AGARDograph 15 r- 4 1 r > * ► pects int is, Fundamental Aspects of Solid Propellant Rockets F.A.WILLIAMS, M. BARRERE nd .C . HUANG Reproduced y he CLEARINGHOSE for ederal cientific Technical Information pringfield Va 22151 N d
Rockets
F.A.WILLIAMS, M. B A R R E R E nd .C . H U
A N G
Reproduced y he
Research and Development AG/RD)
Atlantic reaty rganization nd
mittees from N A T O
countries. hese
panels are responsible fo r
sponsoring
technical meetings and symposia and
th
e publication of technical papers.
Part f h e G A R D ublishing
programme akes he orm of ood
quality etterpress r itho, ard
cover volumes of symposia proceed-
ings r onographs n ignificant
areas f eronautical nd ero-
nautical research and development.
Fundamental Aspects of
Solid ropellant Rockets
Copyright
Research and Development, ATO
Standard Book Number 85102.016. X
Plate 2 . he manufacture of tw
o first tage ocket motor chambers or
the ubmarine-launched Polaris A-3 ballistic missile.
The
whirling arm of the machine in he ear s applying circum-
ferential wraps of glass ilaments over he mandrel while he
machine in ne foreground is applying the final oo
p wraps.
Following wrapping, he chamber is placed n n oven where
the esin mpregnated filaments become
homogeneous
structure.
The mandrel s then withdrawn prior o casting the
chamber with propellam;
courtesy Aerojet General Corporation.
An Introduction to Solid Pro pelU
ni Rocket Motors Generalities
1
2.
Description of Current Solid PropeUant Rocket«
2. 2 History
3.1
Rocket Motors or Lower Stages Boosters)
3 . 2
Rocket Motors or Upper Stages an
d for Space Vehicles
3.3 Satellite Launchers
3 . 4 Sounding Rockets
3 . 5
Auxiliary Rocket Motors to perform Guidance and Control Functions
3 . 6
Assisted Take-off Rocket Motors for Aircraft
3 . 7 Military Applications
3 . 8 Diverse Applications
Chapter 2
1.1 Introduction
1.2.1 ass Conservation
1.2.2 omentum Conservation
1.3 Isentropic Flow
1. 3 . 2
One-component deal Gases with Constant Heat Capacities
1. 4 Nozzle Flow
1. 4 . 2
Flow in deLaval Nozzles
1.4.3 hock W a/es
1. 4 . 5 Nozzle Flow Formulas
1. 5
Thrust and Rocket Performance Parameters
1. 5.1
Derivation of Thrust Formula
1. 5 . 2
Theoretical Thrust Formula;
Maximum Thrust
1. 5 . 3 Thrust Coefficient
1.5.4 Characteristic Velocity
1. 5 . 6 Other Performance an
d Design Parameters
2 ffects of Multicomponent, Reacting Gas Flow
2.1 Introduction
2. 2.1 Frozen Flow
2. 2.3
Comparison of Performance with Equilibrium an
d Frozen
Nozzle Flow
3 . 2.1
Two-Phase Flow Without Particle Lag
3 . 2. 2
Equations f Two-Phase Flow with Particle Lags
3 . 2.3
Dimension less Lag Parameter
3 . 2. 4 Large Lag Limit
>
3 . 2. 5
Small Lag Limit T<<
3 . 2. 6
Numerical Calculations for ntermediate Values of
3 . 2. 7
Influence cf Phenomena Neglected in the Theory of
Section 3.2.2
3 . 4 Experimental Results
4 ozzle Heat Transfer
4 . 4
Additional Heat Transfer Considerations
5 iscussions of Other Deviations from deality
5.1
Influence of Non-One-Dimensional Flow on Nozzle Performance
and Design
5 . 3 Jet Detachment
5.4 nteraction of the External Exhaust Jet with its Surroundings
6 hrust Vector Control
6.3 Fluid Injection
6.3. 2 Theoretical Analyses
6 . 3 . 4
Choice of Injectant Fluid
6.3. 5 H ot Gas Valves
6 . 4 Comparison of Mechanical an
d Fluid Injection Tmus
v
* « * ' >r Control
7 ompatibility of Solid- Propellant Motors *ith New Nozzle Design Concepts
Chapter 3
1. 2
Composition of Propellants and of Products of Combustion
1. 2.1
Propellants Homogeneous Propellants
Heterogeneous Propellants
(a ) Oxidizers
(b ) uels
2 heoretical Performance Calculations
2.1 Equilibrium Composition Equations
2.1.1 Definition of Basis
2. 2.1 Gaseous Combustion Products
2. 2. 2
Combustion Products Containing a Condensed Material
2. 3
General Methods or Calculating Equilibrium Compositions
2.3.1 Huff Method
2.3.3 Brink ley Method
2. 4
Examples f Applications o Propellants Containing H, i, Be, B, ,
Ai, , l, an d F
2. 5
Evaluation of Propellant Performance
2. 5.1
Thermodynamics of Equilibrium Mixtures
2. 5 . 2
Calculation of Performance Adiabatic Flame Temperature
General Aspects f Performance Calculation
Performance for Frozen No Tie Flow
Performance for Equilibrium Nr zzle Flow
Influence Coefficients
2. 6
Research on Solid Propellant Constituents Conducive o Maximal
Performance
2. 6 . 2 Studies of Binders
2.6.3 Studies f Oxidizers
2. 6 . 5 General Remarks
3 xperimental Determination of Performance
3.1 Laboratory Methods
S. 2 Rocket Motor Tests
3 . 2.1 Conventional Test Stands
3 . 2. 2 Measurement of the G
as Velocity n th e Motor
4 omparison of Theory nd Experiment
Chapter 85
Motor Operation
1 ntroduction
2 urning Rate Laws
2.1
Laws Proposed for he Burning Rate Influence of Certain Parameters)
2. 2
Techniques or Burning Rate Measurements
3 otor Operating Characteristics Time Evolution of the Pressure an
d of
th e ¥/eb
3.3 Propellant Grain Geometry
c) egmented Grains
d) lotted Grains
f) Helical Grains
3. 5
Dual-Compositior Solid Propellan"
3 . 6
Remarks on the Pressure-Time Curve obtained during a Motor Firing
4 ptimization of th
e Motor Geometry in Particular Cases
5 jlid Sublimation Motors
1 ntroduction
2,
Experimental Techniques for Studying Homogeneous «olid Propellants
2. 2 Experimental Results
Components of Heterogeneous Propellants
3.1 Linear Pyrolysis Rates
3 .1„ 3
Importance of Pyrolysis Measurements
3 . 2
Deflagration Rates of Certain Oxidizers
4 roposed Model Experiments for the Analysis of the Mechanism of
Heterogeneous Solid Propellant Combustion
4.2 orous-Core Burner
4
. 4 ressed Solid Propellant Strands
4 . 5 Combustion of Metals
5 tudies of Combustion Mechanisms with Heterogeneous Propellants
5.1 Direct Methods
5 . 2.
