.d ~luorine protects
it from contact with atmospheric air.
Gaseous fluorine is neutralized by passing it over dry NaCl and CaC12 salts
and liquid fluorine, by using sodium or a solution of calcina~ed soda.
The application o~ fluorine compounds with chlorine and oxygen is also
possible as an oxidizing agent; they are safer to handle, they can be stored -
in tanks made of ordinary structural materials, and they are less toxic.
150
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;
FOR OFFICIAL USE ONLY
In the United States within the framework of the "Saturn-Apollo" program
research work has been done to replace oxygen by a fluorine-oxygen. mixture
and for testing engines for the upper stage using liquid fluorine. At the
present time the majority of problems connected with transporting liquid
fluorine and the process of handling it have been solved, and its a~pl-lca-
tion as a rocket fuel component is possible after solving quite compleac and
tedious problems with respect to its operation as part of the rocket module~
Liquid hydrogen H2 is one of the most effective cryogenic fuels. This is
a transparent, colorless, low-boiling (it is inf erior only to helium) and
light liquid. It is not toxic or passive in terms of corrosion. Liquid
hydrogen has a low boiling point (only 20 K above absolute zero), which
determines the large losses to evaporation, the low density, which requires
an increase in volume of the booster rocket tanks, energetic impulse fur
ignition 10 times less than for hydrocarbon fuels, high TNT equivalent -
~ (the explosion of 1 kg of liquid hydrogen mixed with ~he correspondin.g
amount of oxygen cr.eates energy equivalent to 10 kg of TNT), and its mi~~t~rPs
with air and oxygen are explosion-hazardous within broad concentratioz~
limits (4.1 to 74.2% of the volume for air and 4.6 to 93.9% for oxygen).
Accordingly, the application of liquid hydrogen in rocket engineering ha.:,
become possible only after the introduction of highly effective thermal
insulation combined with various ~easures to prevent the occurrencP of
dangerous concentrations.
Liquid nitrogen NZ is a cryogenic liquid, it is nontoxic, but improving the
concentration of the gaseous nitrogen in an atmosphere of close facilities
can lead to severe consequences; it is chemically inactive and imposes the
same requirement on the materials as liquid oxygen. Liquid nitrogen is
used as a source of gaseous nitrogen for blowing fuel tanks (the Soviet
"Vostok" booster rocket) and for the ground type gas supply systems.
From the above-investigated physical-chemical properties of cryogenic
liquids it follows that the low boiling point, low heat of vaporization,
large amount of gas obtained during evaporation and the large difference
in densities of the liquid and gas phases are common to them. These prop-
erties are taken into account during planning and design and oFeration of
t:he cryogenic f illing systems, the composition of which includ~as storage,
means of supplying liquid from the storage to the tanks, means of super-
cooling the liquid, the controllable filling and cutoff equipment9 means
of removal and discharge (drainage) of liquid and its vapors to a safe dis-
tance and means of evacuation of the thermal insulation cavities.
The peculiarities of the structural design of the equipment o� cryogenic
filling systems arise from significant changes in the physica.l-mechanical
- properties of the metals and their alloys at low temperatures. With a
decrease in temperature, as a rule, the strength characteristics (yield
point and fatigue point, rupture strength) increase, and the plastic
indexes are worse (impact toughness, relative constriction and elongation).
151
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FOR OFFICiAL USE ONLY
Tlie reduction of impact toughness of carbon steels is so great that it leads
to embrittlement of them and limits the application in assemblies and parts
operating at low temperature. Therefore in cryogenic engineering hydro-
carbon steels are only used to make jackets, supports, fasteners and other
elements not in contact w~.th the cryogenic liquid.
The metals that operate at low temperatures (inside vessels, pipelines and
fittings) have such requirements imposed on them as satisfactory static and
dynamic viscosity, vaccum density and even gas generation, stability of
structure under long-term loads, low capacity for ignition in an oxygen
environment (for oxygen equ3.pment). These requirements are satisfied by
alloyed steels of the austenitic class, aluminum alloys, nonferrous metals
and their alloys, and among the nonmetallic materials, plastics having low
thermal conductivity and high strength. In order to exclude the significant
thermal deformations in the lines, various types of compensators, bellows
or flexible metal sleeves are used.
The schematic diagram for filling tanks with cryogenic components is pre-
sented in Fig 5.2. ~
A storage facility for a cryogenic liquid consists of one or several tanks.
The tank is a structural element made up of an outer jacket to which an
inside vessel is attached through special slits or supports. The thermal _
deformation of the inside vessel is compensated for as a result of the
corresponding fastening of the jacket (suspension or installation on supports).
The evacuated space between the inside vessel and the outside jacket filled
with thermal insulation forms a thermally insulated cavity. The suspensions
and supports and also the pipelines joining the inside vessel and the outside
:jacket are additional elements (heat bridges) through which the heat is
transferred from the environment to the liquid.
The total heat influx to the cryogenic liquid is -
Qtotal - Rrad~gas~heat~heat bridge Qins~heat bridge'
where Qins is the total heat influx through the insulation;
Qrad is the heat influx as a result of radiation;
4gas is the heat influx through the residual gases;
Qheat is the heat influx as a result of thermal conductivity;
Qheat bridge ~s the heat influx through the thertnal bridges.
The thermal influxes are calculated separately for each type of heat trans-
fer, although in reality there is complex interaction of all o� the
components of the total heat in~lux.
~52
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FOR OFFICIAL USE ONLY
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FOR OFFICIAL USE ONLY
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APPROVED FOR RELEASE: 2007/02108: CIA-RDP82-00850R000200040010-4
FOR OFFICIAL USE ONLY
For a sufficiently deep vacuum in a thermally insulated cavity the heat
transf er by the residual gases is very small. By increasing the length
of the suspensi.ons, supports, lines and applying materials with low
therma.l conductivity it is posstble essentially to decrease the heat ;
influxes through the bridges. The heat influxes as a result of radiation
and thermal conductivity are decreased by tY~e application of heat insulating
materials with high reflectivity and low thermal conductivity.
In order to decrea.se the losses to evaporation in some cases the heat
capacity of the departing liquid vapor is used, cooling special shields
by the vapor in the thermal insulating cavity or different thermal bridges.
The losses from evaporation also depend on the geometric shape of the
container: they decrease with an increase in the ratio of the volume to
the surface area. The most advantageous shape is a sphere, but the installa-
tion of spherical structures is connected with defined difficulties.
A cylindrical shape with a length of the cylinder equal to its diameter is
more convenient; in this type of cylinder with electrical bottom and top
the ratio of the volume to the surface area is insignif icantly worse than
for a sphere.
- A stationary tank for prolonged storage of liquid oxygen (Fig 5.3) is a
horizontal cylindrical vessel with electrical ends with a volume of up to _
225 m3. The inside vessel made of alloy steel with multilayered vacuum
shielding insulation is installed on four mounts in the outer ~acket of
carbon steel, and it is fastened by three locators. After f illing the tank
with oxygen the vacuum reaches 0.01 Pa and is maintained for a prolonged
period of time as a result of adsorption of the gas molecules by zeolite
placed in special pockets with coils through which the hot gas is f ed for
regeneration of the zeolite. Between the inside vessel and the jacket in
the f eed and blowing assemblies located near the supports, bellows are
welded out~to provide for sealed exit through the vacuum cavity and compensa-
tion for displacements of the inside vessel with respect to the jacket
during cooling. The outer jacket is a saf ety diaphragm thxough which the
excess pressure is discharged in case of loss of seal in the inside vessel
with the cryogenic liquid.
The stationary vessel for storing liquid nitrogen, oxygen and argon
(Fig 5.4) is a vertical cylindrical vessel 66 m3 in volume which is fixed
in a jacket by two supports and is installed on four supports concentrically
entering into the jacket support.
The portable storages are analogous to the stationary ones, but they are
made of lighter metals and alloys and have a special attachment inside to
extinguish longitudinal displacements of the liquid during transportation.
The cryogen:Cc liquids in the tanks are stored under excess pressure and
without it. Excess pressure created by their own vapor as a result of
evaporation of the liquid in the tank (from thermal influxes from the
environment) or in a special heat exchanger and evaporator, and in some
154
FOR OFFICIAL L'SE ONLY
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APPROVED FOR RELEASE: 2007/02108: CIA-RDP82-00850R000200040010-4
FOR OFFICIAI, USE ONLY
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155
FOR OFFICIAL USE ONLY
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APPROVED FOR RELEASE: 2007/02108: CIA-RDP82-00850R000200040010-4
FOR OFFICIAL USE ONLY
cases also gaseous helium prevents atmospheric air from getting into the
gas cushion of the tank and it prevents variation of the composition of
the stored liquid. The required amount of excess pressure is insured by
a closed drain valve (open periodically to discharge pressure) or a
- special choke disc installed in the drain line.
In the general case the required amount of cryogenic liquid in storage is
Gstorage K~Gtank~pipe~makeup+EGloss~~np~ready~therm~
where K is the margin of safety taking into account possible versions of the -
rocket connected with increasing the filled volume;
Gtank is the amount of liquid put in the rocket tank in the given Yevel;
G i e is the amount of liquid going to fill the pipelines connecting the
s~orage and rocket tanks;
Gmakeup is the amount of liquid going to make up the tank, that is, replace
the losses from evaporation (this component is taken into account for the
f illing system with makeup);
~Gloss is the total irrecoverable losses of liquid to evaporation;
Gn is the amount of liquid remaining as a result of not being picked up
(i~ is determined basically by the geometric characteristics of the tank
and the layer of liquid having a temperature above admissible for feeding
- to the rocket tank as a result of heating from the blowing gas);
- Gready is the amount of liquid in storage insuring that the given readiness
of the storage will be maintained after it is filled to fill the rocket
without additional hauling;
Gtherm is the amount of liquid required to insure the thermostating of the -
f illed tank (this component appears only when thermostating with the appli-
cation of the circulation method under the condition that the thermostating
is realized by the reserves of previously supercooled liquid).
In the g2neral case the irrecoverable losses are as follows:
EG1oss-Gcool~blow~supercool' ~
where G~ool is the amount of liquid evaporated during cooling o� the metal
structural elements of the ~illing system and the rocket tank;
Gbl W is the amount o� liquid evaporated to create blowing of the storage
tan~CS;
156
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Gsupercool is the amount of liquid evaporated during cooling of the filled
liquid by the evacuation method (this component is taken into account only
when cooling the liquid by evacuation).
The amount of liquid required to insure thermostating of the filled tanks:
_ ~
Gtherm G therm'T'~~
where G'therm is the consumption of the liquid during thermostating;
T is the operating time of the thermostating system during one cycle to
obtain the given temperature in the tank;
n is the number of thermostating cycles.
LJhen using the gas cooling units or special cooling fluid insuring super_-
cooling of the filled cryogenic liquid withott losses, the value of
Gtherm can be taken equal to 1 to 3 hour flow rates of thermostating as
the cold storage unit for the initial thermostating cycles. When using
a cooling cryogenic liquid in the ground filling system it is necessary to
have equipment for storing it and f eeding it to the heat exchanger.
The reserves of the cryogenic liquid in storages fox large booster rockets
usually exceed the filled doses by 1.5 to 3 times. Thus, the ratio of the
amount of liquid in the storages providing for launching of the "Atlas-
Centaur," "Saturn-IB," "Saturn-V" booster rockets to its amount in the
tanks is approxi.mately equal to 2, 1.4 dnd 1.8 for oxygen, and 2.8, 2 and
2.3 for hydrogen respectively.
Feed Equipment -
In the cryogenic fuel component fueling systems, forced and pump methods
of supplying the fuel components and also combination methods are used.
In the forced method the blowing is crea.ted by a gas which comes from the
receiver or is obtained during the f illing process in the heat exchange
equipment (a gasifier) as a result of evaporation of some part of the
component with the help of a heat-exchange agent (for example, hot water
from the heating network) or from an electric heater, and in some cases
under t~he effect of ambient heat. The cryogenic liquid is fed to the heat
exchanger by a pump (in the pumped f eed system) or by gravity feed. In
this case the required hydraulic kit for arrival of the liquid is insured
by the location of the evaporator.
From the storages the cryogenic components are fed to the booster rocket
tanks through lines usua.lly made of individual sections, the length of
which is determined by the manufacturing and transporting possibilities.
The section of line (Fig 5.5) consists o� inside and outside (of the
jacket) tubes. The inside tub~~ is oriented with respect to the outside tube
157
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i
- FOR OFFICIAL USE ONLY
and is rigidly connected to it using supports made of material with low
coeffic.ient of thermal conductivity (fiberglass).
The inside tube is made of alloyed austenitic steel, and the outside
(~acket) from ordinary carbon or stainless steel. The space between the
tubes (the thermal insulation cavity) is filled with powdered or multi-
layered insulation with subsequent evacuation through an evacuation line
with a valve. In ~-r~?~:r to maintain and improve the vacuum, on the outside
of the inside line zeolite (for oxygen lines) or activated charcoal (for
hydroger~ lines) is placed in pockets. These materials have good adsorption
properties at low temperatures and reduced pressures. A rupturable saf ety
diaphragm is installed in the housirig in case of increased pressure in
the thermal ~.nsulating cavity of the lin~ (if the inside tube breaks its
seal).
A flexible metal h~se with multilayered vacuum shielding insulation (5.6)
is made up of inside and outside sealed hoses. The inside hose is oriented
with respect to the outer supports of material with low thermal conductivity
and together with the sleeve, tube and pocket with the atisorbent is
insulated by an aluminum-covered film. The outer hose is made of a sealed
= hose, an adapter and a tube, on one of which a bellows-type vacuum valve
and rupturable safety diaphragm are installed.
The sections of lines are connected by means of split bolt or unsplit welded ~
connections. In order to prevent accumulation of static electricity in
- the lines with split connections, special jumpers are installed to insure
reliable electrical contact between t:ie individual sections.
In the pipeline, depending on the specif ic conditions, thermal insulation
of different effectiveness is used. Thus, the lines for liquid hydrogen
- and helium and also the lines for long-term transportation of small flow
r.ates and great extent usually have powdered or multilayered vacuum
insulation, and the lines for brief transportation of large flow rates and
short extent, insulation made of foam plastic, fiberglass foaan and in some
cases (for liquid nitrogen and oxygen) do not have it at all.
The component feed (flow rate) is regulated using ~he valve module having
hydraulic characteristic of the line.
For a large diff erence in the ~illing and makeup flow rate the feed systems
have separate �illing makeup pumps and separate ground fill and topping
off lines. In order to obtai.n an exact level in the tank before finishing ~
up topping off, the drain valve o~ the tank is closed, which sharply
decreases the boiling of the added liquid and permits exact toppi.ng off of
the tank to the required amount.
When feeding the cryogenic liquid through the fill lines, the transitional
nonstationary initial filling regime is the most responsible and complex
dur.ing which the line is cooled to the temperature of the liquid. The
first lots of liquid on evaporation form a lar.~e quantity of vapor and
158~
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159
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fill the entire pipeline with it. The formed vapor is heated almost to
the initial temperature of the line and is discharged through a drain valve
located at the end of the line. As cooling takes place, the initial
section of the pipeline is filled with liquid; in subsequent sections, a
vapor liquid mixture is formed, and the vapor continues to escape through
~ the drain valve.
The cooling of the walls of the pipelines with the vapor makes it possible
to decrease the irrecoverable losses of the cryogenic liquid to cooling
the lines. During the cooling process fluctuations of the pressure are
possible which exceed the feed pressure (which can lead to forcing of the
liquid and vapor into the tank) and pulsations with respect to the flow -
rate respectively. Taking this into account, the longer pipelines are
cooled with low flow rate, and only after cooling and filling them with
liquid are they converted to increased flow rates. In order to prevent the
transfer of the pressure pulsations to the tank or the pump at the beginning
of the line, a check valve is installed, on closing of which provision is
made for transf er of part of the flow after the pump to the tank in order
to avoid disruption of the operation of the pump.
In order to decrea;~e the cooling time and f ill the long lines, the vapor
formed is dischargi,d through additional drain valves or special gas dis-
charge units (for example, the flow type gas-liquid separators) installed
at a defined distance from each other. Considering the great difference in
densities of the vapor and liquid phases, the f eed 'line is located with
so:ne rise in the direction of movement of the liquid, and gas discharge
devices are installed on the upper part of the line. For this system the
vapor is intensely discharged from the line, the liquid f ills the pipeline
faster, but in this case the losses to evaparation increase, for the cooling
takes place in the given case basically at the expense of evaporation of
the liquid. When f illing the separator with liquid the flow, rising,
closes the gas discharge.
During the movement of the cryogenic fluid even through an insulated tube,
heating of it as a result of the thermal influx from the environment, the
~ heat released in overcoming the hydraulic drag, the heat generation during
operation of the pump, and so on is unavoidable. This causes evaporation
of part of the liquid, it leads to the formation of a two-pha.se, gas-liquid
flow and it decreases the carrying capacity of the lines. Therefore, -
during the servicing process, an effort is made to obtain a single-phase
liquid f low which is possible duzing movement of the liquid in the state
of not being heated to the equilibrium temperature. The magnitude of the
underheating is selected so that during the heat release the formation of
the vapor phase is excliided, which makes it possible to stabilize the
hydraulic drag of the feed line, exactly to calculate and ma.intain the
flow parameters of the input to the tank. This regime is insured when
supplying liquid wlch increased pressure. In certain systems where obtain-
ing the single phase flow is impossible, in order to remove the vapor _
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f.ormed on the lines, gas separators are installed. In addition, the filling
of the taZk with a liquid which is cooled in the fueling system (below
the boiling point at atmospheric pressure) makes it possible to r.educe the
amount of gas in the tank as a result of heating of the liquid in the fill
line to a minimum.
Supercooling Means _
In order to cool cryogenic liquids, various refrigerating processes are -
used which can be unconditionally divided into external cooling and evacua-
tion cycles.
In the external cooling cycles, the heat from the cooled liquid is selected
- using the gas refrigerators or heat exchangers with colder cryogenic
liquid (coolant) by direct contact with the cooled liquid with the colder� -
surfaces of the indicated devices.
The operating cycle of the gas refrigErators is based on the compression,
heat exchange and expansion proc.esses of the working medium circulating
through a closed loop (for example, gaseous helium), which on coming to the
refrigeration chamber is heated and takes up the thermai load from the -
cooled liquid.
- In the evacuation cycles, the heat from the cooled liquid is removed at
the expense of evaporation at reduced pressure which is created by special
~ vacuum units.
_ Combinations of the cycles are possible.
In order to store a cryogenic liquid without losses and cool it in the storage
facilities, the "evaporator-condenser" system is used with the application
of gas refrigerators. The liquid vapor in the cushion of the tank is =
condensed by this system in the refrigeration chamber, creating a pressure -
gradient as a a result of which the liquid b egins to evaporate, and the
condensed liquid f lows back to the tank.
The cryogenic liqui.~ is supercooled in the f illing systems by various methods
(Figures 5.7 and 5.8). The most complicated method is the method providing
for cooling of the li.quid without losses to evaporation and requiring the
application o~ re~rigeration units of complex design; the simpler procedure
is tl-~e procedure in which the cooling takes place as a result of evaporation
~ of the basic cryogenic liquid or special coolant. If direct contact of
the basic liquid and the coolant is undesi~'able ~or safety reasons, an
ir,termediate inert heat transfer agent is used. The use of one supercooling
method or another is determined in the f3.na1 analysis by its e~fecti.veness,
reliability and economy.
lhl :
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The most widespread is supercooling based on the use of systems with the
application of coolant, evacuation and combination of them which, in spite
of the deficiencies (losses to evaporation) have relatively simple equiva-
lent. As a coolant in some systems either the basic cryogenic liquid
cooled by evacuation or another cryogenic liquid with lower temperature
(for example, liquid nitrogen or supercooling of oxygen) or mixtures of
cryogenic liquids supercooled by evacuation are used (for example, a mixture
- of liquid nitrogen and oxygen with a composition c,f 77% 02 and 2~% N2 with
hardening point of 50 K).
The evacuation method is based on the properties of phase transition of _
liquid to vapor with heat absorption. If the system is adiabatic, the
evaporation process will take place only as a result of the decrease of its
internal energy which is accom~;anied by a lowering of the temperature.
The equilibrium state ~i the two-phase "liqu~d-vapor" system is character-
ized by a relation d~fined for each liquid of the saturated vapor pressure
as a f unction of temperature. If in the two-phase "liquid-vapor"
system in the equilibrium state the saturated vapor pressure is reduced by
~p, that is, conversion is nade to the nonequilibrium state, the liquid
' turns out to be superheated with respect to the pressure obtained, which C_~R
leads to the initial evaporation process. If the pressure above the
liquid is kept constant, then with an adiabatic system the evaporation
process is accompanied by a decrease in temperature, and it will continue
until the "liquid-vapor" system reaches a new equilibrium state. _
- The use of hydrogen slush as the fuel component differing from boiling
- hydrogen by its high density and longer storage time without losses is of
great interest.
In the United States studies were made within the framework of the "Saturn-
Apollo" program with respect to the problems of developing methods of
obtaining, storing, transporting and fueling with hydrogen slush and also
an estimate bf the eff ectiveness of its application. The studies demon-
strated that hydrogen slush with 50% solid phase content has high absorbing
~ heat capacity. The losses to evaporation can be expected only with a total
heat influx exceeding 112 kjoules/kg.
At the present time the simplest and most economical method of obtaining
this slush is vacuum pumping of the liquid hydrogen vapor with alternate �
processes of freezing-thawing, insuring periodic breaking of the solid _
crust on the sur~ace o~ the liqui.d hydrogen.
A characteristic f eature o~ hydrogen slush obtai.ned in this way is the
' process of "aging" it. Initially the loose particles of unde�ined shape
that were obtained become spheroidal with time, forming a dense mass in
- the pzecipitated slush. Un storage of the slush, in order to eJ.iminate _
thermal strati�ication and local compacting it is necessary to mix it.
In order to maintain a uniform homogeneous mixture in the tank, the con-
centration of the solid phase must not exceed 60%. In order to free the
- slush it is r.ssible to use both pumping and forced means of feed.
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XfY XfY
K
I ~
_ ~
_ .J _ -
a ~ b
E . XfY
- _ - BA BA - K �
~ nx,a _ _ ~ -
K ~ 3
T i
~ ~ .
c d
~ti .
