JPRS ID: 9351 TRANSLATION LANDING OF SPACE VEHICLES ON PLANETS BY V.I. BAZHENOV AND M.I. OSIN

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JPRS L/9351
17 October 1980
- Translation ~
LANDI~lG O~F SPACE VEHICLES ON PLANETS
- By
V.i. Bazhenov and M.I. Osin
~
FBI$ FOREIGN BROADCAST INF~JRMATION SIER~'ICE
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JP RS L/9 351
17 O~tober 19 80 -
LANDING OF SPACE VEHICLES ON PLAiVETS
- Moscow POSADKA KOSMICHESKIKH APPARATOV NA PLANETY (Landing of
Space Vehicles on Planets) in Russian 1978 signed to press
16 May 78 pp 1-160
_ [Book by Vladimir Ivano~rich Bazhenov and Mikhail Ivanovich
Osin, Mashinostroyeniye, 2,500 copies, 160 pages, UDC 629.875.001]
~ CONTENTS
Annotation 1
Preface 1
PART 1. Landing of Space Vehicles on Planets With an Atmosphere...... 2
Chapter 1. Characteristics of Descent to Planets With an Atmosphere 3
Chapter 2. Problems Characteristics for the Return to Earth From
Near-Earth and Interplanetary Space Flights 28
Chapter 3. Choice of Design-Ballistic Parameters for Landing
Vehicles of Martian Planetary Systems 49
Chapter 4. Analysis of Tra~ectory Characteristics for Controlled
Entry and Descent Into the Atmosphere of Venus 64
Chapter 5. Determining Design Parameters for Descent Vehicles
to Sound the Atmosphere of Jupiter 71
PART 2. Methods for Studying Soft Landing of Space Vehicles on
Planets Lacking Atmospher.:s 81
Chapter 6. Theoretical Analysi~ of Sof t-Lrj.nding Dynamics
and oF Space Landing Vehicles 82
Chapter 7. Theoretical Foundations for Phyaical Modeling of
Soft Landing 105
' a - [I - USSR - A FOUO]
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Chapter 8. Experimental Studies of Soft Landir.gs Using
Dynamically Similar Models of Space Landing Vehicles 129
Chapter 9. Peculiarities of Terrestrial Experimental
Development of Soft Landing With the Aid of Full-Scale
Mockups of Space Landing Vehicles 152
Chapter 10. Several Results of Theoretical and Experimental
Studies of Soft Landing 164
Bibliography 156
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ANNOTATION
[Text] This boak explores problems encountered by developers of spa;,e
vehicles for descent and landing in studying vehicle structure and condi-
tions of motion. Methods for the parametric calculation of different
types of descent vehicles are presented, the problem of searchingfor
optimal parameter values is formulated, and methods of physical modeling
are described.
Analytical methods of studying the dynamics of soft Ianding are presented.
The results of applying these methods are fully accep table for practical
. engineering computations in terms of accuracy and scope. The book can be
~ useful to designers of space vehicles, to scientists, and to those engaged
in studying problems of descent and landing, as well as for graduate and
undergraduate students of institutions in the appropriate specialties.
PREFACE
The book is devoted to problems of space vehicles des cending to and landing
on planets. Success of space engineering in this dir ection is obvious;
man-made space vehicles are landing on the Moon, Venusfl and Mars and return-
ing successfully to Earth from orbits and interplanetary tra,jectoriea.
Current publications at homP and abroad describe methods f_or calculating
tra~ectories, gas dynamics, and heat transfer during the motion of a apace
_ vehicle through the atmosphere. As a rule, these books contain techniques
and algorithms that require laboricus numerical computational procedurea to
determine even some of the parametera of such space vehicles. Unfortuf
nately, it is impossible to degcribe precisely all op erations in the process
of space-vehicle dQSign and it ie necessary in practice to search for
analogies and general guidelinea in the conatruction of new space vehicles.
Methods of approximate parametric deaign estimates bas ed on simplified com-
putational models and on the reaults of exPeriments using methods based on
- the theory of similarity and dimensional analysis mus t be used to determine
- configuration.of space vehicles working under new and unusual conditions.
Obviously, it is impossible to encompass Che diversity af probl~:ms connected
with landing on a planet and returning to Earth within the limitations of a
single book. Even a brief enumeration of these problems with a description
and formulation of the basic tasks to be carried out durinq the design and
ground-based development of various landing vehicles would require a multi-
volume monograph. The authors of this book have therefore limited them-
selves to considering only the basic questions connected with the design of
future descent and landing vehicles and with the illumination of problems
that have to be solved in the near future.
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Problems of descent into the atmospheres of various planets are treated in
Part 1, which also presents the results of calculations of design and
research tasks connected with determining the configuration and motions of
descent vehicles. In particular, for descent to Earth, the con3itions of
motion based on load-factor criteria for piloted space vehicles are substan-
tiated for entr~ into the Earth's atmosphere at hyperbolic velocity, and an
; approximate mathematical model is proposed for calculating a controlled
descent tra,jectory. By using parametric r.omputations, optimum loading on
lifting surfaces suitable for a descent to Mars is determined; for Venus,
opcimal descent trajectories are determined based on the conditions at tran-
sition to the suosequent descent into denser layers of the atmosphere; and
modern concepts nave been formulated for design phase calculations of the
characteristics of a planetary sounding vehicle for a descent to Jupiter.