Moderate Pressure Domain 5 to 0 aim)
5.2.3 lateau Domain
5 . 3
Summary of Combustion Domains
Chapter 3 3
1 asic Equations of Aerothermochemistry
1.1 Introduction
1.3
Integral Form of the Governing Equations
1, 4
Differential Form f the Governing Equations
1.5 ransport Phenomena;
Reaction Rates
1.6 herrnodynamic Relations; Counting of Variables
1.7 iservation Conditions at an nterface
2 heories of Homogeneous Solid Propellant Combustion
2. Adiabatic Theories
2.1.1 History
2. 1. 2 Theories f Rice an
d Ginell an d of Parr an
d Crawford
2, 1. 3
Theories f Johnson and Nachbar snd of Spaltung
2. 1.3
Definition of the ohnson-Nachbar Model
2. 1.3. 2
Basic Equations Governing the Gas-Phase Problem
2. 1.3. 3
Bovndary Conditions for he Gas-Phase Problem
2.1.3. 4
Diniensionless Mathematical Formulation of he Gas-Phase
Problem
2.1.3, 5
Bounds or the Solution o the Gas-Phase Problem
2.1.3. 6
Solution of the Gas-Phase
2.1. 3.7
The Surface Gasification Process
2.1. 3.10
Intermediate Surface Boundary Conditions
2.1. 3.11
The Pressure Dependence of th
e Burning Rate for
Unopposed Surface Gasification Process
2. i. 3.12
The ohnson -Nachbar Results for the Adiabatic Burning
Rate of Ammonium Perchlorate
2.1.3. 13 T
he Pressure Dependence of the Burning Rate for Surface
Equilibrium
2. 2 Nonadiabatic Theories
2. 2. 3
Energy Conservation Equations, ncluding Heat Losses
2, 2. 4 T
he Origin of the Influence of Heat Loss on the Burning Rate
2. 2. 5
Dependence of Heat Loss on Surface Temperature
2. 2. 6
The Modifications Produced in the Burning Rate Analysis y
Nonadiabaticity
2.2.8 Interpretation of
Double Eigenvalue Solution
2. 2. 9
Comparison of the Nonadiabatic Theory of Johnson an
d Nachbar
with Experiment
3 heories f the Decomposition of Selected Constituents of Composite
Propellants
3 . 2.1 Porous Plate
3.3.1 uel Constituents
3.3.2 mmonium Nitrate
3.3.3 mmonium Perchlorate
4.1 Introduction
4.3 deas Concerning the Interplay of Diffusion Flamee and Premixed
Flames
5.1 Introduction
5 . 2 Description an
d Classification of Behavior
Various Burning Metals
5 . 3
Theories of the Combustion of Spheres f Metals with Nonvolatile,
Insoluble Oxides
5 . 3 . 2
Metal Sphere Combustion
5.3.3
Remarks Concerning Assumptions for an Improved Theoretical
Treatment of Aluminium-Sphere Combustion
a. Method f Marklund nd Lake
h. Zucrow's method
d. eneral remarks on the Labor?»tor> Methods
2. 2
Direct Measurement of Erosive Burning Velocity in Rocket Motors
a. urn Interruption Technique
c. robe Techi "es
2. 4 Expert mental Results
2. 4.1 General Description
a. nfluence of the Gas-flow
b. nfluence of the Nature of the Propellant and of Graii?