~
BA -
I
Figure 5.7. Cooling systems for cryogenic
_ _ J _ _ liquids in storage:
_ _ n~ a-- with the application of a gas cooler with
- - - heat exchanger ~submerged in the cooled liquid;
_ b-- with the application of a gas refrigerator
~ - by the "evaporator-condenser" system; c-- with
the application of a cooling agent cooled using
a vacuum unit; d-- with ~oint application of
e. the vacuum unit and the gas refr~geration unit
which condenses the vapor af ter the vacuum unit;
e-- with the application of a vacuum unit with loss of cryogenic liquid
as a result o~ discharge of the vapor to the atmosphere; XI'Y gas
refrigeration unit; XA coolant; BA vacuum unit; E-- vapor gathering
tank; 3-- gate valve; n~ liquid losses; IIXA coolant losses;
K condenser.
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. ~ -r
.
XfY BA
I
I
_I_
~ ~ _ nxA
- - XA
~ Q . b~
. , ~ ~
_ ~
nxA - ~ A ~ nxa = xa .
X
_ K - - 3
i B
4
-
- - xa
~ d
~ Figure 5.8. Cooling diagrams of cryogenic
BA liquids when they are flowing in a pipeline:
~ a-- the heat exchanger of the gas refrigera-
- ~ tion unit; b-- heat exchanger with coolant
_ 1__ cooled by evacuation; c-- heat exchanger with
�i - the application of coolant and condensation
- - - XA of its vapors using another coolant with
+ - lower temperature; d-- heat exchanger with
the application of inert gas as the inter-
~ e mediate hea.t-transfer agent cooled by the
3 coolant; e-- heat exchanger with coolant
(liquid picked up from the basic flow) cooled
by evacuation; XTY gas refrigeration unit; K-- condenser; XA coolant;
BA vacuum unit; B-- ventilator fan; 3-- gate valve; IIXA coolant -
losses.
The simplest system of supplying the hydrogen slush to the rocket tank is
the circulation loop through which the hydrogen slush goes from the ground
storage to the tank to a defined level; then the feed is continued with
- simultaneous discharge o~ the cooled hydrogen to the ground f illing system
through a drain hole having a special device for protection from the solid _
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particles. Thus, in the tank it is possible to obtain the required concen-
tration of the solid phase. -
The control fill-cutoff fittings (gate valves, valves) are designed to
control, regulate and stop the feed of cryogenic liquid during the techno~.og- -
ical process operations of storing and filling the tanks in the space rocket
systems, and, as a rule, it has manual, or pneumatic drives and signals of
the required positions of the shutoff (plates). Two separate cavities of
the pneumatic drive provide for fixed extreme positions of the valve pla*_e
as a result of feeding compressed gas to one of the cavities and the
- absence of it i~x the other. In addition, the structural design of the
individual valves provides for one of two extreme positions of the plate
without the presence of compressed gas as a result of the effect of a
spring. On storage between losses, the seal of the inside cavities of the
system is insured by the manual seal valves. In some cases the pn_eumatic
drive of the valve can have a manual, worm transmission by means of which
it is sealed without feeding compressed gas.
Before the beginning of filling, compressed gas is fed to the valve, and
the necessary manual valves are open, as a result of which the hydraulic
f itting of the system is initialized in which the inside cavities of the
lines are protected from the incidence of atmospheric air, and the pressure
in them is increased from evaporation.
The pneumatic valves through which the high liquid flow xate takes place
have special devices providing for slow variation of the flow rate (from
the maximum value to total cutoff) to avoid hydraulic hau~er.
- For this purpose the cutoff f itting of the fill and drain lines of the
system is placed in such a way that in the basic flow line there were no
large blind taps in which the vapor phase can form during movement of the
cryogenic liquid through the ma.in line. The presence of such taps can
lead to unexpect~d hammer phenomena in the response of the f ittings
installed in the blind taps and in the main line. For example, when open-
_ ing the valve of a blind tap (Fig 5.9, a) the vapor accumulated in the
blind tap flows out quickly, and the liquid from the basic line, accelerating
and approaching the still incompletely open valve, is braked sharply, as
a.result of which ~hydraulic hammer takes place. This type of hydraulic
hammer (Fig 5.9, b) can be formed in a blind tap and when the valve is
closed on the main line. In this case, in the basic line when braking the
~low in front of the valve, the pressure rises, the liquid begins to fill
the blind tap, as a result of which the vapor phase in it can be condensed
airl the flow speeded up.
In order to avoid this phenomenon in the blind taps small drain holes are
used to exclude the accumulation o~ vapor, and a carefu~..analysis in
calculation of the response sequence o� the valves is carried out during
the technological operations connected with the feed or discharge of the
cryogenic liquid. Underestimation of this situation can lead to serious
= consequences. Thus, during the process of test filling of a mockup of
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the "Saturn-V" booster rocket during blowing of the tank from the gasifier,
unexpected closure (2 minutes after opening) of the valve occurred on the
section line of the fill pumps, as a result of which hydraulic hammer
- occurred, the magnitude of which (26 MPa) exceeded the margin of strength
of the corrugated pipeline, which led to its rupture and the loss of
2765 m3 of liquid oxygen.
- ~ ,
_ - - -
. a ~
- ~
~ .
- - ~
~ _ - - =I
. b ~
_ Figure 5.9. Schematic of sections of cryogenic pipelines
with blind taps:
a-- when the valve of the blind tap is open; b-- when the _
valve on the main line is closed.
The fittings for cryogenic liquids usually are made according to the follow-
ing scheme: a housing connected with the line using split or unsplit
couplings the shutoff with drive (spindle group) having a split
connection with the housing, which makes it possible to replace it in case
of using up the resarves or failure without dismantling of the line. In
order to decrease the heat influx to the cryogenic liquid the housing has
thermal insulation in the form of the outer powdering with the thermal
insulating material or double walls, the cavity between which is filled
with powder or multilayer insulation and is connected to the vacuum cavity
pipelines. In order to reduce the thermal influxes from the direction
of the spindle group, the cuto~f unit is counected to. the~ p'ush rod of the
drive through a heat insulating bridge, which makes it possible to place~
the drive in the "warm" zone and essentially simplifies its structure and `
servicing.
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High requirements with respect to seal are imposed on the f ittings of the
cryogenic fill system. They are insured by creating the required specif ic
pressure on the seal of the "seat-slide" lock in the application of
sealants of the type of polyfluoroethylene-steel, steel-steel, arul so on.
The means of removal and discharge (drainage) of the cryogenic liquid and
its vapors in the cryogenic fill system consists of drain lines designed to
remove the liquid vapors from the cushions of the tanks and the booster
rocket tanks and for removal of the liquid and its vapors from the drainage
- valves and the valves on the f ill lines and also from the fill columns of
the storage.
The ends of the drain lines are taken out to one location insofar as possible,
located at a safe distance from the launch complex systems. The vapor
- is discharged to the atmosphere here, and the liquid is drained into
special tanks or trenches.
The drainage of liquid and gaseous nitrogen is the simplest and does not
require that special measures be taken. The drainage of the oxygen vapors
as a result of explosion and fire hazard of the mixtures with organic
materials is realized through special fittings which decrease the gaseous
oxygen content to saf e concentrations in the boundaries cf the drainage
area. Liquid oxygen is collected in a special drain tank.
The drainge of the liquid and gaseous hydrogen is most complicated, for the
existing safety engineering rules do not permit the creation of fire-
hazardous concentrations on ground level near the exits from the drainage
- systems. For low flow rate hydrogen is discharged without retardation
by an inert gas or after burning; with large flow rates, with retardation
or with afterburning using a special ignition plane.
Hydrogen is afterburned by the method of combustion through a hydro-
seal, the drain line of which is protected by water from air getting into
it or directly through the drain line and the afterburner. Devices are
installed on all of the drain lines which exclude advancement of the flame
front to the drain line~ Usually before the beginning and after discharge
of gaseous hydrogen, the drain line is purged with 10-15-fold volume of gaseous
hydrogen or helium.
When transporting liquid hydrogen, the drainage regime is selected consider-
ing the minimum speed in the drainage line fitting in which turbulent mixing
~ of gaseous hydrogens with the surrounding air takes place and obtaining
the flow rate of the gaseous hydrogen at the exit which does not require
afterburning or retardation.
The hydrogen vapor is ejected through a special, so-called saf e drainage -
device designed to obtain hydrogen-air mixture with the hydrogen content
excluding inflammation.
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5
6
4 .
- 3 ~ '
_ - Z~
~
�1
IO 9 8
- Figure 5.10. Drainage system of a railroad tanker:
1-- liquid hydrogen tank; 2-- surx~ounding air moved by a
fan; 3-- coil heater; 4-- flow rate regulator; 5-- regulator; -
6-- regulating valve; 7-- incombustible hydrogen-air mixture;
_ 8-- connecting line for discharge of the gaseous hydrogen;
9-- fan; 10 gas turbine.
The diagram of a drainage system (Fig 5.10) developed in the United States
for a railroad tanker that carries liquid hydrogen provides for an automatic
mode of maintaining excess pressure in the gas cushion of the tank. On
achievement of a defined pressure in the cushion, the valve automatically
opens and the gas goes through the coil heater and pressure regulator to
the gas turbine which turns the air fan. On leaving the turbine the hydrogen _
is dispersed in front of tne fan; mixing with the air, which leads ~o a
hydrogen-air mixture in which the hydrogen content is appreciably below
the inflammation limit.
- The evacuation means are used to create and maintain a vacuum within the
required limits in order to insure eff icient operation of the thermal
insulation in the tanks and lines during operation.
`3efore putting the cryogenic liquid in the tank, provision is made for
~:vacuation uf its thermal insulating cavity to a residual pressure of
1.3 Pa through the vacuum lock with the pipeline to the evacuation column.
T.he residual pressure during preliminary evacuation is controlled by the
thermocouple tubes with exit to secondary instruments and through the
electrocontact vacuum meters. After putting the cryogenic liquid in the
tank, the vacuum in the therma3. insulating cavities is brought to the
required value, and it is maintained during the operating process by adsorp-
tion pumps.
Adsorbents are materials capable of adsorbing residual gases by their
sur~ace. The adsorption process is exothermal; therefore the amount of
absorbed gas increases with a reduction in the adsorbent temperature. In
the adsorption pumps the adsorbent is in the zone with lowest temperature,
which provides for maintenance of the gi~ven vacuwn with admissible leak- _
age of gases into the thermal insulating cavity over a long period of time.
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For exa~ple, some industrial tanks with activated charcoal are capable of
maintaining the required vacuum for several years. ~~ith a decrease in
the absorption capacity as a result of saturation with gases the adsorbe~t
is heated to high temperatures (the absorption capacity is restored).
Inmodern tanks, a multilayer vacuum shield and vacuum powder insulation _
are used as thermal insulation.
The multilayer vacuum shield insulation is a set of ser ies-arranged defiect-
ing shields with minimum degree of blackness thermally insulated from
each other by separating inserts. The reflecting shields limit a large
part of the heat influx as a result of radiation, and the separating
inserts decrease *_he thermal conductivity between adjacent shields. The
effectiveness of this insulation is determined primarily by the material
of the reflecting shield, the inserts, the amount of pressure in the
thermal insulation space, the process used to manufacture it and install -
it in the tank. The shields usually are made of aluminum foil several
microns thick or from aluminized (on one or both sides) polymer f ilm, and
_ the inserts are made of various fiberglass materials (glass paper, glass
wall, glass voile, and so on). The multilayer vacuum shield insulation
requires a deep vacuum (to 0.01 Pa), for with a decrease in it, the
coefficient of thermal conductivity increases sharply (by 200 to 300 times
with an increase in pressure to 133.3 Pa). -
The powder vac~ium insulation is an evacuated (to 13.3-1.3 Pa) space f illed
with powdered material (aerogel, perlite, and so on) with low coefficient -
of thermal conductivity. A further increase in the vacuum has no signif i-
cant effect on the magnitude of the thertoal influx which is determined
_ not by the thermal conductivity of the residual gases but by the thermal
radiation and thermal conductivity of: the powder material. The thermal ~
emission through the vacuum powder i~isulation decreases also with addition
of inetal and nonmetal powder to it playing the role of shields. However,
such additives (bronze, aluminum), along with a decrease in thermal radia-
tion, cause growth of the thermal conductivity; therefore the concentrar_ion
of the added powder must insure minimum coefficient of thermal conductivity.
The multilayer vacuum-shield insulation is more eff ective than the powder
vacuum insulation. It has a coefficient of provisional thermal conductivity
(including all the components o~ the influx through the insulation) approx-
imately 10 times lower than in the powder-vacuum insulation, and it is
_ basically used in transport tanks for liquid hydrogen and helium and also
in tanks in which the application of the powder vacuum insulation does not
insure the given requirements with respect to evaporability.
In addition, this insulation makes it possible to have a thickness of the
vacuum space much less than for the powder vacuum insulation, for a
large number of shields can be put in the limited space.
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In practical developments, the combination of these two types of insulation
find application (insulation shields placed in the vacuum space filled with
powdered insulating material).
Safety Measures
It must be remembered that spills of cryogenic liquid are dangerous from
the biological point of view. This is explained by low temperature, easy
evaporability and large concentrations of the vapor formed in the facility.
Entrance into the facility and the structures and remaining in them are -
permitted only with an admissible vapor concentration.
It is necessary to work with cryogenic liquids only in protective clothing
and glasses, avoiding incidence of them on.exposed parts o~ the body, for
this can lead to burns and tissue death..
Oxygen is fire-hazardous; therefore a small spark occurring as a result
of the accumulation of static electricity or on impact can cause ignition
of the gaseous oxygen-saturated clothing and other materials.
When designing and installing the ventilation units at the sensors of the
gas analysis system it is taken into account that the gaseous oxygen is
heavier than air, as a result of which it can fill low places in the
facilities and structsres.
Hydrogen is the lightest element; its vapors are appreciably ligher than
air, and the energy impulse required for ignition is low. This increases
- the possibility of its ignition even for a relatively short time of
dangerous concentration in an open area. A spark or flame in an open
- space causes ignition of a hydrogen-air mixture. The high rate of combus-
tion of hydrogen and the short length of the flame extinguishing section
complicate f ighting such a fire. During the combustion of hydrogen in a
close~ :acility the increase in pressure can lead to an explosion although ,
usually almost ideal mixing of the gases and the presence of an explosion
shock wave are required for an explosion. In addition, the presence of
_ various impurities, especially oxygen in liquid hydrogen is dangerous. ;
The low boiling point of liquid hydrogen and the extremely low solubility
of oxygen in it can lead to the accumulation of oxygen particles and air
in the storage tank, the mixture of which with hydrogen explodes. With
an insignificant oxygen content the mixture is not explosive.
For operating safety with liquid hydrogen, various measures are taken '
which exclude the probability of ignition and the formation of fire-
hazardous and explosive mixtures: the grounding of the tanks and lines,
preventing accumulation of static electricity, protection..against
atmospheric electricity, the application of electric power equipment in
the explosion and fire safe execution or transfer of it to the safe zone; ~
170. ~
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nieasures with respect to fire and explosion prevention consisting in
constant remote gas analysis by low-inertia sensors and the taking of the
required measures with respect to the analysis results (curtailment of
filling, the feed of an inert medium, and so on) using high-speed remotely
controlled equipment.
The formation of dangerous concentrations of hydrogen in the atmosphere
prevent ventilation of possible leaks, spills of liquid hydrogen are pre-
vented by the application of welds, and the formation of f ire-hazardous
mixtures, by maintaining excess pressure inside the system excluding the
incidence of ai.r; by cleaning out the lines and tanks to remove oxygen,
air and other impurities to the admissible amounts before filling them
with liquid hydrogen and installing fine-purification f ilters on the fill
lines of the tanks; periodic cleaning of the tanks to remove accumulated
impurities by draining and heating the tanks with analysis of the residual
gases to determine the accumulated impurities. ~
- Filling the�"Saturn-V" Booster Rocket with Cryogenic Fuel Components
For example, let us consider the procedure for filling the "Saturn-V"
booster rocket. The rocket tanks are f illed first with liquid oxygen,
then liquid hydrogen. The total f ill time is 4.5 hours in this case.
In order to f ill the tanks of the S-I, S-II and S-IV stages with liquid
oxygen, a f ill system is used which includes the following:
A spherical tank with perlite insulation with a volume of about 3400 m3
calculated for inside pressure to 82 kPa consisting of an inside vessel
of alloyed steel of the austenitic class 21 meters in diameter, an outside
vessel of ordinary, unalloyed steel 22.8 meters in diameter; suspension
of the inside vessel executed using vertical and horizontal rods; pipelines
exiting from the inside vessel and located in the thermal insulating space
of the tank along the length in such a way as to insure minimum thermal
influxes and to protect the cutoff fittings from low temperatures during
storage;
Tne heat exchanger for gasification of the liquid oxygen insuring a pressure
in the tank cushion required for normal operation of the centrifugal
pumps; .
A pumping stata.on for feedi,ng oxygen to the tank with two centrifugal
pumps (one reserve) with high output capacity (38 m3/min with a pumping
pressure of 3.72 mPa); the pump drives, through an electromagnetic coupling
which dur~.ng the ~illing process insures adjustment of the flow rate from
9.5 to 38 m3/min for high output pumps and from 0.57 to 3.8 m3/min for -
low-output pumps;
The fill lines under the distribution modules of the valves for each tank
(the fill lines for lar.ge ~low rates of 0.35 m in diameter has no
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insulation; the fill line for small flow rates of 0.152 meters in. diameter
has vacuum insulation);
- The drain line for removal of oxygen to the draina.ge area.
The filling process consists in filling the tank f irst with a low flow
- rate (approximately to 5 to 7% volume), then on the large flow rate
(to 90-96%) with subsequent transition to small flow rate; and on comple-
tion of batching, to the makeup rate. The tanks of the S-l:, S-II, S-IV
- stages are filled to 1300 m3, 320 m3 and 80 m3 of liquid oxygen respectively.
To fill the tanks of the S-II and S-IV stages with liquid hyd.rogen, a _
filling system aahich includes the following is used:
A spherical tank with powder vacuum (perlite) insulation about 3230 m3 in
volume designed for an internal pressure to 0.61 mPa consisting of the
inside tank made of high-alloy steel 18.5 meters in diameter and the
~ outside tank made of ordinary unalloyed steel 21 meters in diameter;
The heat exchanger located near the tank, for gasif ication of liquid hydrogen
- to create the blowing pressure with the forced method of feeding hydrogen
- to the tank; gasification takes place as a result of heating of the surround-
ing air;
The fill lines with vacuum thermal insulation and the distribution blocks
of valves for each tank; -
Drainage lines for removal of hydrogen to the combustion area;
A high-pressure tank receiver for gaseous hydrogen with portable gasifica-
tion unit (the gaseous hydrogen is required to burn the hydrogen coming
from the drain lines of the system).
Before f illing, all the inside volumes for gaseous and liquid hydrogen are
purged successively by nitrogen and helium to remove oxygen. _
The process used to fill the tank consists in cooling the tank with a
small quantity of hydrogen, filling the tank with low rate (to S% of the
volume), then at high rates (to 95%) with subsequent transition to low
rate and on completion of batching, to makeup rate. The tanks of the
S-II and S-IV stages are filled with 1000 m3 and 280 m3 of liquid hydrogen
respectively.
Al1 the filling operations are performed automatically using special con-
trol systems which insure the requixed sequence o� operations both for
normal operating conditions and for various deviations and failures.
This control is duplicated by the remote and automatic control panels for
each control element. During t:~e �illing process, television viewing
and monitoring of the responsible assemblies and units are constantly
carried out.
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5.4. Systems for Filling with High-Boiling Fuel Components
The high-boiling fuel components are used in rocket engineering for thn_
basic and auxiliary engines of the booster rockets and space vehicles. _
Some of the data on the application of high-boiling components are presented
in Table 5.3, and their physical-chemical properties, in Table 5.4.
Table 5.3
Some Data on the Application of High-Boiling Fuel Components
in Space Rocket Systems
Fuel Combustible
Rocket (stage) Oxidant component
"Kosmos" (first stage) Nitric acid Kerosene
"Kosmos" (second stage) Liquid oxygen Asymometric
dimethylhydrazine
"Titan-II", "Titan-III" Nitrogen tetroxide Aerozin-50
"Agena" (stage) Nitric acid Asymmetric di-
methylhydrazine
"Apollo" spacecraft:
Takeoff and landing engines; Nitrogen tetroxide Aerozin-50
Orientation engine The same Monomethyl-
hydrazine
"Surveyor" spacecraft (the
steering engine) The same Aerozin-50
"Centaur" (stage); auxiliary Hydrogen peroxide
engine -
Nitr ic acid HN03 is a high-boiling oxidant with high density. It is
explosion-safe. 100% nitric acid is a colorless liquid with sharp odor,
it is hygroscopic, unstable and tcxic. It decomposes easily i:ito water,
free oxygen and nitrogen ~xides (the lat*_er color it from yell~~w *_o brown).
Additives of nitrogen tetroxide or water are used as stabilize:s. The
nitric acid vapors are harmful to the health (if they get on the skin they
cause diseased, slow-healing ulcers).
Nitric acid and its vapor hava high corrosion activity with respect Lo the
majority of materials, for the reduction of which, the so-called inhibitors
- are introduced into the nitric acid. For storage, stainless chrome and
- chrome-nickel high-alloy steel, the ma~ority o~ aluminum alloys and non-
metallic materials (polyfluoroethylene, asbestos) are used for storage. -
The effectiveness of nitric acid, as an oxidant, increases significantly
on solution of nitrogen oxides.
Nitrogen tetroxide, N2Q4, is a high boiling oxidant. It is more eff ective
than nitric acid. Nitrogen tetroxide is explosion-safe, stable and less _
aggressive than nitric acid, and it is toxic. Insur ing great,~r specific
thrust (by 5%) than nitric acid, it has a narrower range of maintenance -
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Table 5.4
Basic Physical-Chemical Properties of Some High-Boiling
Fuel Components
Bo iling Melting
Chemical Density, point Tk, point T~elt~
Fuel formula k~/m3 �C �C
Nitric acid HN03 1510 86 -44
Nitrogen tetroxide N204 1450 21 -11
Kerosene C1pH2O (provisional) 800 200-250 -49
Hydrazine N2H4 1010 113 + 1.5
Asymmetric dimethyl- H2N-N(CH3)2 790 63~ -57
hydrazine
Monomethylhydrazine H2N-NH(CH3) 875 87.6 -52.4
Aerozin-50 Mixture of 50% 900 70 - 7.3
asymmetric dimethyl-
hydrazine and 50%
hydrazine
Iiydrogen peroxide H2O2 1450 150 - 1
(100%)
a~ liquid state, which is increased by dissolving other nitrogen oxides
in it. For example, the introduction of nitrogen monoxide lowers the
freezing point by approximately 28�.