Geometry
d. ummary
b. heory of Vandenkherckhove
3 . 2
Aerothermochemical Approach to the Problem of Erosive Burning
4 ffect of Erosion Phenomena on the Geometry of the Central Port
Ciapter 8
1 Ignition
1,1 Introduction
1.3 xperimental Results
1.4.1 Introduction
1.4. 2
Ignition by Means of a Stagnant Ho^ Gas
1.4.3
Ignition by Means of a Flowing H
ot Gas
Analysis
Discussion
1. 4.4
Ignition Processes nvolving Heterogeneous Reactions and
Radiant Energy Transfer
Solution for Small Values of t
Numerical Solution for Hypergolic Ignition w hout Radiant
Flux
Flux
Solution and Results for Radiant Ignition without Surface
Reacts (is
Hypergolicity
Generalizations
1.4.6
Critical Comparison of Existing Theoretical Studies
Scale Effects
Motor Extinction by Depressurization
Extinction by Water Injection
Conclusions
1 istory; Suppression Techniques
2 lassification of Instabilities
3.1 Instantaneous Pressure Measurements
3.3 ethods of Data Analysis
3.4 ptical Methods for Combustion Instability Analysis
3 . 5 Other Techniques
4 xperimental Studies of Linear Acoustic nstabilities Fundamental View
point)
4.1
dasis for Fundamental Laboratory Studies
4 . 2
Qualitative Experiments Employing Acoustic Generators
4.3 hock Tube Techniques
4.6.1 efinition
4 . 6.3 Specific Configurations
4 . 8.4 Experimental Strategies
4.6. 5 Experimental Results
6 xperimental Ctudies of Nonacoustic and Nonlinear Combustion Instabilities
6.1 Introduction
6 . 4.1 Introduction
6 . 4.3
Explanations in Terms of Intrinsic Combustion Characteristics;
Strand Observations
6 . 4 . 4 Explanations in Terms of
Combustion Response Residence Time
Interaction; * Instability
Chapter 10
1 Introduction
2. 3
Acoustic Energy in a Sound Field
3 coustic Amplification
3 . 2
Relationship Between Admittance and Energy Growth Rate for
Monochromatic Waves in Cavities
3.3 Alternatives and Generalizations
4 coustic Damping Mechanisms
Discussion of Results
Non-One-Dimensfo^al Oscillations
Sonic Nozzles
4.3.1 Wall Friction
4.3. 2 Wall Heat Transfer
4 . 3 . 3
Complex Wall Loss Phenomena
. 4 Homogeneous Damping
4 . 4 . 2
Chemical and Molecular Relaxation Losses
4.4.3
Other Homogeneous Damping Processes
4 . 5 Particle Damping
4 . 5
. 2 Analysis for Very Small Particles
4 . 5.3 Interpretation of Results
4 . 5 . 4 Accurate Formulas
4 . 5 . 5 Experimental Verification
4.6 iscoelastic Damping in the Solid
4 . 6.1
General Aspects of Solid-Phase Losses
4 . 6 . 2
Vibrations of Gas-Solid Systems
4.6.3
Implications Concerning Attenuation
4.7 ummary
5.1 Introduction
5 . 2.1
Relationship between Admittance and Perturbation of
Mass
Flow Rate
5 . 3 Time Lag Theories
5
. 3.1 implified Time I ag Concept
5 3. 2 mproved Time Lag Theories
5 . 5.1 Low-Frequency Response
5 .7 mall-Amplitude Erosive Effects
5.7.1 Introduction
5.7.3 coustic Erosion without Steady-State Erosion
5.7.4 ombined Steady-State and coustic Erosion
6 heories of Nonlinear and Nonacoustic Instabilities
6.1 Introduction
6.3 nherent Instability
7 omparison of Theory with Experiment
Chapter 11
1 ntroduction
1.1 verview
1.2 iscoelasticUy
2.1
Differential Operator Representation
2.3 omplex Modulus and Complex Compliance Representations
2. 4 Temperature Effects
2.6 Nonlinear Stress-Strain Relations
3 . 2
Stresses in an Encased Viscoelastic Cylinder
3.3 tresses in an Encased Viscoelastic Spinning Cylinder with Ablating
Cavity Surface
3 . 5 Grain Slump du
e to Axial Acceleration
3.6
Viscoelastic Cylinder Reinforced by Elastic Wires
4 ailure and Failure Criteria for Solid Propeliant Rockets
4.1 Preliminary Remarks
4.3 efinition of Failure Criteria
4.4 onclusion
Chapter 12
1 ntroduction
XX
5 echnological Development - mprovement of Flexibility and Thrust Control
Capabilities
Addendum
xxi
Preface
For ten years there has been an acute
a in English
on olid propellant ockets.
Books have been published on propulsion in general
and on rocket propulsion specifically
but these sometimes tend to de-emphasize
propulsion by means of solid propellants, n order to devote more space to topics
such as liquid-propellant or nuclear propulsion, which have occasionally been
deemed more exciting or more exotic in some respects.
The most recent book
in English which is devoted exclusively to solid propellant ockets is the hort
monograph by Wimpress, ntitled Internal Ballistics of Solid Propellant Rockets'
and published in 9 5 0 .
There are tw
o recently published books in Russian on
solid propellant rockets, ne by Kurov and Doijanski (1961) and on
e by Zeldovich
and Rivin 1963), ut these have not been translated into English yet.
The need
for an English-languaf <
; text on solid propellant rockets provided the underlying
motivation for writing the present volume.
The objective of this book is twofold, irst to present basic material on solid pro-
pellant rockets which can be used for classroom instruction and second to carry
the reader to the frontiers of research in a number oi specific areas of solid pro-
pellant rocketry.
Although there is some material (e.g., Chapters , and 12)
which might appropriately be used in undergraduate courses, he instructional
value of the book lies primarily at the graduate level.
An attempt has been made
to enhance the educational utility of the monograph by presenting the more ele-
mentary aspects of the ubject first Chapters
to ), efore proceeding to de-
tailed an
d more advanced treatment f specific areas of research Chapters
to
11).
An attempt has also been made to present the esearch topics in a peda-
gogic manner, o aid the graduate tudent or the practising engineer w
ho is not
familiar with the ubject material.
Research workers in the field of solid propellant rocketry should find this present-
ation useful, oth as a reference to previous research endeavors and as a guide to
desirable avenues for future research. he book delves more deeply into a number
of areas of research than any previous volumes on the ubject have done. ndeed,
progress in the field has been continuing so rapidly that it has not been possible
for earlier books to attain the depth of the
monograph in the pecific areas
of esearch chosen for emphasis herein.
It seems appropriate to record here the parts of the book for which each author
assumed primary esponsibility.
Huang wrote all of Chapter 1, xcept for Sec-
tion 4 n Failure Analysis which was prepared by Barrere and Section 1.1 which
was prepared by Willisms.
The rest of the book was written ointly by Barrere
and Williams, with Barrere preparing the first drafts for Chapters 3,4, 5
, and
for most of Chapter 8
, nd with Williams preparing the first drafts for Chapters
2,6. and 0. Chapters
nd 2 were written jointly by Barrere and Williams,
who als», ointly evised successive versions of the entire monograph in an effort
to make it into a coherent work.
Williams is responsible for the final English-
language editing f the manuscript.
W
e wish to thank many of our colleagues, specially those at the University of
California, an Diego, t the Office National d'Etudes et de Recherches Aero-
spatiales and at the Direction des Poudres for numerous stimulating discussions
relating to many aspects of this work,
The list of names of those to whom w
e are
indebted is to
o long to be presented here.
However, e must explicitly thank
Professor S. S. enner for his nterest an
d for his aid in initiating this work,
Dr. W.R. Maxwell
for his constructive eview of .he monograph and Mr. . W
.
Price for his welcomed review of Chapter .
The many hours spent by Simone
xxii
versions of the manuscript were essential to the successful completion of the book.
One of us, . A. Williams, wishes to thank the Propulsion Division of the Air Force
Office of Scientific Research for continued support of research Grant N
o. AF-
AFOSR-927-67 and also Project THEMIS) on related subjects during the writing of
this material.
Another author, M. Barrere hanks the Direction des Poudres for
granting permission to publish some results which were obtained under contract.
W
t lso extend our thanks o AGARD for supporting this writing under a contract
supervised helpfully by Colonel C
h. Lupold.
M. Barrere Paris, rance
Motors- Generalities
o
F hrust
g * erage acceleration of gravity experienced by vehicle over its flight
path
g
0
K
f onstant of proportionality, ependent on technology
K ^
, onstant f proportionality, ependent n technology
^ ayload weight
M
bV
M L aunch weight
t
b
v PML Mp, atio of payload mass to propellant volume
o t onstant see Eq. 1-5)
A
v elocity change imparted to a given payload
Av
D
ß ropellant density
Motors- Generalities
1 . Introduction
In general a vehicle is propelled by forces, ermed thrusts, which provide a desired
component of acceleration.
These forces can be produced in a variety of ways.
Solid propellant rockets are examples of a pure eaction system in
which the pro-
pulsive forces are produced by the ejection of mass propel«ants) initially contained
in the system.
Self-contained systems f this type are called rocket motors and
can operate n space as well as n atmospheres, ince
iey do not require an ex-
ternal propulsive fluid.
Four categories of rocket motors may bo defined, ccording to the physical state of
the propellant materials carried within the ocket.
These are solid propellant
motors, iquid propellant motors, aseous propellant motors and hybrid motors
(which contain propellants stored in at east tw
o of the three physical states of
matter).
The solid and liquid propellant motors and the hybrid motors employing
solid-liquid combinations are of greatest practical interest because of the heavy
tanks needed to store arge masses of gas.
This book is concerned only with the
first of the four categories, olid propellant rocket motors.
Additionally, ttention
is estricted to chemical propulsion, or w
n
j
ch energy that is necessary for pro-
ducing large hrusts s stored in th
e form of chemical energy of the propeHants.
A solid propellant rocket is he simplest form of chemical propulsion.
The fuel
and oxidizer are both ncorporated in a single solid, alled the propellant grain,
located inside a container called 'he combustion chamber.
This chamber is arge
in comparison with th
e combustion chamber of a iquid propellant rocket motor.
A schematic illustration of this type of motor s hown in Fig. -1 (a).
The grain
shown here s tubular with a star-shaped cross ection.
A device called an igniter,
which is designed to initiate the burning, s placed inside the central cavity of the
combustion chamber.
After gnition ie hot gases, which are produced when the
solid burns, lo
w through the central cavity and are accelerated to a high velocity
by means of a nozzle.
It will be seen in Chapter 2 that the esulting ejection of
gases at high velocity greatly enhances he production of a propulsive thrust on the
motor.
During combustion, ases evolve from the solid propellant grain only at its urface.
Thus, he surface of the solid regresses normal to tself during burning, l the
"linear
regression rate" of the propellant grain.
The combustion gr~es come
into contact with the outer shell or case f the chamber only at the end of a firing.
The purpose of the case s to contain the propellant and to withstand the high cham-
ber pressures that are produced during combustion.
Figure 1-1 (a) emphasises he
simplicity that arises from storing the propellant inside the chamber.
It s apparent from Fig. -1
(a) that the principal parts of a solid propellant rocket
an d -l(c).
In a typical iquid propellant motor, uel and oxidizer are stored in
separate anks.
They are conveyed to the combustion chamber by means of a feed
system and are injected into the chamber an
d partially mixed by means of an injec-
tion system.
The principal pa;.ts of the liquid propellant motor are thus he ttnks,
the leed an
d control) system, he combustion chamber and the nozzle.
The hybrid
moioi llustrated in Fig. ~l(c) contains a solid fuel and a liquid oxidizer.
Its
main components are an oxidizer tank, teed an
d control) system, n njection
system, combustion chamber, solid fuel grain an
d a nozzle.
It will be seen in Chapters
and 3 that the propulsion systems llustrated in Fig. -1
rely on exothermic chemical reactions for their effectiveness.
There are numerous
applications such as space propulsion) fcr which these chemical rockets are cur-
rently the most important propulsive work horses".
The role and current standing
of ol?d propellant ockets among these work horses will be discussed on page 9
after w
e have described solid propellant rockets n somewhat greater detail page 4
an
d classified them according to their use page
).
A brief history of solid pro-
pellant rocketry is given on page 4
and an outline of the development of the rest of
this AGARDograph s presented on page 1
Additional material relevant to solid
propellant rockets may ? found
in the bibliography listed as Ref. -16.
2 Description and Brief History
2.1.
Description ot Current Solid Propellant Rockets
W e shall no
w discuss the solid propellant rocket system more deeply in order o
accustom he reader o some specific terms used in solid rocket technology.