Kerosene C1~H2O is a high-boiling fuel, colorless or yellow liquid. It
is a mixtur.e of hydrocarbons obtained for distillation of petroleum within
defined temperature or cracking limits; out of all of the fuels it is the
most dangerous, simplest and most convenient in operation. It is chemically
stable even at high telnperatures, it has low corrosion activity with respect
to metal and has low toxicity. Kerosene is cheap and available. It is
widespread in engineering, and its production has a broad raw material
base. Kerosene has inconstancy of chemical composition, depending on the
origin of the petroleum. This deficiency is eliminated by creating
artif icial hydrocarbons which have the same characteristics as kerosene
' but have defined chemical composition and constant chemical properties.
Kerosene is inert with respect to construction metals, but admixtures with
water, sulf ur compounds and organic acids increase its corrosiveness.
Hydrazine N2H4 is a high-boiling combustible and a single-component fuel
for liquid-propellant rocket engines. It has the highest density among
the fuels used, and is colorless, it smokes in the air, it is capable of
thermal or catalytic decomposition wi.th the formation of a hot gaseous
mixture of hydrogen, nitrogen and ammonia; it is hygroscopic, it easily
picks up atmospheric moisture; it is toxic, it has an irritating effect
on the mucous membrane of the eyes; on superheating in a closed space or
under the effect of a powerful pulse it is subject to explosive decomposi-
tion; it has high hardening point which complicates use of it. With
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respect to metals it has low corrosion activity, but in the presence of
oxygen it acts on copper and its alloys. The copper ions catalyze the _
decomposit:Ion. In rocket engineering, it is used as a fuel component witii
asymmetric dimethylhydrazine (aerozine-SO), ammonia, a~d so on.
Tl~e .lsymmetric dimethylhydrazine (ADMH) H2N-N(CH3)2 is a high-boiling ~
fuel, a derivative of liydrazine obtzined by replacement of the hydrogen
- atoms by hydrocarbon groups; a colorless liquid with ammonia odor, it is
hygroscopic and toxic; as a fuel it is less eff ective than hydrazine (in.
its molecule in addition to hydrogen atoms it contains less effective
carbons), it is more convenient in operation, f or it retains its liquid
state in a large temperature range; in the presence of ~aater it. is corrosion -
, active with respect to aluminum and its alloys; it is easily oxidized
by oxygen of the air.
Asymmetric dimethylhydrazine is thermally stable, but with an increase in
temperature it decomposes with the release of heat and the formation of
hot gaseous products; it is less explosi~~e than hydrazine, but on super--
heating in a closed space it explodes. It is superior to nitric a::id
with respect to toxicity.
In missile engineering it is used as a basic fuel, a component part of the
combustible (aerozin-50) and as a single-component fuel for the turning
of the pump turbines of the engir.zs.
Monomethylhdrazine H2N-NH(CH3) is a high-boili.ng fuel, a de:ivative of
hydrazine, colorless liquid which fumes in the ~.ir with ammonia odor,
and it is toxic; with respect to its properties, including corrosiveness
it is similar to asymmetric dimethylhydrazine. With respect to eff ective-
ness and ~tability it occupies an intermediate position between hydrazine -
and asymmetric dimethylhydrazine.
Hydrogen peroxide H2O2 is a high-boiling oxidizing agent and a single-
component fuel, a colorless liquid, it is toxic, explosion and fire hazardous
(organic materials are easily burned on contact with it), when it gets un
human skin it causes serious burns and is unstable. Metals (copper, nickel,
silver, the products of iron corrosion, and so on) and chemical manganese e
compounds are catalysts, on contact with which hydrogen peroxide decom- -
poses stormily into water and atomic oxygen with the release of a lar.ge
, quantity of heat. Water evaporates, and the mixture ohtained (vap~rizing
gas) is heated to 520�C with 80% concentration and to 1000�C with 99%
- concentration, The equipment for the hydrogen peroxide 's carefully -
cleaned (degreased, f lushed with distilled water, and so on), and it is
_ passivated.
~ Such properties of the components as high corrosiveness, toxicity,
inclination to decomposition, fire and explosion hazard, self-ignition
of some of the fuel vap~r, high requirements with respect to batching
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accuracy when filling require the corresponding schematic and structural _
solutions when designing the filling systems. Therefore, the schematic
diagram of the filling systems using high-boiling components has such _
snecif ic equipment as the batchers for insuring high accuracy of filling
(for f illing the space vehicles), means of neutralizing the toxic components
and their vapors, thermostating systems to provide for filling the tanks
with components of a def ined temperature.
The high-boiling components stored usually are made up of several tanks
(for especially aggressive liquids a reser.ve tank is provided) and troughs
- for removal of the spilled liquid to a neutralization system.
The reserve of the stored component is selected beginnig with the quantity
consumed for one or several f illings, to fill the lines connecting the
storage and the rocket, for guaranteed remains in the tank and so on.
The tank for high-boiling components is a cylindrical reservoir with semi-
elliptic ends and single walls, the composition of which includes safety
valves operating both on excess pressLre and on rarefaction; the devices
for *_aking samples; the means of monitoring and n~easuring levels, tempera-
ture and pressure both with direct and with remote monitoring; the fill `
lines, drain lines and blowing lines and devices for removing air (deaera- -
tion) from the components.
In some cases (when thermostating the liquid) the tanks have thermal
insulation on the outside made of incombustible insulation material (asbes-
tos, slag cotton, glass cotton, and so on) protected by a housing or hood.
When storing the components having high corrosiveness and hygroscopicity,
measures are used to purify them of possible mechanical impurities and
water.
The liquids oxidized in the air are stored at excess pressure of the
natural vapor or inert gas (nitrogen, helium, and so on).
A mandatory condition of the stable storage of hydrogen peroxide is the
- presence of defined thermal conditions and finish of the ins.~de surfaces =
of the vessels for which the equipment is passivated.
BeFore filling and during the storage process many of the fuel companents
are therm~stated (cooled or heated), which arises from the calculated
density of the component in the tanks or the condition of its stable
storage. As a rule, the fuel components are cooled in the summer and
heated up in the winter.
, The required storage temperat~ire is maintained by thermostating the system
having means of remote monitoring and control. In the launch complex the
~ so urce of cold (heat) is the only refrigeration center; at the f illing
station it is a speci.al thermostating system including freon refrigerators
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and water heating elements. The heat transfer agents are nonfreezing
solutions of salts or antifreezes. The fuel components are thermostated
_ by circulation using a pump in the "tank-heat exchanger" loop.
Before filling, in order to insure reliable starting of the engine, some
of the fuel components are sub~ected to deaeration (removal of air) by
evacuation of the space above the liquid surface or bubbling (passing
an inert gas - nitrogen or helium - through the liquid).
The means of supplying the high-boiling fuel to ;he tanks of the booster
rockets and space vehicles use both forced and pump methods of feed. The
main lines are laid usually with some slope to insure complete drainage of
the comgonents on completion of the filling into special drainage tanks -
from which after taking the analysis the liquid goes back to the storage
tanks or into the neutralization system. For seal, the joints of the
sections of the aggressive liquid lines are made welded or with a minimum
number of flange connections. All of the fittings have seals.
In order .*_o f ill the booster rocket tanks witti toxic components, two systems
are.used. WitYi respect to one system the vapor formed during filling is
run through the drain line into special units (af�terburners) in wnich they
are burned, and the combustion products go to the atmosphere; in accordance
with another scheme the vapor formed goes through the drain line from the
booster rocket to the gas cushion of the storage tank (the so-called
connection). For toxic, aggxessive, fire-safe liquids, field pumps are
used (Fig 5.11) without shafts that go to the outside in the explosion-
protected execution.
t ~ r ~Z -
1
~
~ ~ ~3
~ .
~ -
~t ` 1 �
` `
I ~ _
, 8 ~ 6. s
Figure 5.11. Sealed electric pumps for toxic, aggressive and
fire-hazardous liquids:
1-- i.mpeller; 2-- rotor; 3- shaft; 4-- rear bearing; 5--
coil; 6-- stator; 7- front bearings; 8-- screen
The tanks of the space rocket system are filled with high-boiling fuel
components both at the filling station and the engineering co*nplex (the
tanks of the space vehicle) and in the launch position f,the t~nks of the
booster rocket and the space vehicle).
177
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Each component has its own filling system at th~ filling station (Fig 5.12),
- which includes the basic storage, batcher, f illing columns and liquid,
i gas and vacuum pipelines which co~ect them into a united pneumohydraulic
system. The pipelines are connected to the fill necks of the tanks using
filling connections serving also to drain the component, evacuate the lines
- in the tanks, for spilling of the componer_t before filling, drainage and
purging of the f ill lines before disconnecting, supplying the required
reagents to decontaminate the toxic components. For safety, the fill
connections for the fuel and the oxidizing agent are located on opposite
sides, alongside the fill columns. -
The fillin~ process consists of coupling the lines of the fill systems to
the fill conner_tions of the space vehicle and checking the joint for seal,
thermostating of the component in the given temperature and deaeration
(these operations usually are performed in advance), filling of the batcher,
evacuation of the f ill lines from the batcher to the fill connections and
tanks (when f illing by the drainless system), f illing the fill lines to the
tanks (in order to increase the accuracy of the batching), filling the
tanks from the batcher to the given amount, drainage and purging of the fill
lines to the drain tank with relieving of the batcher of the component,
disconnecting the f ill connections, sealing the fill necks of the tanks and
neutralization of the remains of the components and its vapor.
In the case of internal batching (Fig 5.13) the fill system does not have
a batcher, which determines its theoretical peculiarity and the filling
process.
The neutralization means provide for decontamination of the toxic and
, aggressive liquids and their vapor.
The most widespread are two neutralization methods: physical flushing
and purging the lines based on good solubility of certain liquids and water _
or in organic solvents and chemical - neutralization of the oxidizing _
agents by solutions of alkali, and t':~ combustible fuel components by strong
oxidizing agents (chlorine-containing reagents). When selecting one method
or another the structure of the equipment subject to neutralization is
- taken into account (the tank configuration, the presence of places that are
difficult of access, and so on) and also the corrosion resistance of the ~
materials with respect to neutralizing materials and neutralization products.
The neutralization of toxic and aggressive liquids and vapor is possible
also by burning them in the neutralization chamber (Fig 5.14) with subse-
quent discharg~ of the combusticn products into the atmospheree Such chambers
are made both stationary and portable. The neutralization of the component~
and the vapor by accumulation of them and subsequent e~fect on them by the
corresponding chemical reagents is possible (Fig 5.15). These chemical
reagents decomp~se the component~ and vapor to harmless compounds. The
aqueous solutions of the final reacti~n products are pumped by pumps into
- the evaporation areas. In order tc absorb the vapors of toxic components,
adsorbents are used.
178
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Hydrogen peroxide is neutralized by multiple dilution with water; the -
use of water jet pumps simultaneously pumping out and diluting the hydrogen
peroxide is especially effective.
During operation of the f ill systems, safety measures are carefully observed.
The incidence on the skin of aggressive liquids causes strong slow-healing
burns and ulcers; therefore all operations are performed in special protec--
tive clothing. Before entering into the facility with the fill equipment
usually the gas content of the air is monitored by the gas analysis system,
and if necessary, gas masks are used. _
_ ~ ' Z 3 -
~
m iP 4n ~ I
�w a~~ ~v~ 9
. ~ ~ c ~ '
~ y
~ T
3 i ~ 3
T'aa airA 4 ,
1Ku8KUU ~ ~ ~ xa yea
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caueNOU eMrtocmu ~ (3) -
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x~MnoHeHma S
- Figure 5.14. Neutralization system for the toxic component by
combustion:
1-- combustion chamber with injector; 2-- air fan; 3-- cutoff
valve; 4-- drainage safety valve; 5-- fuel tank; 6-- forechamber
with sparkplug
Key:
1. Liquid componer_* t:om the drain tank
2. Vapor of liquid component
3. Gas for blowing
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Filling the Space Rocket System with High-Boiling Fuel Components
As an example, let us consider the filling of some of the space rocket
systems.
The "Apollo" spacecraft is filled on the launch complex from the service
tower. During the first day of preparation, a tank truck is brought to th~
tower with nitrogen tetroxide; then the tanks of the service module (9.5 m),
the jet control system of the command module (0.23 m3), the takeoff and
` landing stages of the lunar module for the basic and auxiliary devices
(3.8 m3) are filled successively.
On the next day a tank truck with aerozin-50 is brought to the service tower
and the tanks of the service module (8 m3), the tanks of the takeoff
and landing stages (4.5 m3) are filled successively, and the tanks of the -
jet command module system (0.38 m3 of monomethylhydrazine) are filled from
a separate reservoir.
The system for filling the first stage of the "Saturn-IB" booster rocket
with fuel at� the launch complex includes storage, a pump, ground lines
connecting the storage and the booster rocket. The lines to the tank are
connected through the cable mast which is moved to the side before launch.
The storage is made up of a tank, the filtration and pumping system, the
monitoring and measuring system. The tank which holds 215 m3 has safety
valves, lines and f ittings. The filtration system is used for removal of
water and other impurities from the hot watar during the storage process
and feed of the components. The measuring and monitoring system provides
for monitoring the tempexature and pressure. -
After filling the tank from the portable transport means to prevent accumu-
lation of inechanical impurities and water in it the fuel is filtered by
circulation through a separator. filter. The analysis samples are taken
through the quick-removable cover of the tank. -
The booster rocket is filled in several operations. Two days before launch
a somewhat greater amount of fuel than required is put in the tank.
Initially the tank is filled to 15~ of its volume, after which its seal
is checked. Then the tank is fillec? to 98% volume at high flow rate
(7500 ~,/min), and to the full volume at low f low rate (750 R/min), On
the last day, 35 minutes before launch (after �illing the booster rocket
with the oxidants oxygen), the amount of fuel is .finally batched by
the drainage from the tank. The amount of drained fuel is determined by
a computer.
5.5. Gas Sugply Systems
Compressed gases helium, nitrogen and air --are widely used in rocket ~
engineering.
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The compressed gases are used to control the hydraulic equipment of the
filling systems and individual mechanisms of the launch devices; the
filling of the on-board booster rocket tanks; purging of various parts of
the line to create a protective atmosphere in the purged cavities; bubbling _
(mixing) of the fuel components both in the rocket tank and in the storage
tanks; blowing of the tanks before launch and during the drainage process;
provision for the air conditioning systems, fire extinguishing systems,
forced f eed means; the working medium for various ground and on-board
refrigeration units; the creation of the required atmosphere in the rocket
tanks and in the ground storage tanks before filling them and after drain-
age with the application of f ire and explosion-hazardous (with respect to
air) liquids and vapors and also during pneumatic testing of the equipment
of the booster rockets and ground systems (checking the points of connection
of the ground and on-board lines for seal, checking the adjustment of
the reduction gears, and so on).
Such widespread application of compressed gases is explained by the
advantages such as the possibility of supplying from one energy source (the
receiver tanks) of a large number of users, simplicity of the storage of -
the energy by compression to high pressures and convenience nf application
of the electropneumatic devices combining the capacity for ~reating the
required force and the speed of the electrical systems.
The demand for compressed gases is met by the pneumatic systems of the
space center, the specific characteristic~ of which are as follows by
comparison with other f illing systems:
Large number and variety of the performed operations cambined with a
' greater number of users;
Signif icant operating time of the gas supply system (in all stages of
preparation of the space rocket system);
Provision of compressed gases for the concluding, responsible pre-launch
operations (blowing of the tanks, introduction of var~ous ground and on-
board devices, and so on) and the operations performed at the beginning
of movement of the space rocket systems (uncoupling of the split connections
at the beginning of lif toff, the removal of the platforms with uncoupled
fill, drain and other lines);
- A wide range of parameters (to pressure and flow rate) of the gases supplied
to the user.
The cc,:�^Yessed gases used in space rocket complexes have high requirements
imposed on them with respect to their purity with regard to mechanical
admixtures, moisture and oil. The presence of inechanical particles, the
precipitation of ice crystals and oil during choking from wet compressed
air lead to spoiling of the elements of the pneumatic system, loss of
seal of the pneumatic f ittings and failure of them.
184
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The gas supply systems provide compressed gases to the engineering and
launch complexes and are constructed in accordance with the following
scheme: "Compressed gas force-pipelines with distribution, cutoff,
regulating and safety fittings-user" (Fig 5.16).
The application of high pressure gases permits us to have f eed lines wit;~
- relatively small cross section and low metal consumption.
The force of the compressed gases is the central compressor made up of
- autonomous units for each gas an3 receiver and also the liquid nitrogen
storage located alongside.
From the compressor the high pressure gases go through pipeiines to the
receivers of the en~;ineering and launch complexes which also can serve as
portable compressor stations. The gas reserve in the receiver must provide
for the technological process cycle of operations after which the receiver
is refilled.
The compressor station is made up of mobile or stationary multistage
piston type compressors with three-phase asynchronous motors in the stationary
compressor stations and diesel engines in the portable stations. Frequently
the compressor and the diesel engine are combined in a single unit (the
diesel compressor), and their pistons are directly joined to each other.
The diesel compressor has comparatively high efficiency, exceeding by 1.5-2
times the efficiency of the analogous compressor with traditional drive
from a diesel engine.
The sources of gases for the compressors are atmospheric air, nitrogen
coming from the nitrogen extracting unit or from a special liquid nitrogen
storage and helium delivered from the manufacturing plant by special trans-
port units in high pressure tanks.
Atmospheric air and other gases compressed in the compressors contain
mechanical impurities and dust and also water and oil vapor, for removal
of which the drying and cleaning unit is used. The gas moisture is
determined by the "dewpoint" the temperature for which the wr~ter vapor
contained in the gas becomes saturated, and with a further reduction in
temperature supersaturated; in this case the excess moisture falls out in
the form of dew, the time of falluut of which is fixed by a moisture
indicator. _
The moisture is removed From the gas by three methods: cooling of the gas
below the required "dewpoint" by the inertial method in the mnisture and
oil separators (removal of the drop moisture) and absorption of the
moisture by the adsorbents.
In the first method the gas is strongly cooled in the heat exchanger, as a
result of which the maximum amount of moisture which is contained in the
gas before its saturation decreases, and the excess moi:sture falls cut in
the form of snow or frost on the walls.
185
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FOR OFFICIAL USE ONLY
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With the second procedure, the principle of a sharp change in direction of
movement on the gas and loss of speed by it is used, as a result of which
the drops of water under the effect of inertial or centrifugal forces
are discharged to the side surface, they flow off into a speciaY low
settling tank, from which they are removed through the drain line from t:~e
moistur e and oil separator.
The most widespread is the third procedure absorption of moisture by
the adsorbents (alumogel, silicagel, synthetic zeolites). In this case
the drying mo~lules have two adsorbers: one operates, and the other
recovers its absorbing p~operties on purging with hot air. The gas is
purified of inechanical particles by ceramic filters. The purified and
dried gas from the compressor is released to the users or it is accumulat~d
in the receiver tanks.
For large flow rates for gaseous nitrogen and hydrogen, two methods of
obtaining compressed gases are used: with the help of low pressure stationary
gasification units with subsequent release to the user or compression in
compressors and with the help of stationary or portable high pressure
gasification units in which the processes of compression and gasification
are combined.
1 2
3
4 5
S
10' ~ s
9 8 ~
~ Nnompe6umearo ~a~
~
Figure 5.17. Schematic of the gasification unit:
1-- lique�ied gas pu.mp; 2-- liquid feed line �rom the pump to -
the evaporat r; 3-- liquid f eed line from the tank to the pump;
4-- liquefi gas reservoir; 5-- reservoir filling line; 6-- -
evaporator f blowing the reservoir; 7- check valve; 8--
evaporator; 9-- heater; 10 electric bay
Key: ~
- a. to the user
187
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The high pressure gasification units are constructed in accordance with
one schematic diagram (Fig 5.17) wh ich includes the tank for storing the
liquefied gas with the evaporator, creating the required pressure at the
input to the pump serving for compression of the liquefied gas and feeding
it to the evaporator; the evaporator represents a multipass coil put in
heated water. The gas formed in the device heated to 283-303K goes
through the check valve to the receiver tanks, passing through the cleaning
and drying units.
The helium goes to the receiver from the compressor units having good seal-
ing surfaces to prevent its loss (the helium is a very fluid and easily ~
penetrated gas).
The receiver (compressed gas storage) is equipped with means of receiving,
storing an.i f eeding gas to the users. In order to decrease the dimensions
of the receiver, the compressed gases are stored in tanks at a pressure to
41.2 MPa. Each tank usually has two outlets to common manifolds with cut-
off and saf ety f ittings.
The input and output lines of the tanks joined in the sections are led out
to the pneumatic boards with controlled cutoff, regulatable and safety
fictings, f ilters and monitoring and measuring instruments. In order to
determine moisture of the gases and to take samples, the receiver has a
special board. The control of the receiver fitfiings and the pressure
monitoring are possible both manually and remotely.
For convenience of operation the f ittings and the monitoring and measuring
instruments of the individual sections are grouped on individual pneumatic
boards. The elements of the fittings, *he instruments and the tanks are
connected by lines to the lens type connecting devices providing for seal
of the joints even with some misalignment of the pipelines.
The compressed gas is fed from the receiver through the distribution boards
with reducers to the users. For remote control of the output of the gases
from the sections of the receiver and feed of the gases to the specific
users, electropneumatic valves mounted in the pneumatic boards are used.
The primary fittings of the pneumatic systems are the gas reducers, the
safety valves, the electropneumatic valves and the gate valves.
Gas reducers are automatic regulators which step down the pressure to a
given magnitude as a result of choking the gas in the cross section
fnrmed by the valve and its seat. The reducer maintains a given pressure
at the output on variation of the flow through it and the pressure at the
input. The structural design:, of the reducers are varied and depend on _
the requirements imposed on their accuracy, output capacity, dimensions and
weight. The reducers can be spring and unit type, and in turn, the latter
are divided into simple, with control pressure, with hydraulic booster -
and with pneumatic booster. Depending on the direction of the effect of
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~
~ =
BY PROFESSOR R. P. ~OL' SK I Y

? JRNURRY 1980 CFOUO) 3 OF 4
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the incoming gas on the regulating element (valve), the spring reducers
can have direct action valves (the valve is lifted irom the seat under
the effect of the incoming gas) and check-action valves (the valve is
held against the seat by the incoming gas).