The
system s composed of four basic parts:
- he propellant grain
- he case
- he nozzle
- he gniter
The propellant generally consists of an oxidizer and a fuel, nd the most common
type today consists of crystalline ammonium perchlorate dispersed in a plastic
fuel binder.
The performance of this propellant s increased by the addition of
finely ground light metals, uch as luminum.
The basic propellant defined here
is an exampb f a composite propellant, nd the addition of metals is eferred to
as netalization.
The mean size of the ammonium perchlorate crystals is f the
order f 5 0 M
, nd that of aluminum powder is typically
to 30M. The utilisation of
a plastic binder enables one to construct arge-size grains with good
mechanical
properties. These grains may oe
fabrk ited outside the casing or poured into it.
In addition to composite propellants, ouble-base solid propellants are also used;
these were the principal solid propellants
twenty ago They consist f a
mixture of nitrocellulose and nitroglycerin, ach o
i" which possesses both fuel and
oxidizer characteristics.
The periormance of this aouble-foase propellant s
lower than that of the composite propellants just described.
The grain geometry provides a basis for a prelim^ary classification of solid rocket
propulsion systems.
The grain configuration w i'
essentially depend upon the
mission, .e. po n the thrust an
d burning time, r more precisely po
n the thrust-
time history.
Once the grain is ignited, urning generally progresses until the
propellant is completely consumed.
The b ning surface geometry and its time
evolution will then impose the thrust-time history
The grain is called neutral f
the thrust emains constant during the entire burning history,.
This occurs when
the total burning surface area does not vary with time.
There also exist pro-
fressive grains for which the thrust increases with time
an 3 regressive grains
which Ihe thrust decreases with time,
internally burning cylinders or tub*s
FEED SYSTEM
INJECTION SYSTEM
(b ) chematic diagram f a liquid prop
ell aiit rocket motor
G AS GENERATOR ONTROL VALVE
current configurations are star-shaped cylindrical grains which provide a relatively
large propellant surface area for burnt gas emission while maintaining approxi-
mately neutral burning.
The motor burning time depends upon the propellant
thickness and generally increases with increasing rocket ize.
The tar-shaped cross-section is no
t the only grain geometry used;
more complex
cross-sectional shapes are also employed.
Complex geometries correspond to
special applications, .
g. a high gas flow rate may be obtained by increasing the
surface area of the propellant.
These different configurations will be studied in
detail in Chapter 4 .
However, e note here that neutral burning is also achievable
with coaxial cylinders, lthough this geometry poses problems associated with the
necessity of exposing grain supports to the ho
t gaseous combustion products. The
simplest grain configuration is a cylinder burning at its end (cigarette burning},
bu
t this configuration has a low ratio of burning area to nozzle throat area, nd is
applicable only for on
g burning times and low values of the tnrust per unit cross-
sectional area of the rocket.
Besides cylindrical grains, pherical grains are
useful for ome applications.
Grain neutrality is difficult to obtain in spherical
configurations, ut it can be produced by employing grains with tw
o compositions
having different burning rates.
Grain configuration controls the thrust-time pro-
gram of solid rockets and has to be adapted to each mission.
The geometry of the motor case is related to that of the grain.
Case resign also
depends on the application. Tw
o principal types of casing materials are currently
employed,
metalic materials and glassy materials.
Engineering problems arise
in connection with the selection of case materials, ase manufacturing proces*
5
« » «
an
d the tradeoff between conflicting requirements of light weight an
d reliability. Three
important elements must be considered in motor case design:-
a) he mechanical load during motor operLäon the case is subjected to high
pressure, everal tens of atmospheres):
b) he thermal load some parts of
the case which are in contact with burnt
gases must be thermally protected);
c) he auxiliary means of thrust-vector control devices for controlling the
direction of the force vector acting on the rocket are structurally
supported by the case).
Many metalic materials, uch as steel, re useful for motor cases because they
tend to have a high modulus of elasticity and a high yield strength, nd also
because the associated dynamic problems vibration of the structure) are less
difficult to solve.
Expansion of the propellant grain is a less severe problem
with metalic cases than with glassy cases.
Of the many glassy composite mater-
ials that have recently been considered for motor cases, he one that currently is
adopted most often consists of wrapped fiberglass filaments impregnated with geli-
fied epoxy resin.
Many machines have been developed for fiberglass case
winding.
A machine designed by Aerojet-General Corporation for manufacturing
F aris rocket motors is hown in Plate 2;
more than one and a half million miles
of glass filament o into <sich chamber measuring more than 4 feet on
g and 4-
feet in diameter.
The advantage of glassy cases is that they often provide
lighter-weight tructures than metalic cases.
The propellant grain-case bonding must be accomplished with care.
A plastic
liner is generally inserted between grain and case.
The liner has a manifold
purpose:
it acts as a combustion inhibitor, t prevents the burnt gases from
coming into contact with the case, t protects the wall thermally when the burning
surface reaches it
anci it acts as a mechanical bonding between grain and casing.
negligible compared with that of
the empty rocket.
Thermal protection of the for-
ward and rearward portions of the case that re exposed to hot gases during the en-
re firing is achieved by using high-temperature, ften silica-reinforced plandcs.
The nozzle attached to the downstream nd of the rocket motor consists of a
con-
vergent section, narrow-diameter throat
and a divergent ection.
It wiil be
seen in Chapter
that these three elements are needed in order to accelerate the
ho
t combustion gases to the high velocities required for efficient production of
thrust.
Since solid propellant rocket nozzles are generally uncooled, t is neces-
sary to use nozzle materials capable of withstanding a high thermal load.
The
manner of thermal protection of the convergent section will depend upon the aft-end
geometry single or multiple nozzles).
Thermal protection of
the throat is the
greatest nozzle problem because the maximum heat transfer occurs there. he throat
must generally be constructed from layers of different materials.
The layer that
sees the hot gas is composed of a high-temperature, efractory or metal material
that exhibits a good resistance to erosion e.g. ungsten, r graphite covered with
tungsten);
for arge rockets it is sometimes possiMe to permit some erosion of
the throat, o that uncoated graphite throat inserts can be employed.
A material
with a higher specific heat and a lower thermal conductivity is usually placed
underneath the throat insert in order to absorb the heat load and to prevent it from
being transmitted to the rest of the structure.
Finally, material having good
mechanical properties is placed outside in order to resist the transmitted
pressures.
The divergent ection is often composed of an ablative material such
as einforced plastic).
Rapid thrust termination is required in certain applications, .g. at staging times
for satellite launchers. This is often achieved by means of openings located at the
head end of the motor th
e opposite end from the nozzle) and initially blocked by
diaphragms.
These diaphragms burst on command, hereby exposing the chamber
to the ambient atmosphere.
Most of the chamber gases then begin to exhaust
through the head end, ausing first thrust direction reversal and then extinction due
to the rapid gas expansion.
The presence of such openings in the case raises
some structural problems, specially for fiberglass cases;
it would be necessary
either to of
the glass filaments or wind the
fibers the
during fabrication.
Solid propellant rocket motor igniters often consist of electrically initiated, on-
ventional pyrotechniccompositions. The combustion of an auxiliary propellant con-
tained in the igniter generates ho
t gases which come into contact with the grain
surface and induce ignition of the grain.
These pyrotechnic compositions generally
consist of oxidizer-metal mixtures e.g. otassium perchlorate-aluminum).
The
igniter combustion often produces ho
t condensed particles e.g. olid or iquid
alumina) which impinge on the grain surface, ausing high local heat transport and
local ignition of part of the surface area, ollowed by propagation of the flame to
the rest of the grain.
Mechanical properties of pyrotechnic compositions
are
usually such thai for safety they must be shaped into pellets, which are placed in-
side a perforated enclosure, ositioned so that the burnt gases will come into con-
tact with the grain surface.
For long rocket motors, ometimes small auxiliary
rocket engines are used as igniters n order to provide more nearly simultaneous
ignition of
the entire grain surface than could be achieved with conventional tech-
niques that rely on flame spread.
2. 2 History
Although the origin of rocketry is obscure, ndoubtedly the first ockets used solid
propellants.
Rockets are believed to have originated either in China or in Greece.
I rora this date onward, many solid propellant ocket weapons have been
k i > . - v l \ad vised in battles.
Around 800
, ockets varying in weight from 8 to
" f t «-tre
onstructed in England William Congreve)
Fig. 1-2) and the
i asalio.i ;.
projectile was improved by William Hale.
That uch weapons
/ere put o use S demonstrated, or example, y the fact that the British attacked
Copenhagen n 'S O
7 väth some 30,000 rockets.
From the military point of view,
rocket weapons have continually progressed to their current state of sophistication.
It is curious o note hat astronautics pioneers ike Ziolkowsky in Russia. Oberth n
Germany, Goddard ir he United States and Esnault
Pelterie n France, whose
studies were carried out between 1900 and 1
9 3 0
, onsidered only liquid propellants
for «pace missions because they believed that iquids were essential for providing
sufficient energy for such operations.
This opinion rested on the fact that until
1900 black powder, onsisting of charcoal, ulphur and saltpeter, was the material
used as a solid propellant.
Even in 9 3 2
, fter changes n the propellant com-
position that resulted in
smokeless powder double-base composition) no grounds
existed for forecasting performance mprovements that could lead to the possibility
of using solid propellants for space flights.
Later, ouble-base solid propellants
were considered for uch
applications because of their improved mechanical prop-
erties and higher performance.
The mechanical properties of double-base pro-
pellants enabled one c envisage grain geometries with arge combustion areas an
d
consequently high hrusts.
B
ut the most important step in the progress of solid propellant rocketry was taken
in 94
4 by ihe Jet Propulsion Laboratory research workers, w
ho developed the
GALCIT propellant, onsisting of approximately 5
% potassium perchlorate and 5 %
asphalt-oil mixture.
The development of this composite solid propellant was to
open a vast field of research on high-energy solid propellants.
This GALCIT pro-
pellant ha
d some defects uch as poor temperature sensitivity which ed to bad be-
havior at ow temperatures.
However, undreds f housands of JATO-type
rocket engines for assisting aircraft take-offs wen onstructed using this pro-
pellant an
d the working safety of solid propellant ockets was thus demonstrated»
Solid propellant rocket techniques have been widely improved since 9
5 5 , rinci-
pally in tw o directions:
a)
development of propellants with higher performance specific impulse
and volumetric pecific impulse
ee page 9 for preliminary definitions of these
tw
o terms) and with better mechanical and combustion properties;
b)
development of lieht-weight structures for cases and for other motor
components.
These improvements have led to the construction of high performance engines that
can compete well with liquid or hybrid systems.
O
ne end result is the use of
large solid-propellant motors as powerful boosters, uch as in the zero stage"
of the TITAN HI C
, which achieves a total thrust of 2.4 million pounds from tw
o
large solid-propellant, egmented motors trapped on either ide of a first-stage,
liquid-propellant, ore motor parallel taging).
Another esult is the completely
successful static test firing of a 26
0 in. iameter olid-propellant rocket motor
containing , 6 7 3 , 00
0 b. f propellant and producing more than
million b. hrust.
A third consequence is the use of very light olid-propellant engines for upper
stages of satellite aunchers, aving a ratio of propellant weight to total weight
exceeding 90ftb.
3 . Classification
There are several possible bases for categorizing current types of rocket propul-
motor.
A crane lowers the 0 ton segrm
. Photo, ourtesy
of United Technology Center
L
I
ii
classification scheme based on
the field of application.