A spring type zeducer with check valve (Fig 5.18) operates as follows.
Und er the effect of the incoming gas fed to the cavity, to the unbalanced
area of the valve and the spring, the valve, pressed against the seat,
does not pass the gas to the low pressure cavity. When adjusting (loading)
the reducer for a def ined regime by turning the regulating screw the
spring is compressed, and the valve is opened by the pusher. The gas goes
to the cavity VH, where its pressure on the diaphragm~.equalizes the force
of the spring. In the absence of flow beyond the reducer the gas with a
def ined adjustment pressure again clamps the valve agair~st the seat. On
flow of the gas beyond the reducer the output pressure in the cavity Vg
and the pressure of the diaphragm are diminished, as a result of which the
valve opens, choking the gas passing through the slit between the seat anti
the valve. The pressure in the cavity again rises, znd with a defined
magnitude of a~ljustment between the forces acting on the moving system of
the reducer, dynamic equilibrium is established which corresponds to a
def ined gas flow ratP. With variation in the gas flow rate, a new equilibrium
is estab7ished for a diff erent magnitude of the choking slit.
The reducer with the direct-action valve operates anaZogously.
In order t:o reduce the large gas flow rates, unit reducers are applied _
which are made up of two parts; the power and actuating (adjusting) -
reducers; adjustment is realized by loading the actuating reducer. After
the actuating reducer (the ordinary spr~ng reducer) the gas goes to the
- controlling cavity of the power reducer and ~pens the valve of the latter. ~
The pressure at the output of the power reducer compensates for the force -
from the gas pressure in the controlling cavity, and dynamic equilibrium
is established for a def ined f low rate. With a variation in the gas flow
_ rate the dynamic equilibrium is established for a diff~rent magnitude of
_ the clearance between the seat and the valve in the power reducer. A
simple unit reducer operates by the same principle.
' The unit reducers with hydraulic and pneumatic boosters also operate
analogously, but with grea.ter accuracy.
- The safety valv~s are used to prevent the inside cavities of the tanks and
lines from a possib:Le increase in pressure; with an increase in pressure
, the valve moves away from the seat, and the gas i.s released through the
exit opena.ng to the atmosphere or the drain line; the valve closes with
a reduction in pressure.
The various structural elements of the sa~ety valves classified by the
nature of opening of the valve are divided into proportional, nonpropor-
tional, pulsed and mixed type. In addition, the proportional and
189
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.
FOR OFFICIAL USE 0"rTLY
, ' 3
,
2
4
I
.
VN 5
- ~a) ~ea
hue ' �
Be~raoaoe ~b~
~6 "--'aaeAUrue
/ 6
, 7
Figure 5.18. Sp::ing-type reducer with check valve:
1-- valve; 2-- operating spring; 3-- adjustment screw;
4-- sensitive element; 5-- push rod; 6-- sea.t; 7-- valve
spring �
Key:
a. low pr~ssure
b. high pressure
nonproportional valves can be direct and check action. In the direct action
saf ety valves the operating pressure opens the valve, breaking its seal,
and in the safety check valves it pushes the valve against the seat,
tightening (sealing) this connection. Thz flaps in the proportional ~
safety valves open for a flow rate proportional to the rise in pressure;
in the nonproportional safety valves, discontinuously as a result of the
crea.tion of additional forces.
The widest use is made of nonproportional check valves which, by compari-
son with the remair.ing ones, are better sealed. The nonproportional safety
check valve (Fig 5.19) operates as follows. In the absence o� pressure,
the valve receives only a small force from the spring, which only fixes
the valve in a def ine~i extreme posita.on. The operating pressure of the
safety cavity clam~ps the valve tightly against the seat and moves them in
this position upward, clamping the operating spring. With an increase in -
pressure, the sealing force of connect~.on of the valve and the seat
increases until the pressure of the safety cavi.ty is compared with the
adjustment pressure. At the adjustment pressure, the valve, reaching the
stop, halts, and the seat under the effect of che increasing pressure will
190�
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continue its movemEnt. As a result, the seal of the vaJ~ve-seat connecti~zi
is disturbed, and the gas goes to the cavity A under the additiona.l
area f. The seat moves upward without additional rise in pressure, oper~~ing
the output opening of the cavity A. The pressure in t?ae cavity A is
stabilized, and the seat halts. Witt? a reduction in pres~ure in the safety
cavity under the eff ect of the operating spring the siaat clamps the valv~,
removing it from contact with the stop. The pressure in the cavity A drops,
and the seal of the valve-seat connections is restored.
~
3 ~4 v ~b)
~
4
~ ~ , jl,
~ ~ c~~
A 5
~ f
\
2 6
f
� ~ -
'o ~ -
m~ ~a~
Figure 5.19. Nonproportional saf ety check valve:
1-- housing; 2-- valve; 3-- operating spring; 4-- ad~ustment
screw; 5-- valve spring; 6-- seat
Key:
- a. Input
b. Clearance
c. Output
The pulse safety valves are made up of t_k*o valves: the sm:~.ll cross sec.tion
control valve and the basic valve for pa.~sing the entire flow. The
feedback control valve which responds ~utomatically operates on the basic
valve, opening and closing it. This s~stem insures high ac^_uracy of the _
response for large values o� the flow rates and pressures.
The mixed type sa�ety valves in the initial opening step operate in
accordance with the nonproportional check valve scheme, and on complete
opening, by the proportional check v~ive scheme. -
Gate valves and pneumatic valves (sliutoff fittings) are designed for
_ reliable closure of the pneumatic lines, and when they are open they provide
_ the required �low rates of the gas with minimum pressure losses. Pneumatic
191.
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valves with control pressure and electropneumatic valves (,~neumatic valves
with control from an electromagnet) are distinguished.
The gate valves can have both manual and remote drives (electric engine).
The force on the cutoff of the valve is transmitted through the self-
braking screw couple, and the shutoff of the valve retains the established
position after removal of the control input. The ~ralves are set so that
in the closed position the seal of the push rod will not be under the
eff ect of the working medium.
Electropneumatic valves provide for remotely controlled feed of the gases
to the various users. The characteristic of the electropneumatic valve
~is given in the deenergized state. Dependi.ng on the position of a cutoff,
che electr~pneumatic valve can be closed (deenergized-closed) and open
(deenergized-open) and also they can be made with and without drainage.
'~he electromagnetic valve without drainage simply stops feeding gas to the
user when it is closed, and the electromagnetic valve with drainage not
_ only stops feeding the gas, but discharges the gas remaining between the
- ~iser and the e?ectropneumatic valve to tre atmosphere. The electropneumatic
valve with drainagF is used primarily to control thE pneumatic valves,
electropneumatic valves without drainage for various types of purging, _
blcwing and opening of pneumatic locks.
= When developing the layout of the pneumatic system and its operating condi-
tions, usually electropneuma.tic valves are selected, the electromagnets
of which will be deenergized for the maximum amount of time. However, in
practice, cases of prolonged (to several days) staying of the electromagnet
of an electromagnetic valve under current are possible.
- The direct-action electropneumatic valve and with pneumatic booster are
distinguished. In the direct-action electropneumatic valve, the electro-
magnet shifts the push rod of the basic valve directly covering the gas
cavities; in the electropneumatic valve with pneumatic booster, the push
rod of the servovalve (the unloading valve) having smaller working areas
than the basic valve. The displacement of the servovalve causes redistribu-
tion of the for.:es acting on the sensitive elements of the electropneumatic
- valve ;~hich leads to a displacement of the basic valve, the e?ectro-
pneumatic valve with pneumatic booster is used for ~arge cross sections
of the pipelines. The double-action electropneumatic valves used to control
double-action pneumatic valves have one input and two puts for the gas.
Gas is always used in one of the output cavities. The displacement of the
electromagnet c,auses gas feed to the other output cavity and drainage of
the ~as from the filled cavity. The application of one double-action
electromagnetic valve replaces the use of two electromagnetic valves with
drainage. This decreases the number of fittings in the system and increases
its reliability.
The ele~tropneumatic valves usually are made up of an electromagnet, a
housing with seats for the shutoff assemblies and connections for ~oining
to the pipelines, a shutoff containing a push rod with a valve or system
192
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!
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of valves and special devices the mechanism for manual inclusion (if
it occurs), signalling the position of the valve, the fixing device (if
it exists), and so on.
~ The pneumatic valves with controlling pressure are used for remote closure
of the lines with high compressed gas flow rate usually paired with the
electromagneti~ valves, which makes it possible to operate by the following
scheme: "an electric signal from the controY panel electromagnetic
- valve f eeding of the control air from ~the elect~omagnetic valve
response of the pneumatic valve (closing or opening the line)."
With respect to structural designs of the pneumatic drive the pneumatic
. valves are divided into simple and double action valves. The simple action
pneunatic valves have one controlling pressure cavity, after discharge of
which the valve shutoff is shifted to the other position by the pressure
of the medium and an elastic element (spring). These pneumatic valves
ha.ve one fixed position without feeding a controlling pressure. The
position where the controlling pressure is not fed to the valve is con-
sidered to be normal. The double-action pneumatic valves have two con-
trolling pressure cavities. The position of the shutoff of the pneumatic
valve is 3etermined by the fact that a controlling gas is fed to each =
cavity. With~ut the controlling pressure and without the clamping springs
the double action valve does not have a fixed position. The position is
considered normal when the controlling pressure is .f ed, and the electro-
pneumatic valves are deenergized. These pneumatic valves are controlled
either by two electropneumatic valves (one deenergized open, the other
deenergized closed) or by one double-action electropneumatic valve.
It must be noted that the speed of the pneumatic valves can be altered in
the 1�equired direction by using replaceable in~ectors installed on the
contrcilling pressure line at the input to the pneumatic drive.
Gas Supply System of the "Saturn-V" Booster Rocket
The gas supply system provides for the production, storage and distribution =
of the compressed gases nitrogen and helium. The low pressure compressed
air is used in the conditioning systems.
The high pressure gaseous nitrogen and helium are obtained using converter-
compressor equipment which includes the liquid nitrogen storage (a spherical
tank with perlite nonvacuum insulation); the high and low pressure liqu~d
nitrogen pumps and gasifier; the filtering and drying unit; the helium
high-pressure compressors a.nd distribution boards,
The gaseous nitrogen is obtained by gasification of liquid nitrogen which _
is passed through the deep cleaning filters to the high and low pressure _
- pumps; then it goes to the gasif iers and after them to the f iltering _
and drying units where already in the g,aseous �orm it is purified of vapor
and hydrocarbons and on passing throug~h the fine clea~ing filters, it is
fed to the distribution boards under a pressure of 41.2 and 0.98 MPa.
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The gaseous helium arrives in the transport units under a pressure of
15.2 I~a. The helium is fed through the. fine purifying filters at signifi-
cantly lower pressure into the compressor units having oil and water traps,
after which at a pressure of 41.2 MPa it is fed through the cZeaner to the
distribution unit.
From the distribution devices the high pressure gaseous helium and nitrogen
are fed to the compressed gas storage of the vertical assembly building
and to the launch complex storage. 7'he storages include the receivers
from nitrogen and helium made up of several tens of tanks. The collectors
and the tanks of the receiver are equipped with safet~ valves and rupturable
diaphragms for the case of an emergency rise in pressure. ~
The amount of compressed (to a pressure of 2/~.5 MPa) nitrogen an3 helium
is insured by performanc of all of the operations of assembly and testing
of the "Saturn-V" booster rocket in the vertical assembly building and
also the demand of all of the users of the iaunch complex and the booster
rocket itself.
In addition, the gaseous nitrogen and helium are used during the operations:
' The nitrogen for control of the valves of the filling system for the fuel
components and the valves of the rocket during the pre-launch preparation; _
the purging of the different equipment of the booster rocket and the ground
systems in order to insure explosion safety; the operations of the
pneumatic cylinders of the mechanisms for turning the moving platforms
on the ser~;ice cable mast and blowing the tanks of the S-I fuel stage;
Helium for pre-launch blowing of the oxyoen in hydrogen tanks; charging of -
the on-board tanks; purging of the different equipment of the booster rocket
and the ground systems; bubbling of the oxygen in the rocket. tanks ana
control of the individual mechanisms.
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CHAPTER 6. THERMOSTATING SYSTEMS
6.1. Purpflse, Structure and Composition
Thermostating is insurance of the given thermal conditions of the space
rocket system or its elements during the process of their ground prepara-
tion in order to create conditions for normal functioning of the on-board
equipment and systems. The given regime is insured at the engineering
complex, when transporting the space rocket system within the boundaries
of the space center and at the launch complex.
- As a rule, it is not the space rocket system as a whole that needs maintenance
of thermal conditions, but only individual components of it.l
The space rocket systems include equipment requiring def ined temperature
conditions for operation, variation of which lowers its characteristics
and disturbs the normal functioning. The temperature also determines the
characteristics of the on-board electric power supply sources, the operating
reliability of the engine assemblies, the thrust of the solid fuel boosters
in the rockets and the engines of the emergency rescue system for the
space vehicles.
The thermostating of the fuel components insures the given temperatures of
the oxidant and the combustible component going into the engine and the
required density, and the thermoseating of the cryogenic components,
reducti~n of their losses from evaporation. The space vehicles are thermo-
stated in order to maintain th~ required air temperature in the compartments
- with the equipment, the structural elements and individual assemblies and
and units (the instruments, power supplies, the serviced tanks and so on)
and als~, what is extremely important, in order to provide life support for
the cosmonauts during the pre-launch preparation time.
The optiutal temperature range depends on the cotnposition o~ the bo~os*.er
xocket and space vehicle systems, the type o� installed equipment, the fuel
1Hereafter all these elements will be called "the thermostating targets."
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components used, the structural design of the solid fuel engines,'and so
on. Usually for the instrument compartments and the engines of booster
rockets the optimal range is from +5 to +25�C, and for solid fuel booster
rockets and the emergency recovery engines of the space vehicle, from
0 to +40�C or a positive temperature range without restriction of the upper
limit. For some booster rockets the temperatu.re rai~~e is not limited, and
such rockets can be prEpared and launched without the application of
special thermostating means at any surrounding air temperature. _
For space vehicles with ground preparations the range from +15 to +25�C
is considered optimal although deviations from these values are possible.
Thus, the air temperature in the operating zones of the installers and
testers in the spacecraft or space station is permitted from +10 to +30�C.
The preferable temperature for the containers with food installed in the
space vehicle (ship or station) is considered to be the temperature from
0 to +15�C. The admissible temperature range of the space vehicle after
filling its fuel tank at the filling station is significantly limited.
The thermostating problems include both supplying heat to the elements of
the space rocket systems and removal of it. The heat is usually supplied
when the rocket system is outside the facility at low surrounding air
- teMperature (when transporting, in the launch position, and so on), and it
is removed to the engineering complex; when transpcrting the space rocket
system (the~.top module) within the boundaries of the space center and in
the launch complex. In the engineering complex usually excess fuel is
removed from the inside volumes of the space vehicles during electrica.t
testing and also from the compartments where the installation men and test
people are operating. When transporting and at the launch complex the
rocket system (top module) is protected i.n the summer from high surrounding
air temperatures and solar radiation.
The given thermal conditiions of the space vehicles sometimes are insured
by jaint operation of the ground ~hermostating means with the on-board
thermal regulation system. The on-board thermal regulation system of the
- space vehicle usually is made up of two hydraulic loops cooling and
heating. The excess heat is removed from the internal volumes of the -
vehicle by the cooling loop which, by means of the intermediate liquid-
liquid thermostating heat exchanger is connected to the outer loop the
ground liquid system for supporting the thermal conditions.
In the hydraulic main of the inner loop, a temperature is maintained
which is required ~or operation of the assemblies of the vehicle realizing
heat removal from the atmosphere of the compartments. The temperature
in the loop is re~ulated either by temperature variation and ~he amount
of cooled heat exchange agent in the outer line or by the regulating
elements in the loop itself. The heating loop is also connected to the
outer loop using the intermediate liquid-liquid thermostating heat exchanger
(ZhZhTT); in this case, the heated heat exchange agent is fed to the outer
loop.
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In order to thermostat the elements of the space rocket system, air systems
are also used to maintain the thermal conditions (VSOTR) which both supply
- and remove heat. The thermostating air is fed to the instrument compart-
ments and the engines of the booster rocket, under the nose conP or the
space vehicle elements are blown by it.
Maintenance of the thermal con3itions during the period after completion
of the operation of the ground thermostating means is insured as a result
of preliminary bringing of the temperatures of the structural Elements and
air in the compartments of the r~cket sys~em to the given levels during
thermostating process, determined by the outside temperature conditions and
the degree of thermal installation of the elements of the rocket system
and also as a result of selecting the time for disconnecting the thermo-
stating means.
During the thermostating process, the temperature is monitored at the most
important (from the point of view of the thermal conditions) Foints of thn
vehicle by using special temperature gauges connected with the ground
panels. The readings of the gauges are used to control the thermostating
conditions.
6.2. Classification of the Systems
The ground thermostating systems are classified by the method of thermo-
stating, the heat-exchange agent used, and mobility.
With respect to the method of thermostating the systems are divided into
active and passive thermostating systems and cambined systems.
The active thermostating systems provide for supplying heat to (removing
from) the vehicle and have sources of heat or cold and equipment for
supplying the heat-exchange agent in their composition. These include -
the air (VSOTR) and liquid (ZhSOTR) thermal conditioning systems.
The passive thermostating systems insure the given ~hermal conditions as
a result of insulati:ng the vehi.cle from the environment. They include
the thermal insulat~.ng hoods for the top modules, the space vehicles and
the engines of the emergency rescue systems which decrease the heat
exchange between the vehicle and the environment and also various coatings
with different coefficients of reflection and absorption.
The systems of combined means usP methods of both active and passive
thermostating. These are the electrothermal hoods in which the limitation
of the heat exchange with the environment is realized as a result of
thermal insu].at3.on, and the heating, as a result of the electric heaters.
With respect to the heat-exchange agent used the thermostating systems
are divided into air and liquid systems; in the air system the heat
exchange agent is air; in the liquid systems, it is different liquids
(brines, freons, antifreezes).
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With respect to mobility the thermostati.ng systems are divided into portable
and s:.ationary. The choice of one version or another depends on the cold
and heat requirement of the thermostating object, the required flow rate
of the heat transfer agent, the conditions of applying the system and also
the composition and structure of the ground complex.
The systems with low cold and heat output capacity are usually portable.
" They have a source of electric power (they can also be fed from an outside
source), and they are used both at the engineering and launch complexes.
They are universal and mobile, which makes it possible with a limited
number of units of equi~ment to provide f or thermostating in the different
stages of the technological process cqcle for preparation of the space
rocket system. Such systems are used, as a rule, for light class rockets.
The portable units are used also to insure the thermal conditions of the
space r~cket system for its elements during transportation within the
boundaries of the space center. Thus, using the portable units, the space -
vehicle or the top module is thermostated on delivery to the filling station
and also when being hauled from the engineering complex to the launch ~
complex as part of the completely assembled rocket system if the transport
time exceeds the admissible for which the normal conditions of the vehicle
or the top module will remain within the given limits.
For rockets of inedium, heavy and superheavy classes, the equipment of which
is basically stationary, the thermostating systems are also stationary.
The support of the thermal conditions of such rockets will require high
cold and heat output capacity which it is impossible to achieve by using
mobile units. The stationary systems are placed in special structures
and part of their equipment, in other ground units and systems (for example,
on the service tower of the launch complex, and so on). The portable
thermostating units are used in these cases only totransport the space
rocket systems or elements of them.
6.3. Sources of Cold and Heat
In the thermostating systems the sources of cold are compression type
refrigeratior. units, turbocompressc:rs and turborefrigeration units, systems
which use choking of gas and devices with the application of the vortex
effect, ard the sources of heat are electric and water heaters, and so on.
In the thermostating systems, in order to cool the heat-transfer agents,
the piston reftigeration units have received the greatest application.
In these units, the heat is picked up as a result of boiling the cooling
agent (usually f reon) in an evaporator with subsequent compression of its
vapor in the piston compressor; in this case the picked up heat is
transferred to the water or air. The piston compression refrigeration
units, in spite of their complexity, have been quite reliably developed -
at the present time, and they do not cause any special di~ficulties in
operation or maintenance.
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In the turbocompressor the compression of the coolant vapor takes place as
2. result of the creation of a centrifugal force on rotation of the impel7er;
directing the vapar successively through a number of impellers, it is
possible to obtain the required degree of coaipression.
The turborefrigeration units are a combination of a turbocompressor with
regenerative heat exchangers, and they operate by the thermodynamic cycle
called the "Russian cycle." The turborefrigeration unit which produces
1 kg of cold air persecond raith a temperature from -80 to =135�C and a
cold output capacity to 30 kwt is made up of a turbine (the turbine expansion
engine) and compressor on a single shaft, regenerator, the refrigeration
chamber, valves, air bypass units, a booater and switching mechanism.
The advantages of the turborefrigeration units are the low mass, small
size and the possibility of using atmospheri~ air instead of expensive
refrigeration agents,and the deficiencies low air pressure at the output
which complicates their application in the thermostating system.
The systeas using choking of gas are based on the principle of a reduction _
in gas pressure on passage of it through a constric.ted opening with simul--
taneous reduction of the temperature. These syste:n.s include *_he pneumatic
panels with reducers (valves, diaphragms), which step down the pressure and
hoses with sprayers; the air which comes out of the sprayers cools the
thermostating object. The advantages of these systems are simplicity of
structural design, hioh reliability and eass of servicing, and the
deficiencies include the low eff iciency of the cycle, the nonregulatabili.ty -
of the air temperature and the possibility that moisture will get into the ~
object, which is condemned in the structural elements of the various
assemblies, hoses and sprayers during the air cooling process.
The devices using the vortex eff ect (the Rank-Khilsh eff ect) are used to -
obtain a flow of air cooled to -(10 to 60)� and heated to +(50 to 100)�C.
The main part of these units is an eddy tube (the vortex refrigeration
unit) into which the air that has been compressed in the compressor in
advance goes. In the tube the air acquires a.n eddy motion, as a result of
which the inner layers are cooled, and the ou*_2r ones are heated.