The principal applic ions
of solid propellant ocket ystems can be listed
under the following headings:
Rocket motors for ower stages boosters)
Rocket motors for upper tages of ballistic missiles and for space vehicles
Satellite aunchers complete vehicles)
guidance and control of vehicles
Assisted take-off rocket motors for aircraft
Military applications, .g. uided weapons, nguided bombardment and
air-to-air ockets
Diverse applications
3.1
Rocket Motors for Lower Stages Boosters) - Solid propellant systems are of
interest for u s «
» a3 boosters because their high propellant density leads o improved
overall performance which in some instances exceeds that of competitive liquid
propellant systems see page 9 .
After the February 3 96
6 successful test of
a 26
0 in. diameter rocket motor, t is no
w possible to consider very powerful
first tages using solid propellants.
The era of solid propellant boosters is in its
infancy;
in some such applications they are already beginning to replace liquid
propellants.
The 6 0 in. motor provides on
e example of a large solid propellant booster ocket.
It employs a composite propellant consisting of polybutadiene, mmonium perchlo-
rate and aluminum.
Some data that we have not already mentioned is that the
pressure measured
in the chamber during the full-scale test was very close to the
calculated one;
it reached its marimura value of 01
psia after 0 . 5 seconds. The
effective burning time 14 seconds
the peak
was . 51 million
The ignition period lasted 3 3
5 milliseconds.
The nozzle throat ablated at a rate of
4.8 0"
in. er second.
Another example of a large solid-propellant booster ocket is the strap-on motor
developed by United Technology Corporation for the TITAN II
I C , which is an im-
proved version of the TITAN II
missile.
This application illustrates the "zero-
stage" concept for improving capabilities of existing boosters;
viz., n operational
launch vehicle can effectively be launched from a high altitude by providing the
entire vehicle with a zero-stage) booster.
Often this approach affords a less
expensive solution to the problem of achieving higher weights n orbit than is
afforded by the development f a new launch vehicle.
The TITAN IIIC strap on
zero-stage motors are each 20 in. n diameter and develop more than one million
nounds thrust each.
The complete vehicle was successfully fired on December 1,
.965.
Each Strap-on motor for the TITAN II
I C consists f 5
segments, weighs about 25 0
tons
and burns for about 20 seconds.
The direction of the thrust vector is con-
trolled by injecting nitrogen tetroxide through the wall of the divergent ection of
the nozzle. In so larg«.
rocket motor, he thicknesses of the walls and of the
liner are mall in compar'son with the grain thickness, nd the question of wall
thermal protection can be solved relatively easily by using a protective coating
about inch thick.
Three mm-thick layers of liner are attached to the case by
centrifugal coating, he last layer bemg put in just before loading. The case is made of
LADISH D6AC teel. The
propellant grain is attached
segment. The cen-
tral core through which burnt gases flow is a circular cylinder, he ateral faces of
which are no
t inhibited from burning.
Adjacent egments are held together by
Clevis pin joints.
Tbc egment-joining technique is shown in Fig. 1-3, nd Fig. 1-4
12
Another example of the use of solid propeliant boosters is to accelerate a amjet to
its operating speed.
The experimental missile 3TATALTEX, ntended for studying
high-altitude, igh-speed ramjet propulsion flight Mach number greater than ) is
shown on Plate L
This ramjet is launched by a solid propeliant booster.
There are, f course, variety of other large ^cale and small-scale booster appli-
cation*» of solid propel
Umt rockets.
3.2. Rocket Motors for Upper Stages and for Space Vehicles - Solid propeliant rockets
are also used injection
satellited and as slow-down stages to
atmospheric e-entry- The change in vehicle speed (Av) required in these applica-
tions is obtained most easily by usinf; very ügiit ocket motors rather than pro-
pellants with high pecific impulse.
Reinforced piastic cases can leac o propeli-
ant mass fractions ratios of propeliant mass to total vehicle mass) higher than
0 . 9.
An example of a solid propeliant rocket used in such space applications is
shown in Fig. 1-5;
this motor has been developed by Hercules Powder Company for
orbiting a spacecraft whose purpose is tc detect high-altitude nuclear explosions.
The third stage of SEREB Diamant launcher is shown on Fig. 1-6;
tics are as follows
4 5 seconds
varying between 2 , 7 00 and 5 , 3
00 kg
9.6 cm
orthostrasyl ablative plastic)
Solid propeliant ystems are ill-adapted to the propulsion of space vehicles that e-
quire rocket motors with long burning times and low thrusts.
However, ertain
space missions are well-suited to solid rocket propulsion, articularly when the
trajectory or orbit of the vehicle is to be modified impulsively (large thrust for
short duration).
propeliant motors are
used as the primary means
transmitting a specified Av to a space vehicle.
Guidance gas rockets are also
needed in such systems to provide fine orbit adjustments after the solid rocket fir-
ings.
cations satellites) use solid propeliant apogee rockets that are usually spherical in
shape and <hat place the satellite in a desired position with respect to the earth, n
order for the atellite to be able to receive and emit signals in correct directions.
The mission or these apogee injection motors is to increase
the payload speed.
lft. n diameter, sed to nject satellite
from orbiting vehicle nto higher-altitude
orbit, courtesy of Hercules Powder Com-
pany.
mant auncher, ourtesy f ONERA
Fig. -7
Nike-Cajun ocket, ourtesy of NASA
3.3
Satellite Launchers -Since olid propellant systems are attractive as boosters
because of heir high propellant density, and ince hey have become competitive as
injection motors because of he achievement of ight structures, hey
possess moüt
of he qualities hat are necebsary or a complete atellite auncher.
Advantages
of uch yF*°mbstem rom their ow manufacturing costs and from heir eliability.