The eddy tube (Fig 6.1) is a smooth cylirn~rical tube equippe3 with a unit
with tangential nozzles, a diaphra~n with axial opening and a choke. On
escape of air through the nozzle, an intensive circt~lar flow is created,
the axial layers of which are cooled, and they leave in the form of a cold
flow through the opening in the diaphragm, the hose, a muffler~and
a fitting; the peripheral la5~ers are also heated in the form of a hot flow
and they exit through the tube, the hose and the muffler. As the
_ choke is covered, the cold flow through the opening of the diaphragm
increases with a corresponding increase in the flow rate of the hot flow.
The devices usually have several tubes, at the exit from which air is
obtained with different temperature. The devices also include the air
preparation unit which includes filters, oil separator and water heat
exchanger for preliminary ~~ooling of the air after the compressor.
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The advantages of the devices with the application of an eddy effect are
simplicity of structural design, operating reliability and possibility of
fast temperature regulation, and the deficiencies include a low eff iciency
cycle and powerful noise occurring during operations; the noise is reducsd
by the application of mufflers on the fittin~s of the "cold" and "hot"
ends and also sound insulation of the tube casing.
Electric heaters with different power serving both to heat the air (gas)
and the liquid heat-exchange agents are used as the sources of heat in
- thermostating systems. In order to regulate the temperature, the heaters
are usually made of several sections included in various combinations.
The water heaters use hot water from the boiler room of the space center or
from other systems where it is a byproduct. In systesns with the application
of the eddy eff ect for heating, air emitting from the "hot" end of the eddy
current is used.
6.4. Structure of the Thermostating Systems
The air thermostating systems, independently of the structural design, have
sources of cold and heat, tanks with coolant, a system of lines with regu-
lating f ittings, pipelines for supplying air, a mect~anism for removal of
- the on-board split connection and a con~trol system.
The largest VSOTR systems used for thermostating apace rocket systems ~t -
the launch complex have a significant influence on the composition and
placement of the structures and other systems.
In addition to the VSOTR and the ZhSOTR systems, the cold users can also
be the thermostatin~, systems for the fuel components, the air conditioning
system for the facilities, and so on.
_ In the stationary thermostating systems all of the refrigeration equipment
is placed in a single refrigeration center, which insures the demand for
cold and heat. This composition of the equipment makes it possible to
use it more eff iciently, increase the efficiency, and the neutralizer
servicing and operation.
Th~ cold center can be placed both in an individual structure and in the
launch facilities under the pad. The special structure for the refrigera-
tion center is usually of an arch type, semiburied with the necessary
protection in case of explosion of the rocket on launch. ~
For more ~ff icient use, the refrigeration equipment is placed as close as
possible to the user (rocket), since with an increase in distance the
heat losses increase significantly. The equiptnent must be compact, fire
and eacplosion safe and automated to the maximum. The large refx.igeration
equipment (the refrigeration unit, the heat exchange units, the tanks with
a system of li~es and the regulati.ng fittings for intermediate heat-
exchange agent, the pumps, water supply systems) and also the ccntrol
y 201
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panels with the monitoring and measuring instruments requiring the
presence of service personnel usually are placed in an individual structure,
and the channels for supplying air, in the facilities under the pad and
on the service tower.
~
The air feed channels include fans, filters, air coolers, electric heaters
and air ducts. The fans (air blowers) are designed to create the required
- air pressure and feed; the filters are designed to remove dust and mechani-
cal impurities from them, the air coolers are made to cool the equipment
to the required temperature and settle out muisture contained in it; the
electric heaters are used for heating. The air ducts are placed on the
service tower and lead to the rocket at the corresponding levels.
.
The VSOTR systems operate both by the open and closed cycles. In the first
case the air fed to the rocket is discharged to the atmosphere through
one of the hatches; in the second case, it returns to the system.
In order to remove the on-board connections with the air ducts on the
service tower platforms disconnect mechanisms are provided. These mechanisms
~ remove the on-board connections with the air ducts to a safe distance,
excluding the possibility of collision of them with the rocket under the
effect of wind loads. At some launch complexes (basically for light and
medium class rockets) the d~sconnection and removal of the on-board split
connections are accomplished manually.
The air systems for supporting the thermal operating conditions are
operating both in the manual and in automatic modes. In the manual mode
the air temperature is selected by the operator and maintained by variation
of the temperature of the heat-transfer agent, its consumption in the air
cooler, disconn~ction (connection) of ane of the air coolers or inclusion
- of the required number of sections of the electric heaters. In the auto-
matic mode the air temperature is maintained by a sp.ecial device which on
disconnection of it from the given one sends a signal to the servomechanism
regulating the consumption of the heat transfer agent through the air
- coolers or it changes the number of connected (disconnected sections of
the electric heaters.
The means of supporting the thermal regime of the engineering complex
usually have separate air and liquid thermostating systems (although the
- possibility of combining them into a united refr3.geration center is not
excluded), and wiCh respect to construction principle they are analogous
ta the systems of the launch complex. However, as a result of the fact
that the cold and heat requirements at the engineering complex are less,
and the remote and automatic control frequently is not required, they are
si.mpler ~aith respect to structural design. The system equipment is
placed in the installation and test facilities for the booster rockets
and the facility for installation and testing of the space vehicles, in
additions on the buildings or in special buildings.
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- Let us consider the structure and the operating grinciple of the
stationary air thermostating system of the top module aC the launch
complex with air flow rate of 4 m3/sec and temperature from -10 to +40�C
(Fig 6.2). The system operates by the open cycle, that is, with discharge -
of the air to the outside, and it is fed from the coQling center, in
which the refrigation units, brine tanks for the heat-exchange agent, the
pump group, the system of brine and water lines with adjustable gate
f ittings, control panels and other equipment are located. The fans for
supplying air, the filter, the air cooler, the electric heater, the air
duct and the mechan~sm foT� remova.l of the on-board connections are placed
on the service tower.
The air is cooled in two steps: in the first step the atmospheric air
- forced by the fan and passing through the filter is cooled in the air cooler -
to a temperature of 2-5�C as a result of heat exchange with the heat-
exchange agent (27-29% calcium chloride solution) coming from the cooling -
center; si.multaneously the moisture precipitates out (to 95%) which is
contained in the collected air. It runs off to the bottom of air cooler
and is removed. In the second step the air in the air c~oler is cooled
to a temperature below 0�C with precipitation of moisture on the surface
- of the air cooler in the form of "frost." As the cross section of the air
cooler is decreased as a result of the formation of the "frost" the ai~
f eed from one air cooler is switched to the other, and in order to defrost
the first air cooler, a special fan and electric heater are switched on;
this air is not used for thermostating and is discharged through the
connection to the atmosphere.
_ During operation of the system and the heating mode, the air fed to the
top niodule is heated in an electric heater. In order to obtain the air
with given "dewpoint" it is first cooled in the air coolers where precip~ta-
tion of the moisture takes place.
The system operates both in the manual and in automatic modes. In the
manual mode the given ai'r temperature at the input to the top module is
selected and maintained by the operator by varying the flow rate of the
heat-exchange agent fed to the air cooler. A defined temperature of the
heat-exchange agent is maintained in the brine tanks. In the automatic
mode the given air temperature is maintained by instru~ents in accordance
with the readings of the temperature gauges~.installed at the input to the
top module. In this case, the oper~tor of the control system adjusts the
thermostat to the required temperature (variation in the temperature
automatically changes the flow rate of the heat=exchange agent), and the
given temperature of the heat-exchange agent is maintained in the brine
baths by varying the amount of coolant going to the evaporators. The
air flow rate is adjusted remotely by opening (closing) the air valves
or increa.sing (decreasing) the number o� fans~put into operation.
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204
FOR OFFICIAL USE ONLY
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APPR~VED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000200040010-4
FOR OFFICIAL USE ONLY
The portable air thermostating units are mounted in the body of a railroad
car or truck. Such units are fed both from the outside network and from
the inside source of electric power, which permits them to operate during
movement (when transporting the rocket or the top module within the limits
of the space center). The portable air units usually operate by closed
cycle without application of the intermediate feed-transfer agent, with
cooling of the air in the evaporator of the refrigerator; connection and
disconnection of the lines, as a rule, is manual. � ,
_ Let us consider one of the units designed for thermostating the top module
of the rocket when transporting it from the installation and test facility
to the filling station and also within the composition of the fully
assembled space rocket system during transportation from the engineering
_ complex to the launch complex.
All of the equipment of the unit (Fig 6.3) is placed in the railroad car
divided into several compartments. The source of cold is freon refrigeratc:rs
with air cooling condensers. A diesel generator is installed in the unit;
for connection to the outside current source there is a coil with a cable;
the unit operates by a closed cycle. The cooled or heated air is fed to
the top module located o n the railroad carrier (in the fully assembled
space rocket system, on the transport-erection unit), using an electric ~
fan by the system of stationary and flexible air ducts and adapters, and
then ie again goes to the air cooler of the unit. The control panel is
used to control the operation of the cooling and heat3.ng unit in the manual
and automatic regimes. In the automatic regime the given air temperature -
is maintained by the instruments in accordance with the gauge readings. -
Liquid Thermostating Systems. In the stationary ZhSOTR systems the heat-
exchange agent is cooled both from tlie cooling center (in common with the
VSOTR system) and from the autonomous source of cold. The feed lines of
the heat-transf er agent are placed in facilities under the pad on the
service tower. For protection of the ZhZhTT [liquid-l~quid thermostating
heat exchanger] of the space vehicle from excess pressure in the ZhSOTR
systems, a pressure relay is provided. In the portable units all of the
equipment is placed in the body of the truck, and the heat transf er agent
is cooled by heat exchange with the coolant of the refrigerator.
The liquid systems operate only by a closed cycle. On completion of
operation of the system, before disconnection and removal of the hydraulic
block from the rocket, the heat transfer agent is drained off, and the
lines are blown out with compressed nitrogen in order to protect the on-
board lines from corrosion and exclude incidence of the heat transfer
agent on board the rocket or space vehicle.
205.
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FOR OFFICIAL USE ONLY
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I
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APPR~VED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000200040010-4
FOR OFFICIAL USE ONLY
The operation of the ZhSOTR in the manual and automatic modes is analogous
to the operation of the VSOTR.
As an exanplP let us consider a portable liquid thermostating unit. The
~ unit (Fig o.4 and 6.5) is used to feed thQ heat-exchange agent (antifreeze)
with a temperature from -5 to +50�C to th~ ZhZhTT of the space vehicle ~
with a flow rate of 0.27�10-3 m3/sec, and it is mounted on the body
installed on the automobile chassis. The unit can serve as both the
launch complex and the engineering complex. When working on the launch
complex, in order to feed the heat-transfer agent, pipelines are used
which are laid on the service tower (truss); in the engineering complex
there are special pipeline~ for the installation and teat complex.
' During operation of the unit in the cooling mode, the heat-transfer agent
is moved by an electric pump from one division of the mixing tank to the
other; it passes through the evaporator where it is cooled by a coolant
(freon) which boils at low pressur e and temperature. the cooled heat ~
~ transf er agent is moved by an eleetric pump through a flexible hose, through
the pipeline on the service truss, through the distributor and the pressure
delivery hose to the tube space of the liquid-liquid heat-exchanger of the
- target. It picks up heat only from the space vehicle and is again
drained ~nto the mixing tank of the unit. At the input to the ZhZhTT
heat exchanger there is a pressure relay which shuts off the pump with a -
rise in pressure above admissible. ~
n
u
16
t J f b i 7 C 9
~
/ n
_ 20�
~ li Il R
Figure 6.4. Portable liquid thermostating unit:
1-- truck chassis; 2-- body; 3-- mixing tank; 4,5 electric
pumps; 6-- compressor; 7-- f3.lter-drier; 8-- electric heater;
9-- evaporator; 10 heat exchanger; .11 receiver; 12
heat regulating valve; 13 solenoid valve; 14 manual regulating
valve; 15 flexible hose; 16 pressure hose; 17 service
. trtlss; 18 resistance thermometer; 19 distributor; 20
pipeline on the service truss
207
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APPR~VED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000200040010-4
FOR C~'FICIAL USE ONLY
~
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Figure 6.5. Schematic of the portable liquid thermostating unit:
1-- ZhZhTT.~ 2-- pressure hose; 3-- distributor; 4-- pressure
relay; 5-- pipeline on the service truss; 6-- flexible hose;
7-- electric pump; 8-- shutoff valve (coupling); 9-- an~ifreeze
line; 10 flow rate relay; 11 resistance thermometer;
12 mixing tank; 13 electric heater; 14 manometer;
15 electric pump; 16 shutoff valve (angular); 17 evaporator;
18 safety valve; 19 heat regulating valve; 20 manual
regulating valve; 21 solenoid valve; 22 filter-drier;
23 heat exchanger; 24 receiver; 25 freon section line;
26 condenser; 27 compressor; 28 freon delivery line
Key:
a-- filling with antifreeze; b-- cold division; c-- warm division;
d-- heat exchange agent drain; e-- water output; f-- water input.
The coolant vapor formed during boiling in the evaporator is removed by
the compressors through the intermediate space o� the heat exchanger in
which they were superheated as a result of the counterflow of freon from
the receiver. The compressor compresses the freon vapor and pumps it into
the condenser where it ~.s cooled by water and condensed.
208
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~ ' FOR OFFICIAL USE ONLY
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209
_ FOR OF~ICIAL USE ONLY
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APPR~VED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000200040010-4
~OR OFFICIAL USE ONLY
Liqutd Ereon tormed in tlie condenser drains into the receiver, f rom wliict~
it goes through the coil of the heat exchsnger, the filter-drier and ttie
solenoid valve to two parallel heat regula~ting valves. The supplied
- high-pressure coolant (at condensation pressure) is fed through the choke -
cross section of the heat regulating valve to the evaporator wl?ere low
pressure is maintained by compressors sucking vapor out of the evapora~or.
The liquid freon, running through the evaporator space between the tubes,
boils at low temperature as a result of the heat influxes from the heat-
transfer agent circulating through the evaporator tubes. Then the cycle
is repeated.
The temperature of the heat-transfer agent is monitored by a resistance
thermometer installed in a mixing tank and an electron bridge; on deviation
of the temperature from the given temperature, the campressor is switched
on or off.
When operating in the heating mode, the heat-transfer agent is heated by
the electric heater of the mixing tank, af ter which the pump is used to _
f eed it to the ZhZhTT of the vehicle and from there again to the mixing
tank.
The stationary ZhSOTR system (Fig 6.6) receives cold from the refrigeration
center of the launch complex. In the hea.t exchan~er, heat exchange takes
~ place between the hea.t-transfer agent (antifreeze) and the intermediate
heat-transfer agent (freon) fed by a centrifugal pump from the expansion
tank. The antifreeze is heated in the tank of the electric heater and the
piston pump feeds it through the system lines to the ZhZhTT of the vehicle
from which it is drained back into the tank. The flow rate of the anti-
freeze is established by regulating valve; the flow rate is monitored by
a flow gauge. When the pressure is exceeded at the input to the ZhZhTT
heat exchanger the pressure relay switches the pump off. _
Thermal Jackets. A thermal jacket without electric heating (passive thermo-
stating) is designed to decrease the heat echange between the elements of
- the space rocket system and the environment and also for protection from
. the meteorological effects and solar radiation. The thermal jacket is
made up of strips of cloth between there is a�iller (foam plastic or
other insulating material), with a split along the generatrix covered by
the fastening locks with traction belts. _
The electrothermal jacket (thermostating by a combined method) insures the
thermal regime o~ the eng3nes of the emergency rescue system, the solid-
propellant boosters, fuel tanks and other elements of the space rocket
system.
The electrotherm~.l jacket (Fig 6.7) for the engine of the emergency rescue
system mainta~.ns the temperature above +15�C and is a strip to which _
electric heaters, resistance thermometers and thermal resistances making
up the electrical part o� the jacket are fastened. The strip is made of
foam plastic covered with rubberized balloon material with a split along
the generatrix. Locks are attached on one side of the strip, and on the
other, holders with turn buckles which are fastened to rubber shock
210 �
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FOR OFFICIAL USE ONLY
absorbers. The holders are f ixed in the locks by jaws connected to each
ather by the opening line. At the top there is a ring by which the jacket
is held by the service truss crane after opening, and then it is lowered
on the platform.
The electric heater is in the form of two strips between which wires are
glued which make up the heating elements; the ends of the wires are soldered
to the contacts which are taken out through a plug. .
The resistance thermometers are the sensors of tY~e ratiometer of the
control panel indicating the temperature under the ~acket; the thermal
resistances are the instrument sensors of the panel that automatically
regulates temperature.
During the summer, the top of the jacket is f itted with a protective shell
made of rubberized ma.terial to which metallized film is glued which has a
high coefficient of reflection of sun rays.
The thermal conditions under the ~acket are maintained automatically;
when the temperature deviates from the given one, sound and light signals
are sent. The visual monitoring and manual regulation of the temperature
, are provided from the monitoring and control panel.
c
.
211
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FOR OFFICIAL USE ONLY
. f
2
/6
~ 3 -
A A
` -
~S
~ A-A
s -
~a ' e ~
J4
~
13 1? .
l1 1Q
Figure 6.7. Electrothermal jacket:
1-- ring; 2-- fabric strip; 3-- shock absorber; 4-- plug; 5-- opening _
line; 6-- lock; 7-- jaws; 8-- holder; 9-- turn buckle; 10 thermal
resistance; 11 beds; 12 resistance thermometer; 13 electric
heater; 14 balloon fabric; 15 foam plastic; 16 protective shell
212
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APPR~VED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000200040010-4
FOR OFFICIAL USE ONLY
CHAPTER 7. COMMUNICATIONS OF THE GROUND SYSTEMS WITH THE ON-BOARD SYSTEMS
GROUND-ON-BOARD" COMMi1NICATIONS)
7.1. Nature of the "Ground-On-Board" Communications
In order to perform the various operations during preparations for launch
(filling with fuel components and compressed gases, various checks and
' measurements, purging, pre-launch purging of the tanks, maintenance of the
thermal conditions of the on-board systems, and so un) the rocket installed
on the launch system is connected to the ground launch systems through
electrical, pneumatic and hydromechanical plug connections, f.orming the
so-called "ground-on-board" communications.
As the pre-launch operations are performed, the number of couplings with
the ground systems decrease as a result of uncoupling the plug cannections
and removal of the ground lines to a saf e distance. Since some of the pre-
launch ground operations are joined or close in time to the startup of
the first stage engine, many of the plug connections are disconnected in
the initial phase of movement of the rocket (Fig 7.1). The organization
of these communications, laid still in the design stage, to a significant
degree determines the operating convenience, reliability and efficiency
. of the rocket complex. The classification of "ground-on-board" communica-
tions is presented in Fig 7.2.
In order to simplify servicing and decrease the number of plug connections
of the numerous lines brought to the rocket,they usually are combined into
several groups (trunks).
Both related (that is, hydraulic, pneumatic or electric only) and unrelated
lines are run through the plug connections; for safety reasons the combining
of the lines, which when damaged will possibly cause an emergency (for
example, mixing of self-igniting components) is undesirable.
The lines brought through the plug connections of the rocket are laid along
speci.al units the service towers (trusses), the fill cable and the
fill-drain tnasts, and so on.
21:i
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FOR OFFICIAL USE ONLY
~
:
.
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Figure 7.1. Disconnecting the "ground-on-board" communications
in the initial stage of flight of the rocket
The plug connections basically have a common design. They are made up of
two parts (panels) joined along the plane of the plug, and they insure
the required connection of the lines running through them. The lines
running from the ground systems are connected to one part of the plug
(the ground plug), and the lines of the space rocket system are connected
to the other part of the plug (the on-board part). Both parts are ke~t
tightly connected b3~ the special device (lock) which separates them at
the required ti.me, pushing the ground part away from the on-board part.
The structural design.of the plug connection is determined by such factors
as the method (manual or remote) of coupling (uncoupling) the parts of
the plug connection, the number and the sizes of the transverse cross
section of the lines running through the plugs, their purpose and type.
The plug connections must satisfy the requirements of simplicity and
convenience in operation and maintenance, seal of the pneumohydraulic
lines and contact coupling of the electric lines, the maxi.mum possible
214
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- FOR OFFICIAL USE ONLY
distribution of the mass of the structure on the ground (removable) part
and also insurance of reliable remote monitoring of the positions of the
parts of the plufi, high speed operation of the lock, independent redundancy
in the response means and protection of the on-board and ~round lines
from the environment on uncoupling.
. ~1 ~ ~"nQCCU~iu~raqu,v ccr+~eu
,,,e,waA-Qopm" (DopmaBae
~OG.lbCMNb/C COCB(lNCNUA~
(2) ~3)
~o muny KOMMj~N(M'Q[(U(L ~a
~
_ (6) (8) (9) b~ ~ ~ (10) -
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F s $ O o~ ~"`F
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_ `
no ~+tcmy na xapa7mepy npa~ademrr~v nn ,ruu~uu{enMOUnn ~
pacnoaa,~eNUa n~ueccne crnaRO pac~r,on~eae~,na.~ cmei~_
_ na pa~reme (11 paccineiKae,ra~~) ~aeMnax caedFiNe~ruu 13) _
~'y 5$~~ o i o
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1 17 18) 19 ~ 20 21 22)
Figure 7.2. Classification of the "ground-on-board" communications
Key:
1. Classif ication of "ground-on-board" communications (on-board plug
connections)
2. with respect to type of line 13. by the protection of the
3. with respect to time of uncoupling open joints of the open
4. electric connections
5. pneumatic 14. on the lateral surface of
6. hydraulic each s+~age
7. mechanical 15. on the end or lower part
8. mixed type of the lateral surface
9. bef ore the beginning of takeoff of the f irst stage
of the rockets 16. mixed type
10. in the initial takeoff phase 17. manual
of the rocket 18. remote
11. by the location on the rocket 19. mixed type
12. by the nature of performing 20. with protection
the connecting and disconnecting 21. without protection
process 22. mixed type
215
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The manual connection (disconnection) is used for simple plugs of small
mass in easily accessible locations and serviced before the beginning of
filling the rocket with fuel. In this case the ground part of the plug
- is connected to the on-board part manually using flanged or threaded
connections. The ground part is the end of the line of the corresponding
ground system.
The plug connections with a large number of lines having signif icant
transverse cross sections and rigid requirements with respect to seal are
most complex with respect to structural design. In order to facilitate
the servicing, the coupling of these plugs is completely or to a significant
degree mechanized, and sometimes they are remotely controlled and mon~tored.
In this case it is necessary to have complex additional equipment which
is placed directly in the service zone of the plug. In order ta improve
the quality af coupling, it is done at the engineering complex with subse-
quent connection of the ground lines at the launch complex to the adapters
= using simple flanges or threaded connections.
Remote uncoupling arises from the fact that some of the plugs must be
connected until a defined time of pre-launch preparation of the rocket,
and access to them by the service personnel is forbidden by saf ety
engineering requirements and also the effort to decrease the labor consump-
tion of the work done at the launch complex.
The ground lines (especially the large-di.ameter hydraulic and pneumatic
lines) must be suff iciently flexible and strong and provide for signifi-
cant, especially under high wind blows, mutual displacements ~.,f the
- booster rocket and the service unit on which they are laid. ~n order
to compensate for the mutual displacements with sma11 ampJ.itudes and
frequencies af the oscillations, flexible hoses are used; for signif icant
amplitudes and oscillations, a combination of flexible hoses, hinges and
other a~semblies are used providing for rotation of the lines in the
required planes.
The heavy lines require special mechanisms for bringing them in and remov-
ing them during the connecting and unconnecting process; the mechanis~s
take the greater part of the weight of the connected lines, which lowers
the load on the plugs and at the same time simplifies their design.
From the point of view of reliability it is desirable to have a system in
which the "ground-on-board" couplings are disconnected in advance (before
starting the f irst stage engine), for a failure or delay in uncoupling
and removal of the ground lines to a safe (from collision with the
ascending rocket) place can result in an emergency. The use of this
system is connected with reducing the efficiency of the rocket, for the
preliminary discontinuation of the feed of cryogenic components and
compressed gases to make up the tanks and bottles and the electric power
for the on-board user leads to partial consumption of it before launch
and also to the necessity for repea.ted remotely controlled and monitored
coupling of the connection (in case of emergency shutdown of the engine
216
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during launcli) For the dralu lines, che inert gas feed, and so on, wliich
significantly complicates the design of both on~-board and the ground
parts of the "ground-on-board" communications. Therefore, the rocket has
a number of couplings which are disconnected only after starting the
engine of the first stage, which in practice coincides with the beginning
of the rocket flight. These comnunications include the topping lines
for the cryogenic fuel components, the lines for pre-launch blowing of
the tanks, the lines for purging the protective cavities and the systems
for f ire and explosion prevention, the electric line of the measurement
and control systems participafing in the launch operations, the lines for
which repeated coupling of the connections is complicated as a result of -
safety requirements and the devices holding the rocket on the launch system
until the engines reach full thrust and in certain cases insuring a given
change in G-load wr~n separating it from the launch system.
Uncoupling the plug connections and removal of the uncoupled lines at the
beginning of flight of the rocket are technically difficult and for
implementation require well thought-out and well-developed structural
schemes. Primary attention has been given to the problems of high speed
~ of the blocks where the plug connections and the mechanisms for removal
with insurance of independent duplication in them of the response, exclu-
sion of the collision of the removed lines and the rocket.
The uncoupled connections of the first stage of the booster rocket usually
are taken after the lower part of the lateral surface of the booster
rocket or to its end. They are serviced from the launch system or from
small size units. -
In order to provide couplin~s for the upper stages of the booster rocket
and the space vehicle, two versions are used. In the first version all
the lines or the greater part of them are taken out to the first stage,
which although it simplifies servicing, significantly complicates and
, increases the weight of the structure of the booster rocket as a result
of the placement of the li~les and auxiliary equipment on the lower stages
required only for pre-launch preparation of the upper stages. In the
second version the plug connections are on the side surface of each stage
and are connected to the ground ~ystems through the service tower or the
service cable mast which, complicating the servicing, decreases the length
of the on-board lines and do not require complex plug connections between
s tages .
- The service towers usually are pulled back from the rocket a significant
time before launch; therefore lines are put on them which can be unplugged
in advance. The trusses of the service cable towers and the cable mast
are pulled back from the rocket directly before launch or during launch;
therefore the lines are laid on them which are disconnected in practice -
when starting the engine. '
For modern space rocket systems, as a rule, a combination of various
systems of maintaining "ground-on-board" communications is used:
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the couplings which are disconnected when the rocket system lifts off
are led out through the lower part of the f irst stage (usually the
electrical and pneumatic lines for the entire rocket and the fluid lines
Eor ttie first stage); through the upper stages couplings are made with
advance disconnection (the thermostating system, the system for servicing
with high-boiling fuel components, the adjustable plug connections, and
so on) and couplings with disconnection of the lines directly before
launch or at the time of launch (electrical, pneumatic, topping off, and
so on). This variety of systems is caused by the effort to create the
most eff icient on-board systems for the rocket systems and to mainCain
operating reliability of all the "ground-on-board" communications. -
7.2. Standard "Ground-On-Board" Communications Layouts
The layouts for the "ground-on-board" communications will be considered
in the example of the communications over the thermostating line. The
= on-board plug connections for the gas and liquid lines usuallya~e called ~
pneumatic and hydraulic blocks. The lines for the air thermostating
system (VSOTR) are co~ected to the pneumatic blocks, and the liquici
thermostating system (ZhSOTR), to the hydraulic blocks.
The layouts for the "ground-on-board" communications with respect to the
VSOTR line (Fig 7.3) operate on the following principle. The air ~uct -
of the VSOTR is fastened through the adapter by means of the frame and ~
guide to the sliding lift mechanism which serves not only to feed the
air duct, but also to compensate for mutual displacements of the rocket ~
and service tower. $
The pneumatic lock of the block is opened when compressed gas is fed to ~
it, and it repels the ground part of the pneumatic block to a short j
distance from the rocket. The pneumatic block together with the connected -
air duct is moved by the withdrawal mechanism to the required distance
" and is f ixed by a catch in the terminal position. The kinetic energy of ~
the withdrawal mass is absorbed by the shock absorber. Completion of f
withdrawal is monitored by a signal which is sent to the system that F
controls these operations. {
~
The layout of the "ground-on-baard" communications with respect to the -
- ZhSOTR line (Fig 7.4) has by comparisun with the precedin device all the ~
degrees of freedom for displacement of the hydraulic block with respect ~
to the service tower. Before uncoupling, in order to avoid spilling the ;
heat-exchange agent on the side of the rocket, the lines for the hydraulic
block are purged with gas to completely remove the rema.ins of the heat-
transf er agent. After uncoupling the withdrawal mechanism is rotated ~
by means of the pneumatic drive, the final position of which is fixed "
by the signal unit. On completion of rotation, the disconnected part ~
_ of the hqdraul.ic block with the ground lines is lifted to the extreme
upper position. ;R
~
~
218
~
,7
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~ Z 3
+ + ~
+ ' ~
1
4
S .
~ ' ' .
~ 7,
~o s ~
~ B
~
. Ca, � r . .
_ 14 ~ ~ ~ ~
. ~ !2~ . ~ -
. ~
Figure 7.3. Layout of the "ground-on-board" communications
with respect to the VSOTR line:
1-- pneumatic block; 2-- adapter; 3-- air duct; 4-- frame;
5-- guide; 6-- lift mechanism; 7-- catch; 8--shock absorber;
9-- support; 10 pneumatic drive; 11 signal unit;
12 base; 13 service tower; 14 top module
Key:
a pressure f eed _
The "ground-on-board" communications are recognized using ordinary, split
(ShR) and contact-breaking (Sh0) plug connections and contact-breaking
plates which diff er from each other by purpose, structural design and
method of separation.
The split plug connection is used to provide a coupling when preparing
the space-rocket system up to launch time, including the initial phase
of liftoff, and its separation occurs as a result of the movement of
the space-rocket system. The operating principles of such plugs are
different: one of them splits as a result of simple separation of the
on-board and ground parts; other.s split as a result of the response of
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a special lock with a line attached to the launch system. Usually the
split plug connections are placed at the end of the booster rocket.
The contact-breaking plugs (Fig 7.5) are designed to provide electrical -
- communications which are broken before liftoff of the rocket. If the
- communications are separated ahead of time, the cables are laid on the
service tower, and the ground parts of the plugs are attached to the
corresponding trusses; if the couplings are broken several seconds before
launch, the cables most frequently are laid on the cable masts.
SQL(~NA 06CNJl~C!leQJW11 . ~8~
1 2 3 4 MBXQ~
MQ ONlQOBd ~b~
5 g
~
x
x
~ 10 il '
3
~ :
. + (
;
I
- 9 t 8 '
� � ' ~
� . % i
_ ~
~
. ~ ~ ,f
~ 1
j: '
, 1
' ~ r,
f .
Figure 7.4. Schematic of the "ground-on-board" couplings with
respect to the ZhSOTR line:
1-- connection for f eeding the heat-transfer agent; 2-- connection
for the pneuma.tic block; 3-- contact sensor; 4-- clamp;
5-- line; 6-- moving part of the withdrawal mechanism;
7-- signal unit; 8-- pneumatic drive; 9-- top module of the
rocket; 10 hydraulic block; 11 connection for removal of
the heat-transfer agent (the hoses for supplying the transfer
agent are not shown)
_ Key:
a-- service tower; b-- axis of rotation of the withdrawal mechanism
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,h,x~
Figure 7.5. Contact-breaking plug connection:
a-- view from the contact sealed side; b-- view from the cable
entrance side
- The contact-breaking plugg of different designs have the special lock of
the pressure (by hand) or electromagnetic type. Usually before beginning
to fuel the rocket the plug connections are disconnected manually; after
fueling this is done remotely by sending a signal to the electromagnet
= of the lock from the control panel. The lock responds, the round part of
- the plug separates under the effect of springs from the on-board part and
it is trapped by a basket (trap) on the cable mast.
The contact-breaking plate is used to provide for coupling a large number
of electric circuits. It is a massive metal plate with plug connections
and it is ma.de up of on-board and ground parts held in the coupled state
by a breakaway bolt. The coupling of the plates requires special attach-
ments and usually it is done at the engineering complex. For coupling to
the ground cable network the contact-breaking part of the plates has
cable adapters.
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At the launch complex after issuing the command to the breakaway bolt,
separation of the contact-breaking plate into two parts takes place.
The released ground part of the plate with the cable adapters is ejected
by springs from the side of the booster rocket and withdrawn by special
mechanisms to the service truss. The contact-breaking plates have contact
devices which signal the control panel about the execution of the command _
to separate the plates.
7.3. "Ground-On-Board" Communications of the "Saturn-V-Apollo" Space
Rocket System
Let us consider the schematic of the organization of the "ground-on-board"
communications and the characteristics of its basic elements in the example
of the "Saturn-V-Apollo" space rocket system.
The "Saturn-V-Apollo" space rocket system is installed before removal from
the vertical assembly building on the upper part of the launch system
made up of the launch platform and the cable service tower. All of the
"ground-on-board" communications are coupled.
On the launch platform are four supporting clamps at an angle of 90� to
each other which hold the rocket system during transportation, while it
- is at the launch complex and for several seconds after starting the first
stage engine. In addition, in the same area there are three tail service
cable masts attached to the rocket providing for (in addition to electro-
pneumatic feed) drainage of the liquid oxygen, filling and drainage of
fuel and air feed for air conditioning. Uncoupling of the ground communi-
cations at liftoff of the rocket takes place through these service cable
masts.
- The electrical, pneumatic and hydraulic communications, telephone and tele-
vision cables required for servicing and pre-launch preparation of the
booster rocket and the space vehicle at the launch com~lex are laid on
- the service cable tower. The coupli.ng of the on-board systems to these
lines takes place through the pre-launch (separated befo~e launch) and
launch (withdrawn during launch) service truss, the distribution of the
lines on which is presented in Table 7.1.
.
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Table 7.1
Distribution of Lines with Respect to the Service Trusses of
the Service Cable Tower for the "Saturn-V-Apollo" Space
Rocket System
Lines _
Service trusses Electri- High- Air con-
cal Fuel pressure ditioning
Command module - - 8 1
Service module 5 - 8 1
Instrument compartment 22 1 22 1
S-IVB engine compartment 8 2 42 1
S-II tool compartment 7 1 20 1
S-II intertank compartment 15 2 46 2
S-IC instrument compartment 3 - 8 2
S-IC intertank compartment - 2 5 -
Tail compartment (the tail 18 2 21 1
service cable mast)
All of the lines of the upper part of the launch system are led out to the
service zones of the launch stand (Fig 7.6); after installing the launch
platform on the supports of the stand they are connected to t:~e ground
systems of the launch complex through the coupling units, and they are
disconnected after launching the rocket before withdrawal of the launch
platform.
The supporting clamp arms of the launch platform (Fig 7.7) hold the space
rocket system until all of the engines develop the required thrust. If
one of the engines does not reach operating conditions during this time,
the f irst stage engine is shut down. With normal starting, after a defined
time the mechanisms for withdrawing the clamps (the withdrawal of the
clamps is made redundant by a pyrobolt if necessary)respond from two
identical (redundant) pneumatic systems (high-pressure helium). The
_ liftoff of the rocket is monitored by contact signal elements of diametrically
arranged clamps; in this case the signals generate a command to disconnect
the fast-disconnect couplings and withdraw the tail service cable masts
and the launch service trusses of the cable service tower; this takes
place when the rocket has lif ted approximately 20 cm.
The tail service cable mast (Fig 7.8) is a balanced structure with pneumo-
electric control and hydraulic drive and it is made up of a base, a
lever with a counterweight on which the corresponding lines are placed
with high-speed plugs and a protective housing. The fast-disconnect
coupling has two parts: one is on the booster rocket side and after dis-
connect is ~covered by a cover; the other part located on the service
cable tower consists of the housing with a special collet-type lock
- providing for connecting the plugs and uncoupling them with repelling of
the ground part away from the on-board part.
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_ ~ ~
8
a
+
f
2
3
S ' 4
Figure 7.6. Schematic of the exit of the ground lines to the
service zones of the launch stand of the "Saturn-
V-Apollo" space rocket system:
1-- electric power mains; 2-- auxiliary equipment; 3-- air
conditioning mains; 4-- electric cables; 5-- liquid-oxygen
lines; 6-- liquid and gaseous hydrogen lines
~ ! 2 3 4
~
~ I 5
a' . ~ i B
~ I
~ I ~
~ ~I ~
~ 8
1 1 9
~
\
. - ~ ~o
Figure 7.7. Supporting clamp arm:
1-- adjustable support; 2-- upper element of the arm; 3--
stop plate; 4-- cover; 5-- central element; 6-- leveling
attachment; 7-- pnewnatic distributor; 8-- winch; 9--
lower element of the arm; 10 bearing beam; a-- end of
bo~ster rocket
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0
0
D
~
6
,
5- ~
4 ~~3
Figure 7.8. Tail service cable mast of the launch platform
of the "Saturn-V-Apollo" space rocket system:
1-- protective jacket and truss; 2-- f eed line; 3-- base;
4-- hydraulic and pneuma.tic lines of the system; 5-- electric
line; 6-- arm with counterweight; 7-- ground part of the
flow connection
Before launching the space rocket system the pre-launch service trusses
of the intercompartment of the second stage are withdrawn (11 hours
_ 30 minutes), the spacecraft compartment (preliminary 43 minutes and
f inal 5 minutes), the intertank compartment of the first stage (50 seconds),
the instrument compartment of the second stage (16 seconds before launch),
and at launch time, the launch service trusses with the fill and drain
lines and the basic electrical and pneumatic couplings.
Provision has been made for mechanical redundancy of the uncoupling of
the plugs and the withdrawal of the service trusses operating at liftoff
of the rocket system in case of failure of withdrawal system.
Part of the couplings required for servicing the engines of the space-
craft and the auxiliary engine of the third stage operating on long-
storable high-boiling fuel components are supported from the movable
service tower which is withdrawn from the launch system 10 haurs before
launch.
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CHAPTER 8. GUIDANCE SYSTEMS OF THE SPACE ROCKET SYSTEM
8.1. General Information
The gyrostabilized platform of the booster rocket control system must be
oriented a defined way in space relative to tlie direction of the earth's
meridian and the coordinates of the launch pad determined during geodetic
preparation for launch in order to support insertion of the space vehicle
into a given orbit. The set of operations with respect to orientation
of the booster rocket or the elements of its on-board control system
before launch to obtain the given f light parameters is called guidance.
As a rule, additional orientation of the space vehicle with respect eo
the ground geodetic network before launch is not required, for it is
structurall connected with the booster rocket, and its location with
respect to the rocket is known.
The booster rocket and the control system sensors are oriented during
guidance relative to the launch coordinate system OX~Y~Z~ (Fig 8.1),
the origin of which coincides with the center of mass of the space rocket
system installed on the launch system. The OX~ axis indicates the f.light
direction and its position is determined by the launch azimuth A~, the
OY~ axis is directed vertically upward, and the Y~OX~ plane tangent to
the trajectory of ~otion of the space rocket system with the location of
the launch system is called the launch plane.
As a result of rotation of the earth and other factors the tra~ectory of _
motion of the space rocket system is a line of double curvature; therefore
it does not coincide with the launch plane and deviates from it.
The so-called bound coordinate system OX1Y1Z1(Fig 8.2, a) is connected
with the space rocket system, the origin of which is placed at the center
of mass of the space rocket system. The OX1 axis coincides with the
axis of the rocket system, and the direction of the remaining axes is ,
determi.ned by the location of the steering elements placed in the
; stabilization planes which usually are numbered in roman numerals. The
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I-III plane passing through the longitudinal axis of the rocket and
the steering elements is called the basic plane of symmetry.
The direction of the axes of sensitivity of the gyroscopic and inertial
sensors of the control system determines the inertial coordinate system
OXYZ (Fig 8.2, b), and the XOY plane is called the basic stabilization
plane. At the time of launching the space rocket system the axes of the
bound and inertial coordinate systems are oriented in a defined way with
respect to the coordinate axes of the laLncher. The required mutual
arrangement of all three axes of the coordinate systems is achieved by
verticalization, azimuthal guidance and ad3ustment of the rocket gyro-
platform.
r (1) b ny~'`a .
c nOGxocm X~
n
A~ N
A
:
o ~ ~u : : .+t
. ~
-
I
:~.f, ~ ~i
: : ~
. Ih ~ iK:j~
j
~ M ~ � I ~ _ ~ ~ ~ :
Z~ _
S
Figure 8.1. Launch coordinate system
Key:
1. Launch plane
- As was pointed out previously, verticalization of the rocket is a set of
operations with respect to bringing the space rocket system installed
on the launch system to a strictly vertical position. The greatest =
deflection of the axis of the rocket or the element of its on-board control
system in the vertical position must not exceed several angular minutes.
Verticalization is achieved by rotation of the support plane of the launch
system around two mutually perpendicular axes using lift mechanisms, and
it is performed either directly during erection or immediately after
erection of the rocket.
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- Azimuthal guidance is orientation of the OYl axis of the bound coordinate
system and the space rocket system installed on the launch system, in the
horizontal plane to obtain a given flight direction. It is realized either
by turning the rocket in the horizontal plane or orientation af individual
elements of its on-board control system.
The adjustment of the gyroplatform has as its purpose the matching of the
basic stabilization plane with the basic plane of symmetry. It is
accomplished by rotation of the base of the gyroplatform with respect to
the hull of the rocket af ter it. In order to guide the space rocket
system at the launch complex it is necessary to perform two preliminary
- operations: g~odetic operation of the launch and preparation of initial
data.
In the case of geodetic preparation of the launch, the coordinates of the
launch system are determined, and the orientation of the geodetic direc-
tions at the launch complex is carried out. The coordinates of the
launch system together with the coordinates of the flight trajectory of
the space rocket system are used when preparing the initial data for
launch, and the oriented geodetic directions, directly for azimuthal
guidance.
_ XI ~ X~
n,~~q , Y ,
_
U 2
'(1) i ��~~eNQ( )
? \ 3 ~+aq~~~WW
1 2 ~
~ ,
Y~ ~ 3 .
~ 4
~ f
I
~
.
i - q X
i Z Z ~~s ~ Y!
~ I
Z~ f II ~ .
II fi b
~
~ 'Y ~
a
Figure 8.2. Coordinate systems:
a-- bound; b-- inertial; 1-- stabilization engine; 2-- angle
gauge; 3-- gyroscopic; 4-- accelerometer; 5-- control prism
- Key: 1-- basic plane of symmetry; 2-- basic plane of stabilization
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The orientation of the directions is determination of the azimuth of
any straight line takzn as the orientation line; for the space rocket
system, as a rule, orientation is performed from the geodetic network.
When guiding at the launch complex, a large volume of various operations
are performed connected with determining the directions, measurement
of the angles and rotation of the instruments of the gyroplatform. In
order to decrease the time, the operations not connected with the location
of the space rocket system at launch are performed in advance, and an
effort is made at maximum automation of the guidance system itself.
8.2. Basic Devices of the Guidance System
With respect to physical principles used as the basis for the operation
of the elements of the guidance systems, their devices are divided into
optical-mechanical, photoelectric, electronic, electromechanical and
gyroscopic devices. The optial-mechanical devices are used to determine
*_he azimuth of the orientatian directions; the photoelectric devices,
for measuring the mismatch angle, the electronic and electromechan~cal
devices, for generation, amplification and converaion of the signals
during measurements of the angles, remote transmission of them and process-
ing of the angular mismatches, and gyroscopic, primarily as the measuring
elements of the control system or as gyrocompasses for providing azimuthal
guidanc e.
Optical-Mechanical Devices. An example of the optical-mechanical devices
is the theodolites which are widely used to determine the azimuth of
the oriented directions and for verticalization of the rocket.
For guidance of the space rocket system it is necessary to f ix both the
position of the launch plane and the position of the basic stabilization
_ plane usually fixed by mirrors and mirror prisms. The mirrors and prisms
are fastened to the stabilized platform during manufacture of it and they
are oriented with great accuracy with respect to the basic stabilization
plane of the rocket.