One of the earliest examples of a vehicle propelled olely by olid propellants and
capablo of orbiting a satellite s he SCOUT, which contains our olid propellant
stages.
characteristics of he ystem are the
Total ength
Initial weight
Length
Diameter
Length
Diameter
3 8 econds
With his vehicle t s possible o place 300 b
in a low) 300 n.mi. rbit.
The
principal purpose f he SCOUT s o orbit cientific ayloads for nternational
uses.
3.4
Sounding Rockets - Sounding ockets are used for the exploration of the upper at
mosphere and space n order o gather cientific data.
There exist many ypes of
such vehicles, ot ess han 0 n he United States alone.
They are characterized
by heir pay load and by he maximum altitude that they can each.
One of the ear-
liest meteorological ockets s the one-stage ARCAS ocket which weighs 5 b and
lifts a 2-lb payload to an altitude of 0 mi.
The two-stage ARCAS ocket weighs
10 0 b and
an ift 2 b o an altitude of 0 mi.
As an example of a arger ounding
rocket, we describe he NIKE-CAJUN used by NASA.
The characteristics of this
rocket, which s hown n Fig. 1-7, are he ollowing
Length
Diameter
00 mi
4 200 mi/h
The vehicle consists of wo olid propellant tages, he first HERCULES) having a
thrust of 8 0
0 b and the econd having a thrust of
620 b The econd tage
weighs 257 lb.
5
3.5. Auxiliary Rocket Motors o perform Guidance nd Control Functions - Fe
w solid
propellant rocket motors are specifically designed for he purpose of controlling
th
e hrust vector of a main engine or of otherwise providing th
e attitude control
forces led for steering a vehicle under
hrust.
In ne exception, orces or
th
e control of attitude nd hence the direction of main engine hrust are obtained
by means of four small solid rocket motors ocated at the aft end of the vehicle.
Their nozzles are s *t at an adjustable angle with espect o the axis of he main
rocket motor.
Torques o change attitude, nd hence hrust direction modifications,
are achieved by varyng the orientations of the nozzles of these ocket motors.
This in
d of vehicle atitude control s employed in the four-stage olid propellant
rocket BERENICE, hich is designed for studying in flight ihfi kinetic heating of
hypersonic vehicles during atmospheric re-entry.
In the photograph of the
BERENICE that is shown n Fig. 1-8, ne can see he foar control ocket motors
at th e aft en
d of the vehicle. T
he burning time of these control motors s f the
same order of magnitude as hat of the irst stage.
The propellant grain geometry
of each s cylindrical with cigarette burning.
The ourning rate is ncreased by
placing silver wires n the propellant perpendicular o the burning surface„
Solid propellant rocket motors are
irely used o supp'y the attitude control forces
for satellite guidance because of their short burning nines. ompressed gases or
subliming solid substances can perform this unction better.
Solid propellant mpulse ockets have been constructed for pinning up certain miss-
ile tages PET rocket) an
d also for accomplishing stage separation.
3.6. Assisted Take-off RocketMotors or
Aircraft- Few rockets are ow usedtor n-
creasing th
e ake-off power of aircraft because of improvements n the ea
k power
output of turbojets.
The most famous ake-off rocket is he Aerojet JA
TO, he
propellant of which s an extruded composite.
The ATO motor has he following
characteristics:
Thrust
2
Rocket motor weight 2.5 kg
3.7. Military Applications - The most mportant weapons composing the arsenals f
armies nowadays use solid propellant systems as a propulsion mode.
Such ystems
have indisputable superior.'ies: per: lanent readiness f th
e engine, oo d reliability,
low maintainance, asy preparation of a firing.
Military ockets are distinguished
according to heir mission, .e. urface o-surface, urface-to-air, ir-to-surface,
air-to-air an d anti-submarine.
For surface-to-surface missions, hree ypes f missiles are distinguished.
a)
There are ntercontinental ba istic missiles ICBM), uch as he MIN-
UTEivIAN which consists f three olid propellant stages.
The missile ange
is about 6000 to 7 000n.mi.
TheMINUTEMAN an e prepared for iring n
about 3
2 seconds nd equires about 0 seconds or he selection of an objec-
tive.
The overall ength of the hree stages or different missiles n his
series s A-53.17ft, B- 5 9ft,
The missile s about 65000
A MINUTEMAN is hown n Iight n Fig. -9.
The principal contractors or
MINUTEMAN propulsion are:
THIOKOL for he irst stage: AEROJET for the
second an d HERCULES for the hird. T
he irst stage has iameter of
1.65m nd a propellant weight of about 0 tons.
The case s made of LADISH
D6AC steel.
Four nozzles with axes aligned with he vehicle are used for he
first tage;
their divergent sections employ phenolic efrasil or heat protec-
tion.
It has been difficult to solve problems associated with he hermal pro-
tectionof the interior chamber walls.
The single-composition, olid propellant
testing re-entry vehicles at Mach 2, ourtesy f ONERA.
IS
grain ha.3 a star-shaped configuration 6 arms).
The propellant consists of
PBAA polybutadiene acry^c acid) and ammonium perchlorate.
A press-
ure peak observed at the beginning of a firing has been decreased by a part-
ial inhibition of the grain.
b) here are surface o-surface missiles of intermediate range IRBM) such
as the POLARIS series of strategic missiles which are fired from submerged
submarines. The three variations
propellant stages.
These engines have the following characteristics:
A i ength
t
ype is about 1 00 n.
mi., hat of the A
2
3
is 5 00 n.mi.
c) here are short ange missiles, uch as the ERSHING rocket range
from 00 to 4 00 n. mi.) with tw
o solid propellant stages, r the SERGENT
(range 25 to 7
5 n. mi.) consisting of only on
e solid propellant stage. The
smaller HONEST J O H
N range 2 n. mi.), onsisting of one solid propellant
stage, lso falls into this category;
it has the following characteristics:
28 ft
4.5 ft
: 4 00 lb
O ne t the small