The rectangular mirror prism is most widely used (Fig 8.3), a characteris-
tic feature of which is the fact tha.t the light beam incident on the
hypotenuse face of the prism in some plane P exits from it back in the
0 plane parallel to the P plane. At points and d the beam is refracted
on the hypotenuse face of the prism, and at points b and c, it is
reflected from the silvered faces of the prism making up its legs. As a
result of this property it is not necessary exactly to verticalize the
prism in the YOX plane, for even if the entering beam will not lie in the
XOZ plane, the reflected beam goes in the opposite direction in the plane
parallel to its plane of incidence. If the mismatch plane measured by
the theodolite between its viewing axis and the perpendicular to the
edge of th~ right angle of the prism is not equal to zero, then the angle
between the incident and reflected beams is equal to twice the mismatch
angle; if the mismatch angle is equal to zero, then in the case of
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parallelness of the incident and reflected beams their images in the
vertical plane of the eyepiece coincide; consequently; the axis of the
_ theodolite is also perpendicular to the edge of the prism.
Y
X
.
.
b .
~i ~s: 9
- i~ Lr
.o " ~
~ .
O ~ ~
. Z
~
Figure 8.3. Rectangular prism
f Y k
~ S.
i
~
~ ~ ~
, . -
i ~ '
. a
,
, ,
_ . Figure 8.4. Autocollimator:
1-- mirror; 2-- objective; 3-- grid; [s light; S--eye-
piece; a-- field of view ~
The azimutha2 position of the monitoring elements is determined by the
autocollimation principle, that is the path of the light beams for
which they exit from the instrument as a parallel beam and, on being
reflected from the mirror surface, they pass through the elements of the
instrument in the opposite direction. If the surface of the mirror is
perpendicular to the viewing axis of the autocollimator (Fig 8.4),
the direct and autocollimation images of its grid coincide with each
other. Using the autocollimation principle, it is possible also to
solve the inverse problem setting~ of the control mirror or the
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prism of the gyroplatform i.n accordance with the given azimuth, for which
it is necessary to turn it with respect to the verticaY axis and,
observing the autocollimator, to achieve matching of the direct and
autocollimation images of the grid.
Photoelectric Devices. The source of the light signals in the photo-
electric devices is the gas light tubes, incandescent lights and lasers.
In order to convert the light signals to electric signals, various
_ photoelectric radiation receivers are used (photoelements, photoresistances,
photomultipliers, photodiodes and so on). The optical systems of these -
. instruments are designed to create parallel light beams, for focusing,
separation, connectian and change in direction of the light beams.
_ The photoelectric devices are used for automatic measurement of the small
mismatch angles between the basic stabilization plane of the space rocket
system and the launch plane, the generation of electric signals which
depend on the measured angular mismatches, the transmission of oriented
directions in the vertical plane and measurement of the azimuthal angles -
in a large range of variation of them.
Depen~~ing on th:: purpose, the photoelectric devices are divided into
goniometers, synchronous transmissions and angle gauges.
The goniometers solve the f irst two problems and are of two types wit":
external and inte~nal light signal source; they consist of a source of
radiations an optical system, a radiation receiver, signal amplifier and
converter. The light signal r~eceived from the radiation source is incident
- on the control prism ins~alled on board the rocket and, being reFlected
from it, is received and analyzed by the goniometer.
The goniometers can operate in the zero and measuring regimes: in the
zero regime the mismatch signal generated by the goniometer is fed to the
drive of the gyroplat�orm which is rotated until the base stabilization
plane coincides with the lau ach plane. The mismatch angle is measured
for the measuring regime, and the electric signa.l proportional to the
measured angle is generated.
Synchronous transmissions are designed to transmit oriented directions in `
_ the vertical plane from the base of the launch system where the guidance
instruments are located to the instruments on the space rocket system.
The angle gauges are used to measure large angles of rotation of the
various instruments and devices. Their measurement range reaches 360�
in this case.
The polarization devices (a version of the photoelectric devices) operate
on a polarized light signal and are used in optical synchronous trans-
missions and autocollimation goniometers. By comparison with the photo- y
electric synchronous transmissions, these devices have hi$h accuracy.
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Their application in the goniometers permits use of splitting of the
polarized light beam to generate a mismatch signal in two mutually perpendic-
ular planes, that ie, to measure the angles of deflection of the viewing
axis of the goniometer from the perpendicular to the mirror surface in -
two directions azimuthal and vertical.
The light beams emitted by ordinary sources are not polarized; therefore
the light signal in the polarization devices is converted in advance
using polaroids, polarizing and double-refracting prisms. In order to
exclude interference from outside light sources the polarized light
signal is modulated by the polarization element which changes its optical
properties under the effect of an electric or magnetic field.
The amplifier-converters are used to amplify the electric signals picked
up from the photoelectric radiation receiv~rs and other sensitive elements -
having low power insufficient for direct actuation of the regulating and
servounits and also for signal conversion. The electron, semiconductor
and magnetic amplifiers are the most widespread.
The electron and semiconductor amplifiers are distinguished by high
sensitivity. They are capable of amplifying the low-power signals. The
ma.gnetic amplifiers permit us to obtain high output power of the signal
and they have high reliability.
~Iodulators and demodulators are most widely used among the conversion
units in guidance systems. If the sensitive element operates on direct
current, and the servoelement, on alternating current, then the DC signal
is converted in the amplifying channel to an AC signal using modulators.
Systems i:~ which an AC signal is picked up from the sensitive element are
more wic~Pspread, for the light f lux itself is modulated, and the servo-
element o~~erates on direct current; the conversion of the AC signal to
DC takes place using demodulators.
The induction synchronous transmissions are designed for remote measuce-
ment of the angles of rotation of the various elements, remote rotation
of the elements themselves by defined angles and synchronous rotation
nf several axes mechanically not connected with each other.
The induction synchronous transmissions, in contrast to ~hotoelectric
and polarization devices, do not have the property of rigidity (one-to-
one spatial correspondence between the orientation of the sensor and _
the receiver). The sensor and the receiver of the induction synchrQnous
transmission in the matched position can have a spatial orientation and
cannot be used for vertical transmission of the orientation directions.
In addition, the induction synchronous transmissions have lower accurac~r
by comparison with the polarization transmissions.
The gyroscopic devices usually are used as measuring devices in the
- inertial flight control systems. The operation of the guidance systems
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is clcsely connected with the on-board gyroscopic instruments, for the
goal of azimuthal guidance is orientation of the axes of sensitivity of
these instruments with respect to the launch plane.
~
The application of gyroscopes is based on such properties as stability
consisting in an effort to keep the position of the axis of rotation in-
- variant in space and precession consisting in the fact that when applying
the moment along one of the axes of the gi,~mbal frames, rotation of the
gyroscope around the other gymbal axis takes place. A gyroscope with
3� of freedom (rotation around its own axis, horizontal axis and vertical
axis of the g~mbal) has these properties. On restriction of one of the
degree of freedom the gyroscope loses the property of stability and the
property of precession.
Gyroscopic devices are used for guidance of space rocket systems to main-
tain the oriented geodetic directions, for autonomous determination of
the azimuths of the orientation directions and stabilization in space of
the elements of the guidance system under the effect of various mechanical
disturbances on them.
8.3. Nonautomated Guidance Systems -
In the nonautomated guidance systems, the principle of visual determination
- of the position of the control element of the space rocket system with
respect to the ground geodetic network is used. For example, let us
consider th e method of guidance, the base for which is transfer of the
reference direction with the help of a two-channel autocollirna.tion tele-
scope with established base angle (90�) between the viewing lines of the
obj ectives of the two channels (Fig 8. 5) .
The two-channel telescope is in the form of two autocollimation tubes,
the viewing axes of the ob~ectives of which are at an angle of 90� to
each other; in this case the image planes of the two ob3ects are matched
by special optical elements in the field of view of one eyepiece. In
the autocollimation mode only one channel operates, the objective of which
~ is aimed at the control element of the guidance system the prism
the objective of the second channel is aimed at the electric stake giving
a defined geodetic direct.ion. At the center of the field of view of the
eyepiece on the grid of angular units a line is plotted which must be
matched with the autocollimation image from the control prism. The edge
of the control prism must be perpendicular to the viewing axis of the
autocollimation tube. The second image on the grid of the eyepiece will
be from the electric stake. The reading between the two images in
' provisional units is the angle which must be taken into account in the
g~xidance formula.
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4 S
, ~a~~~?
~ ~ B .
.
7
Q
2 ro
3 8
_
~ ~
~ , ~ .
. .
~r �
p -
- , .
ID
Figure 8.6. Schematic of the arrangement of the visual
guidance system equipment "goniometer-
prism-mark":
1,5 collimators; 2-- goniometer; 3-- pentaprism;
_ 4,10 stakes; 6-- mark; 7-- prism; 8-- space rocket
syste.~?; 9 angular guides
Key:
' a directinns of launch
In the visual nonautomated guidance systems, the "goniometer-prism-mark"
method is used (Fig 8.6) based on matching the images of two marks
located in two planes at diff er~nt distances in the field of view of the
goniometer. The oriented geodetic directions are formed by two collimators
and two stakes (to provide guidance along any azimuth). The "gontometer-
mark" direction perpendicular to the launch direction is established
through the pentaprism (pentagonal pri~m) with respect to these directions
and considering the direction (azimuth) of launch. The pentaprism pro-
vides for rotation of the angle by 90� independently of the degree of
perpendicularity of the beam to the plane of the prism face. Rotating
the rocket with the prism installed on the gyro, the edge of the prism
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is matched with the "goniometFr-mark" direction and thus guidance is
reduced to matching the images of the edge of the prism and the mark
in the field of view of the goniometer.
All the guidance operations are performed before putting the fuel
components on board the booster rocket, for after servicing, the presence
of service personnel in the direct proximity of the rocket is forbidden.
Therefore the visual guidance system cannot exclude the errors in
azimuthal guidance which occur as a result of deformation (twisting) of
the hull of the booster rocket after it is filled with fuel,.and it can
be used in cases where high accuracy of guidance in the launch plane is
not required.
8.4. Automated Guidance Systems
Considering that in the process of preparing the space rocket system at
the launch complex complex automation is f inding greater and greater
application, the nonautomated guidance systemu cannot provide for the
given requirements.
_ In modern space rocket complexes for azimu'thal guidance of the rockets,
completely automated systems are used which provide for automatic output `
of the required i.nformation to the flight control system of the space
rocket system. These systems can be of two types: single-channel and
double-channel.
The single-channel guidance system (Fig 8.7) includes the autocollimation
goniometer which tracks the reflector with the drive, the prism which
f ixes the orientation geodetic direction, the amplifier-converter unit,
the on-board control prism, the drive of the gyrostabilized platform,
the television transmitter and the guidance system control unit. Before
guidance geodetic gridding of the position of the viewing axis is carried
out, and a special prism is used for periodic monitoring of the position
of the goniometer.
Using the tracking reflector, the light beams leaving the objective of
the goniometer are rotated by 90�, and at an angle of 25� to the horizon
_ they are directed at the on-board control prism. On the basis of the
analysis of the light flux reflected from tb:: prism, the control signal
- is ~enerated which is fed to the drive fo~ rotating the gyrostabilized
- platform azimuthally. When developi_r_g this signal the on-board prism
_ takes up the position in whicy~ the perpendicular to it will be perpendicu-
lar to the vie~:~~~g axis of the goniometer.
The on-board control prism does not have a fixed position with respect
to the basic stabilization plane of the rocket and is fastened to the
- stabilized base of the gyropZatform in the gimbal which can rotate by
360� with respect to the gyroplatform. Varying the position of the
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prism with respect to the basic stabilization plane, it is possible to
change the launch direction with a fixed position of the viewing axis
of the goniometer.
In order to control the operation of the system, a special control unit
is used from which commands are given to the tracking reflector with
primary "lock-on" of the on-board prism by the guidance system. The
"lock-on" signal is generated in the goniometer and is f ed to the control
unit. In order to monitor the operation of the guidance system on
"lock-on" of the on-board prism and in the mismatch signal generation
regime the tracking system for the rotation of the gyroplatform is the
television with transmitting camera placed in the goniom~ter. The
_ receiving chamber is fed part of the ~ight mismatch signal generated by
the goniometer. Direct visual control of the accuracy of the guidance is
provided for in the goniometer, for whic.h part of the light flux from the
goniometer is fed to the viewer.
1
i -
? ~ 3 ~
4
~
f
li~ �S' .
?
\ /Z
' ~d
~ ? ' �
~
` 10 6
~
~
i ~ '
\
- , i i 7
\ `
9 ~
~ 9
~ 1
~
~ ~ _
~
~ t . 13 ~
\
Fi;ure 8.7. Single-channel guidance system:
1-- control prism; 2-- precession angle gauge; 3-- gyroscope;
- 4-- moment gauge; 5-- television receiver; 6-- television
transmitter; 7-- objective; 8-- reflector; 9-- prism;
10 azimuthal error signal amplifier; 11 control unit;
12 engine power booster; 13 autocollimation goniometer
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.
3 ~ ~I~
, 4 .
Z 8 . . .
1 ~ '
. ~ -
~
.
.
I \
.
! `
s~
I ~
i
.7 ~
,
d 5
. A~ ~
~
A ~ ~
~ o ~
C s
~ ~
dA ~
~
B C ~
Figure 8.8. Two-channel guidance system:
1-- control prism; 2-- moment gauge; 3-- gyroscope;
4-- precessiun angle gauge; 5-- orientation points;
6-- long-range goniometer; 7-- short-range goniometer;
8 amplifier; 9 drive
The two-channel guidance system (Fig 8.8) includes two autocollimation
goniometers and two tracking systems, one of which is used to rotate
the space rocket system and the other, to rotate the gyrostabilized
platform.
The goniometer of the first tracking system is installed in direct
proximity to the launcher and is viewed along the control prism attached
to the rotating part of the launcher. The mismatch signal generated by
the close range goniometer is fed to the drive for rotating the launch
system together with the rocket. The close range goniometer is designed
for rough guidance and provision for operation of the nonrange goniometer
and also for guidance of the space rocket system with a different
direction of launch, for which the launcher has two control prisms:
one corr.esponds to the guidance azimuth in the basic direction and the
other, the auxiliary direction.
The long-range goniometer which enters in to the precision guidance
tracking system is installed 130 to 150 meters from the launcher. The
light flux transmitted by this goniometer is directed at the on-board
control prism attached to the gyrostabilized platform. The mismatch
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signal generated by the goniometer goes after amplification to the drive
for rotating the gyroplatform which is rotated azimuthally until the
basic stabilization plane of the rocket coincides with the launch plane.
Before guidance the short and long range goniometers are installed so
that their viewing axes will coincide with the launch plane. For
orientation of it, the launch azimuth Ap and the azimuths of the oriented
geodetic direction A1 are used. The guidance angle (the angle between
the direction of~.the launch plane and the direction of the arientation
point)
~A = Ap - A1,
if this value is negative, it is increased by 360�.
A characteristic feature of the two-channel system is the relation between
the launch azi.muth of the space rocket system and the location point of
the goniometer which must be selected so that the direction of the viewing
angle of the goniometer on matching with the launch plane will simul-
taneously coincide with the direction of the control prism. This means
that the launch plane must pass through the axis of the launch system and
the location of the goniometer. If the launch direction changes, the
point of location of the goniometer must be shifted along with the arc of
a circle.
At the present time when building a launch complex and its structures,
the launch direction is taken into account. Here the launcher and the
goniometer are arranged so as to exclude preliminary (rough) guidance.
In addition, the modern control systems provide for a change in flight
direction by rotation of the control pris~ with respect to the gyroplat-
form.
In cases where the flight control system has two or three autonomous -
gyroplatforms in order to increase its reliability, the guidance system
also must ha~~a the corresponding number of independent azimuthal guidance
channels.
Special attention by the specialists is attracted by the guidance method
using gyroscopic compasses, the axis of which has selectivity with
respect to the direction of the north thanks to the effect of the direc-
tional moment manifested as a result of rotation of the earth.
The on-board gyroscopic instruments of the control system (accelerometers,
gyroscopes) can operate in the general compass mode. In this case the -
guidance becomes autonomous and the presence of oriented geodetic directions
and ground equipment at the launch complex is not required. The
def iciencies of this guidance method are the relatively long time (20 to
- 40 minutes) required to determine the directian of the north and the
technical complexity connected with obtaining the required accuracy
characteristics of the gyrocompasses.
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CHAPTER 9. MONITORING AND CONTROL SYSTEMS FOR TECHNOLOGICAL PROCESS
OPERATIONS
9.1. General Information
The automatic monitorng and control systems for technological operations
of preparing the space rocket system for launch at the engineering and
the launch complexes with respect to their organizational and functional -
characteristics belong to the large systems, for they contain a signifi-
cant number of different servo, power and measuring equipment and control
units connected to each other by braa.~hed, multifaceted co~unications
for automatic performance of a complex of functions under conditions of
complex environment in the presence of interference and counteracting
factors.
. The automatic control systems usually are classified by the information
_ about the control process or system used. The information plays a
significant role in the control processes, and the means of obtaining it
are important elements of the control systems. Two types of information
are distinguished: initial (a priori) and operating (arriving during the
process of performance of given functions by the system). Beginning
with the characteristics of the initial and operating information the
automatic control systems are divided into ordinary, adaptive and game.
The technological process of preparing the space rocket systems for launch
has a game nature. The problems of controlling the preparatir~n operations
can be interpreted as the problems of automatic playing of a game of two
sides, of which the first is the control system, and the second, the
object of control. The actions of the control system are subject to a
defined program within the limits of a number of solutions depending on
the action of the second side. The actions of the object of control are
also subject to certain rules, but there can also be random deviations
The object of control is not antagonistic with the control system. This
type of game system belongs to the class of games with nature.
_ 240
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In systems with a set of standard solutions from a set of versions of
actions of the f irst side, an optimal choice is made in advance, and
in systems with automatic search for the solutions, the control machine
itself solves the problem of optimal choice for each current step of the
control operations.
For example, let us consider the operation of fueling two tanks of a
- booster rocket (Fig 9.1). Depending on the order of receiving the signal
"nl" of passage of the level in tank A or B, valve 2K or 3K must be
- shut off; the sequence of shutting off the valves depends on the random
causes leading to various combinations of passage of the level. An
analogous situation occurs also with respect to the "Yp" signal in
tanks A and B with valves 4K and 5K.
~ ~:c'~ (b)
n~~. ~
if"' ~ ~
_ ni 8~
. ~K ~~4~K1
2/f ~ ~
6aKA (a) 14K
np.
e ~/1 I
m 3~
. ~K
� sx
9K lOK 6K
Figure 9.1. Pneumohydraulic system for filling the tanks of a
booster rocket
Key:
a-- Tank A; b-- Tank B; c-- (DPK) drainage safety valve
In case of failure of one of the valves 2K, 3K, 4K, 5K (failure to close)
the filling is stopped and transfer of the fuel components from the tank
- with the emergency to the other tank and f illing of it to the required
level begins and also topping off the tank with the emergency under
special conditions. Depending on which valve has failed, four different
combinations of operation of the systems are possible; consequently, only
for the given simplest system do eight game situations occur, and the
entire technological process of preparation has significantly of them.
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The control system for pre-launch preparation, beginning with the game
nature of the technological process would be expediently carried out
with automatic search for the optimal solution, but in this case it will
become very complex and awkward.
The insufficient reliability of the control elements forces us to do awai
with tha embodiment of all possibilities of game nature of the preparation
process, and one or several standard solutions, the optimalness of which
has been checked in advance, are written into its operating program.
The operation of the automatic launch preparation system (ASPS) consists
of successive steps, as a result of which the preparation control takes
place discretely by formation of a sequence of control instructions
~ for the systems of the space rocket system and the space center.
The automated launch preparahion systems must maintain fitness with respect
to their purpose and eff ectiveness for any action on the part of the object
of control, and by this attribute they belong to systems with minimum
necessary information about the target (at the beginning of operation
they only have minimum primary information based on the result of the pre-
ceding adjustment operations or checks, and minimum initial information
about the state of the systems, the failure of which can lead to emergencies).
Thus, for example, in the system for controlling the fueling of booster
rockets with fuel components, the initial information is the readiness of
the power supply, air supply and fuel storage systems, and as a rule,
the initial state of the level control system in the booster rocket tanks,
although this process is participated in by the set of electropneumatic
valves, pumps, a large number of elements in the automatic control system
itself, the pnewnohydraulic system of the booster rocket, the information
about the fitness of which is not available at the beginning of servicing.
The operating information obtained during the control process about the
state of the object of control comes to the control machine (system). In
the ASPS, the game algorithms of the operation are directly put in the
special modules (or autonomous systems) which control the individual
technological processes in the form of a defined set of standard solutions.
The modules themselves (autonomous systems) interact with each other by
a program given in advance.
Beginning with the investigated peculiarities of the operation of the
ASPS, they are classified as game systems with program control and the
set of standard solutions (Fig 9.2). The most characteristic feature of
such systems is the use of the control instructions obtained from the
operating information on the basis of the algorithms. In the ASPS, there
is no "struggle" of two or more algorithms in the operating process, but
a"struggle" of the algorithm with the random disturbing factors. The
criterion on the basis of which the various versions of the algorithms
are compared usually can be expressed in the form of the basic function
of state of the operation, the so-called "payoff function" and additional
~ 242 .
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conditions. The choice of an effic ient "payoff function" for the operation
of the ASPS is the most important p art of the development of the alogrithm
and requires profound study of the technological processes of launch prep-
aration considering all factors, circumstances and relations existing under
actual conditions. When designing the ASPS for this purpose the results
of experimental and full-scale tests are used.
~ CucmeMai
Qemo,+.rumtt yecKOZo
yn pa e neyuA (1)
O66JKHOBBNNb/C CQMONQC1Il/1QflCQA~U{lfCCA
~2) \ ~~3~
/1~/10/l63yA7- R730MKNy- 3A'C/ll MQ/Ib CCQMONQC/11-
W~e moie cucme- ~ ~OQUOLLpI[(l!M CQMONQC//l-
npuyqunb~ Mai aemon+a pe
y~_ Roppr.xmupy p~uearou~u-
om~r~aveHas mavecKnao eaHU~ '~!'M 1'~m' ec~v
4 y/J,QIQB/ICN!!A C6, 7 d/bOM $
l~)
~ .
Nspoebie ~
1 _ _
r /1~0'P~QMMNOZO ~
~ C HQ60~OOM I yR~B/1BNUA I C QO/IIOMQ/AUVBCK(lA~l
/UQ6/IOHHb/X� ~ ~ yQ6p~p,~ ~ qOUCIt'OM
PC!!!BN(llL I UlQ6/IONNb/X ` ~OiCUlBNU(L
10 ~ , L Peuietii~u ~1~.1 12
Figure 9.2. Classification of automatic control systems
Key:
3. Automatic control systems 10. With a set of standard
2. Ordinary sclutions
3. Adaptive 11. P~bgram control with a set
4. Using the principles of deviation of standard solutions
5. Open automatic control systems 12. With automatic search for
6. Experimental regulation the solution
7. With adaptive correction device . ~
8. Adapi ive
9. Game ,
The automatic monitoring systems in the general theory are classified
with respect to the most varied attributes, for example, with respect to
the type of controlled variables, the purpose, sphere of application,
technical execution, and so on. However, classification by these
attributes has practical significance which would be common to all the
automatic monitoring syst~ms and which would characterize their internal
structure and functional peculiarities (Fig 9.3).
~ 243
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CucmeMei -
acmo~+tam~rvec~roro
~rvamponA ~ )
1
C OBNOA'pQ/ANb/M C MNOtOA'PQ/l1N6/A/
UC/!0/lb300QNUBM UC/!0/16JOOQN!/C~i/
ycmparicme ycmpodcme
- ~ . RQNL7AYl IrQHQ/1Q
KON/ll~plpr/A 2 KoympoAR3 `
~ 1
C HC/1pICQb/eNb/iiI C a!!CA'QB//INb/M ri U.l,NCpL'N!/CA!
npedcmoe~uruc~?i n,oedcmaeneHUe,:~
Koympoeupyr.~ax Koym,oonupye~fax
napa~rempio~ napa,?~empoe
u ycmQOn~r ~4~ ycmaeoK~S~ 6ea ~~weFe~r
C ucnnaa~aeovNVe,v
Cna,orrn~enayoiM Cnocn~oaamenayai~ 'n~~yu�em
conocmQeAe~rue.~r canocmQene~uea.~ cvna~cmae~,eauA(10
Koym,vonUpye,rrnzo ~roNm,oonupyeMOZo
nQpu,rrempa napuMemAcr
c eao ycinao~qq~te , c eto y~mae~ra.~u~9~ . 6e,~ ucnoAOSOeaNa.v
pea ~amomoe
nAeo~adyu~em (1
conocmae~eau.v
C e~daved C oadave~i
pe~yAamQm~ pesyabmamae �
KON/IIpO/lA KON/lIpOAA
� /10 A'QAYO~OMy r0 COOOKy/INOC/JJlL
KoympoAUpyeMVNy noympv~vpye.~ax
naAQMempy 12 ~Q,oia'n~emp~ve. 13 ) .
Figure 9.3. Classification of automatic monitoring systems
Key:
1. Automatic monitoring system
2. With single use of the monitoring channel devices
3. With multiple use of the monitoring channel devices
4. With continuous representation of the monitored parameters and
settings
5. With discrete representation of the monitored parameters and
settin~s
6. With measurement
7. Without measurement
8. With parallel resistance of the monitored parameter with its
settings
9. With series comparison of the monitored parameter with its settings
10. With the use of the results of the preceding comparison
11. Without use of the results of the preceding comparison
12. With output of the monitoring results with respect to each
controlled parameter
13. With output of the monitoring results with respect to the
� set of controlle~.'. parameters
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The process of monitoring the pre-launch preparation of the space rocket
complex can be realized by a sysstem of any one type, for each technologi- -
cal process imposes its requirements on the system. Even for one
technological process certain parameters must be monitored continuously; _
others must be monitored discretely, and a third set, only during the
process of deviations from the given values. Therefore, the exiGting
classifications do not reflect the functional purpose of the systems for
monitoring the technological operations of pre-launch preparation.
9.2. Purpose of the Systems
The systems for monitoring and controlling technological process opera-
tions ~re a set of equipment designed for the performance of technological
= operations of prrI.aunch preparation of the space rocket system and also
monitoring its state and the state of the ground system during the
preparation process.
During the preparation of the booster rocket for launch, control reduces
to the performance of various operations causing def ined, previously
_ provid~d for process conditions. These conditions are repeated under the
given conditions always in the same form. In addition, the control sys-
t~ms maintain a constant value of the regulatable variables with respect
to a given law, and they protect the complex from emergency situations
_ on occurrence of uncounteci operating conditions of both the ground systems
and the booster rocket systems. For example, if for any reason the
launch of a booster rocket filled with cryogenic fuel components (liquid
oxygen or hydrogen) is delayed, then heating of the fu.el begins and, as
a consequence, an increase in volume and level in the tanks, which can
cause an emergency. In this case the level and temperature gauges located
in the booster rocket tanks generate emergency signals for t:~e ground
~ service control system which provides for correction of the level and
*hermostating of the fuel in the tanks.
_ The monitoring and control are continuously related to each other. The -
objects of mcnitoring are not only the Lechnological systems, but also
the control systems themselves and even the space rocket system.
The purposes of ~monitoring and preparir~g the sFace rocket systems at the
~ engineering and launch complexes can be the fol~.owing:
Output of informa.tion to the operator about the condition of the space -
_ rocket system, the technological process systems and their operating
_ conditions in the process of pre-launch pregaration; -
Obtaining information about the state of the object for variation of the "
control conditions or generation of the required contral input;
_ Correctness of the execution by the control system of the process
algorithm and also correspondence of the system parameters to the given
values; ~ `
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Determination of the location an~i causes of failure of the control system
in case of its failure;
Mechanical and electrical failure, seal and reliability of fastening the
elements and modules;
The state of good repair of the control system in storage, without repro-
duction of its actual operating conditions.
As a rule, in order to perform each technological operation there are
special monitoring and control systems which, depending on the degree of
automation are divided into manual, semiautomatic and automatic. ~
The manual monitoring control system is a control system which requires
the participation of an operator. Estimation of the monitoring results
and observation of the defined sequence in the output of the control
commands are also the business of ~he operator.
The semiautomatic monitoring and control system is characterized by the
fact that the main part of the operations are performed automatically.
_ The operator only switches the individual monitoring and control elements
on and off, but he cannot introduce changes into ttie process of execution
of the cycle and its sequence. When wc~rking with such systems the operator
usually ma.nually controls more than 50% of all of the operating time of
the systems.
Automatic monitoring and control systems do not require operator interven-
- tion except to switch the systems to a given regime and individual manual
operations as a rule amount to less than 2% of the total operating time.
The selection of the operation, the control and the decision making of
such systems are all automatic.
The complexity of the space rocket systems, the large number of operations,
the limited test time and the performance of the launch at a previously
established time give rise to the necessity for maximum automation of the
pre-launch preparation process. This reduces the preparation time for
the launch, it increases the accuracy and reliability of monitoring, it
permits operations to be performed which cannot be performec~. by man on the _
basis of his limited capabilities, it decreases the wear of the equipment
and also essentially reduces the service personnel.
Beginning with the necessity for complex solution of the control problems,
- it is expedient to develop a united system including the entire complex
of automated control devices for individual units and systems which -
participate in the pre-launch preparation. Such a system includes both
the control systems and the systems for monitoring Che general engineering
and special technological process ground systems. On campletion of the
operation, it outputs the general availability to the ground equipment
- of the engine starting control system.
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- In individual cases, for very large complexes it is expedient to crF_ate
, not an automatic, but automated preparation system, for a number of
operations, especially decision making, are more efficiently performed
by the human operator.
The control of the ASPS units and systems is realized from a central
preparation panel (TsPP) designed to monitor the operation of the pre-
launch preparation operations control system and locate it at the command _
post of the launch complex. Usually all of the information about the
course of the performance of the pre-launch preparation operations is -
concentrated at the TsPP. The signa.ls and commands from the central
preparation panel go in generalized form to the launch panel which is
located. at the command post of the space center or in the launch control
center building.
A flow chart for the control of the technological operations of launch
preparation of space rocket systems is sho�~an in Fig 9.4.
~ 1~ PaKemNO - xocMUyec~ras cucme.?~a ' .
' (2) TexNOnotuvuKUe cucme.+~ai na&nmoeKu
N�1 N~f N+d N�4 . N~n
_ . � ~
N!1 N~1 N~~ N~4 . , N~I1. ~ -
' ~ g~ CpcmeMai //npae~eNU,v u ~roympcNA _
0/AdC96Nb/X /liCXHO~IOtuyCCA'UX C(/C/7l
~t,~ acnc . .
~ . ~5~ q~~ ~ - . , : ~
/ly~em
. ' (6)ny~c,rQ -
Figure 9.4. Flow chart of the control of the technological
operations of launch preparation
Key: 1-- space rocket system; 2-- technological preparation systems;
3-- monitoring and control systems for individual technological
systems; 4-- ASP S; 5- TsPP; 6- launch panel
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The command to p~rform the f inal operations of pre-launch preparation
are fed from the launch panel, the subsequent hurning of the transparencies
on which permits monitoring of the course of their performance. By the
command of the operations director the operator sets the LAUNCH SWITCH
- to the required position and presses the LAUNCH button, after which the
pneumatic valves of the drain lines of the oxidant and fuel tanka are
closed, the fuel tanks are blown, the engine is started, and on reaching -
a def ined thrust the space rocket system separates from the launch system.
1 9.3. Classif ication of Systems
The most specific attributes of the ASPS are the principle of their con-
struction and monitori.ng techniques (Fig 9.5).
~ _
.
~ Aemo,~ramuvecKUe
; cucmeMb~ ,
_ iroBiomaa~ru pm� .
;
' C B!!C/71QX((!!64'NNM C /lI~IC.MCXQN!/ -
i yp~~ fCCAV~/
3)Y~,a~nrrw
~
; .
; C~M,rqu ara~w C ap~,~nxaaa C~yy,rquoNO.?e~ro- -
: ~r~ 4 ~ro
m~(5) ~
~ro~wm~a ~6~
~
l
~ - np/1Moaa
; ' C L7~//0/1~irDYl ~
hlOUA H `
~ 7 ' J .
~
~
` C /lpqd~lTpl/IAGANaY% C COMO/lQAa!',OAGY!
$ c~,�a+q~paai 9~ no aeap~vu. ~10~' : �
~
~
Figure 9.5. Classif ication of automatic launch preparation
systems (ASPA) by the principle of consCruction
and method of control
Key:
- 1. Automa.tic launch preparation system
2. With re~note control
3. With telemechanical control
4. With the functional monitoring technique
; 5. With time monitoring technique
i 6. With functiona.l monitoring technique
7. With self-checking
; 8. Direct action (without self-checking)
~ 9. With preliminary self-checking
~ 10. With self-checking on emergency
l
1
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The automated control systems for the pre-launch preparation of space
rocket complexes can be executed both in the form of a remote control
system and a telemechanical control system. The former are characterized
by the fact that the control of the monitoring and adjustment equipment
take place over wire communication channels, and the equipment is at a
comparatively short distance from the ob~ect of preparation (to 500 meters)
and it is placed at the command post of the launch complex, and at the
servoelements of the technological process systems. In the second group
the control of the monitoring and regulation equipment takes place over
the telemechanical channels. The command post, as a rule, is located at
a significant distance from the technological preparation and launch
systems.
With respect to principle of construction the automatic moni.toring systems
are divided into systems with fux~ctional, time and functional-time
monitoring techniques.
For the ASPS with functional monitoring technique the sequence of trans-
mission of the commands and the beginning of operation of the individual
systems are related to each other by strong functional relation: each
subsequent command can be generated only after monitoring the execution
of the preceding one. If for any reason there is no sign3l of completion
of the operation, then the next command is not output, and a transparency
burns on the operator panel signalling emergency shutdown of the process.
The ASPS with the time method of monitoring are distinguished by the fact
that the next command will be issued after a strictly def ined time follow-
ing the preceding one, in spite of the fact that the precedirg command
has already been executed and all the conditions for executing the next
operation have been met. If the preceding command has not been exec~~ted
and the conditions have not been set up for execution of the next command,
then automatic disconnection of the system and return to a gafe (initial)
position are provided for.
For the ASPS with functional-time methods of monitoring it is characteristic
that there is a rigid functional relation between the control commands
which, in addition, is time controlled.
All of the investigated types of ASPS can be executed with self-checking
and without self-checking. Self-checking is realized either before the
beginning of operation of the system or during operation by the monitoring
system signals. In the first case on the "preparation" command initially -
the equipment self-checks, and after receiving a positive result, per-
mission is given for further operations; in the second case, the system
_ halts its operating cycle and begins a self-check to find and indicate
the location of the failure. The systems with self-checking are more
complicated and expensive, but this is compensated for by the convenience
of their operation. When preparing to launch space rocket systems and
especially in emergencies when the decision-making time is strictly
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limited, the systems with self-checking offer the possibility of quickly
finding the failure and prevention of abortion of the launch or an
emergency.
The automatic launch preparation system usually includes a number of
individual systems executing functionally different missions; such systems
have come to be called subsystems (Fig 9.6). The ASPS subsystems can
operate both in the pre-launch preparation process and in the period
between launches (the so-called duty period).
The systems operating during the duty period mai.ntain the technological
systems of the launch complex in a state of readiness for reception of
the space rocket system. These launch complex systems include the storage
areas for the fuel components, the receivers, compressor stations, the
~ power feed systems, thermostating systems and so on.
The duty systems can be divided into the systems with cyclic effect,
systems operating by deviation of the regulatable variable from the given
racing, and tlie systems with combined effect.
The cyclic action systems operate not during the entire duty time, but
periodically, with defined cyclicity (once a day, every hour, and so on);
here all of the measurement, regulation and control processes in them
are performed only during the operating cycle. Such systems as a rule
are used for technological processes having high inertia in which the
failure of the individual elements cannot lead to emergency during the
period between the monitoring systems. Thus, the temperature in the level
of the fuel components in the storage tanks are usually monitored once a
day, for their vaxiation takes place slowly, and even in the case of
failure of the thermostating means, tens of hours are required for them to
gobeyond the admissible limits.
The systems operating by the deviation of a regulatable variable begin
to function only when this variable goes beyond the admissible limits,
after which it will be monitored continuously. Such systems are used
when the deviations of the given variable from the rated value can lead
to an emergency. Thus, for example, the vacuum insulation of the liquid
hydrogen storages is monitored, for on loss of seal, an explosion-hazardous
mixture can be formed, and an explosion can occur.
The combination-action systems combine both periodicity of action and the
principle of beginning operation on deviation of the regulatable variable
_ from the given rated value.
The systems for monitoring the technological operations of launch prepara-
tion are divided by purpose into systems for functional and operative
monitoring and systems for monitoring the control process (Fig 9.7). :
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AemoMamuve~rue
cucme~e~
nadzomoeKU cmQpma ~1)
~~y~ ~ 2 ~ Peay~upooaHUA ynpoe~e~mv
cur~vu,/u~aqau
Qe~rypyom ' /lyc~raieruto
pearaav~5~ p~,~Q ~6~ .
qu~rauvec~roao ~o om~r~oNeyuro Ko,~OuyuAv-
~ ~zynupyeMOU eaHHOto ~
~~j~ aenuvaya~g dedemet~v(9
Figure 9.6. Classification of the ASPS subsystems
Key:
1. Automatic launch preparation 6. Launch regi.me
systems 7. Cyclic effect
2. Measurements and signal units 8. On deviation of the
3. Regulation regulatable variable -
4. Control 9. Combined effect
5. Duty regime ~
The functional monitoring systems provide informa.tion about the state of
the object for the generation of def ined control inputs conditioned by
the technological process algorithm. These systems usually measure the
physical parameters (temperature, level, pressure, vacuum, flow rate, -
displacement, and so on). The interaction of monitoring systems with the
control system takes place automatically and is determined only by the
technological preparation process.
The operative monitoring systems provide inforwation about the state of
the technological launch complex systems anii all the systems of the space
rocket system. As a rule, these are several multichannel systems capable
of recording and monitoring parameters (from several tens to several
hundreds) determining the degree of readiriess of the space rocket system
for launch, the temperature and pressure in the various compartments of
the booster rocket, the condition of tHe pneumatic and hydraulic equipment
of the engine, the seal of the instrument compartment and the operation
_ of the on-board electrical systems. Considering the large volume of
parameters which must be encompassed by visual observation and also the
necessity for document recording of the preparation process, the majority
of the operative monitoring systems are designed considering the recording
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C(LC/17BMb/ KON/%Ip0/IA �
/11eXN0.I0EUSlBCKflX (1)
onepauuu
~Aedcmapmoeoci nodiomoeK~
~yy,rquoycrna- u~epumueHOao KoympvnA
Hoto ,roNmpoAn ~3 ,~p~aqecca
~roNmQo~~2~ ynpcrane~uA 4~
euayrraayaa ao~ry"'e"- Ta,rmoeaao navmanNae `
~m~~A manenoao ,r~Mmp~onA ~ronmpam~
Koym ~r ~ ~8
Figure 9.7. Classification of systems for monitoring the
technological process operations of pre-launch
preparation
Key:
1~ Systems for monitoring the 5. Visual monitoring
technological operations of 6. Document monitoring
pre-launch preparation 7. Cycle monitoring
2. Functional monitoring 8. Step by step monitoring
3. Operative control
4. Monitoring of the control
_ process
_ of the operating process on photographic film, paper or magnetic tape.
In these systems, along with the recording equipment preparation is made
for the possibility of visual monitoring of the interesting parameter
as the operator desires. The visual monitoring systems are primarily
used to observe the operation of the ground technological systems or for
monitoring the auxiliary parameters of the space rocket system during the
development and f irst flight testing of them.
The control process monitoring systems provide information about the
correctness of execution of the given algorithm and they form a signal
- to stop preparation in case of emergencies. These systems usually are
divided into two independent types: the cycle monitoring systems which
monitor each discrete change 1.n state of the ob~ect and the control
system and by the monitoring results,~permitting or forbidning subsequent
operations, and the step-by-step monitoring systems which monitor a
defined completely technological step cycle including part of the
- ' overall technological process. The step-by-step monitoring is used in
cases where the technological process can be broken down into individual
steps and there is a possibility of halting the process or repeating it.
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9.4. Estimation of the Eff iciency of the ASPS
Ttr~ ef f iciency of the technical device is the cY~sracteristic of the degree
- and quality of performance by this device of the functions for the execu- -
tion of which it is intended. Consequently, for the automatic launch
preparation system the efficiency criterion will be the probability of
the execution of technological preparation algorithm with estimation of -
the basic pa?-ameters of the monitoring and control systems for the tech-
nological operations by their criteria in difFerent stages of their
development.
, The parameters of the monitoring and control systems most significantly
inf luencing the structure and the direction of the developments usually
are estima.ted in the stage of preli.minary design with simultaneous
selection of the control principles, the construction of the structural
diagram of the system and its elements, determination ~f the structure
of the subsystems, and so on.
~ The variation of the relative number of estimates in different stages of
development and their importance are illustrated in Fig 9.8. As is
obvious from the f igure, the frequency of the estimates in the detailed
design stage increa.ses, and their importance is reduced. The cost of
redoing the system caused by the i.mplementatian of suggestions increases
as it is developed. _
The concept of efficiency of the systems can include various components
which reflect the time and cost of development, the cost of manufacture
and servicing, the degree of realization of the basic specifications
of the system. The efficiency of monitoring control systems for
technological operations at launch can be estimated by the procedure
constructed on the basis of the model of estimating the efficiency of
such systems where the model of the eff iciency is depicted in the form
of a graph not containing loops (Fig 9.9) and having three branches:
readiness, reliability and compatibility.
The readiness branch for the space rocket complex Pr=1, for the prepara-
tion and operation of the space rocket system take place in a previously
def ined time and there is no necessity for keeping it constantly in a
ready condition. For the monitoring control systems for the preparation
of space rocket complexes, the basic criterion probability of insurance
of the launch at the given time
Pe - ~H ' Pc ~
where Pe is the eff iciency system (the probability of insuring the launch
in the given time);
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. �
~A , ~ ' .
t ' 3 -
,
, � b, : c~ . e ~ ~~d (B)
Figure 9.8. Graph of the variations of the relative number
of estimates in different development stages:
a-- selection of basic parameters of the system; b-- pre-
liminary designs; c-- detailed design; d-- manufacture;
e-- test; f-- operation; 1-- relative importance of the
estimate; 2-- relative frequency of the estimate;
3-- relative cost and delay connected with changes
Key:
A. estimates
B. operating periods
PH probability that the system will aperate for a given period of time
insuring the characteristics within the tolerance limit; ~
P~ compatibility def ined by the probability that the actual conditions
of application will correspond to the conditions under which the control
system carries out its mission.
This equation is valid if PH and P~ are independent.
The reliability PH determines the probability that the ASPS systems will
function without the parameters within the tolerance limits for a given
_ time. Consequently, -
pH ~t~ = e t/Tmean,
where t is the given time;
Tmean is the mean fail-safe operating time.
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$
P~ ~ , . ?'c
. ~ . PH ~ .
Figure 9.9. Graph of the efficiency model:
Pe efficiency of the system (probability of insuring a launch
_ at the given time); Pr readiness (probability that the system
will be rea.dy at the required point in time); P~ compatibility
(probability that the actual operating conditions will correspond
to the conditions under which the system will carry out its
mission); Pg reliability (the probability that the system will
operate for the given period of time insuring the output
characteristics within the tolerance limits).
This equation is valid if adjustment and repair operations are not made
during the process of regular functioning of the system.
For the ASPS in the modular synthesis step (the drawing design) the goals
of the system are defined, the systems are broken down into individual
modules, the general plan is made for the exchange of information in
commands between the systems and modules. The general reliability of the
performance of the stated goals between individual systems and subsystems
of the ASPS is also distributed in this step.
The equal reliability of all systems (modules) means identical basing when
distributing the general reliability among the individual systems or modules.
This distribution is not uniform, for as a result of the difference in
functional problems and their complexity it is impossible to insure identi-
cal reliability of all systems in practice. For systems carrying out
- simple functional missions and having few component elements, it is simpler
to obtain high reliability than for systems with a large number of elements.
Therefore the overall reliability of the ASPS between individual systems
is distributed differentially, for which the concept of the provisional
"weight" of the system is introduced
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B = aN (~+K),
where N is the number of servoelements which the system controls (for the
- monitoring syste:n, the number of monitored parameters);
~ is the number of functional operations performed by the system;
k is the number of monitored states or operating conditions of the system
which lead to emergency shutdown of the technological operations or a
change in operating conditions of the control system;
a is the coefficient reflecting the importance of the individual system
or module overall; usually a=1, but for especially important systems or -
modules, the reliabi]ity of which must be apprec~ably higher than in the
remaining systems, 0