Prablems associated with space vehicles making contact with the surfaces of
planets and the related problems of landing-gear design can, to a certain
known ;iegree, be separated and considered independently of the problem of
- ruvtion in the atmosphere and during the touchdown braking phase. Part 2 of
the book is devoted to the set of pro'blems that arise during the development
of landing gears, including theoretical methods of calculdting the dynamics
of soft landing, methods for modeling and designing support kinematics,
designing shock absorbers, and improving the stability of landing systems,
and the experimental verification of theoretical results using full-scale
' mockups and dynamically similar models.
Part 2 of the book substantiatea the techniques for the physical modeling
of soft landir.g under Earth conditions. The derivation and analysis of con-
- ditions of similarity are presented, which are necessary for the simulation
of processes of lan3ing on the Moon and other planets. Such simulations are
r~quired for tne ground-based development of landing vehicles.
The book can be useful to specialists in designing and developing landing
systems for lunar, Venusian, Martian, and other descent vehicles, to tho9e
studying the specific probl~ms of the mechanics of landing and thermodynam-
ics of descent, and to students of appropriate specialties.
The authors are grateful to candidates of technical sciences A. F. Yevich
and R. P. Belonogov for useful counsel and the comments they made in review-
ing the manuscript, and to engineers A. I. Goncharov, A. I. Burtsev, Yu. V.
Zakharov, and V. F. Malykhin for help in preparing manuscript materials for
publication.
The authors wili gratefully accept all comments that readers might have
about this book.
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- PART ONE
LANDING OF SPACE VEHICLES ON PLANET~
WITH AN ATMOSPHERE
This portion of the book is devoted to ~roblems of space-vehicle descent
through tfie atmospheres of Earth and planets. The wide circle of special-
ists is familiar with books and texts that present methods for computing
and generalizing the problems of thermodynamics of descent (for example,
::~efs. 1, 20, 40, 43~. In these works, the emphasis is on obtaining approxi-
, mate relationships of a general character suitable for qualitative conclu-
j sions at the preliminary design stage of analysis. The methods and results
of analytical assessment~ presented in this literature are no doubt useful
to peraons beginning to study the problems of motion of traditional descent
vehicles in an atmospfiere.
However, space engineering continues to develop; new vehicles use modern
~echnology and employ many different specialists. These specialists need
more than ,just general information on balliatics and thermodynamics of
descent; they are interested in the professional aspects of designing new
flight vehicles. It is just for this reason that the present book attempts
to thruw light ~n the problems of designing future descent vehicles. The
characteristics of descent to each of the planets are considered witl-~y~ only
one of the two parts, with the authors intentionally limiting themaelves to
the exposition of general relationships and basica and to specific engineer-
ing aspects of descent under conditions characteristic ror each planet. The
goal was to capture the problem as a wh~le, to compare the various possible
engineering solutions for each type of descent vehicle, and to establish the
general outline inherent to the conditions of motion and shape of future
descent vehicles.
Part 1 of the book explores questions of selecting the descent mode and cal-
culating design characteristics of future descent vehicles. To start with,
the design features peculiar to descent vehicles and which differ from those
of other component systems of space rocket systems are pre~ented in general-
ized form. The typical design search problem that has to be solved in any
mechanized design at the top level of research and des~.gn selection is
described, that is, in an automated synthesis of the configuration and
structure, urlth consideration given to the conditions of motion in the
, atmosphere. Methods and results of analysis of the conditions of motion are
considered as they apply to the requirements of modeling future vehicles
with weight as the control parameter. Such vehicles, in the authors' opin-
ion, include rEUSable space rocket systems that return to Earth, as well a~
vehicles for landing on Mars or returning to Earth after an interplanetary
, expedition. The problem of deacent for heavy controlled automated planetary
systems to study Mars and Venus and questions of automated sounding of the
Jovian 3tmosphe�re are decidedly of interest.
A11 of these problems, to some detail, are sequentially illuminated in
Part 1 of the book,
Thus, we will start with the characteristica of descent vehicles that deter-
mine the spec:tfics of designing them, in contrast to the design of ottxer
space vehicles.
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CHAPTER 1
CHARACTERISTICS OF DESCENT TO PLANETS WITH AN ATMOSPHERE
Landing on planets with an atmosphere or returning to Earth from space is
one of the most complex and important stages of space flight. The task of
the descent vehicle is to land on the surface of a planet or to --eturn to
- Earth with some payload. In landing on a planet, the payload c~nsists of
the scientific research equipment; with returning to Earth, it can be the
fruits of a particular flight program. For manned space vehicles whose pay-
load is the crew and passengers either returning from space or about to land
on a planet, the pE.-io d of descent through the atmosphere represents one of
the most crii.ical stages of the expediti,on. This is because the descent
vehicle experiences the extreme effects of space flighL. The descent vehi-
cle, as well as orbiting vehicles and interplanetary and planetary space
_ vehicles, operate und er conditions of vacuum, weightlessness, solar radia-
tion, and meteoric eros ion. However, all theae effects, common to all apace
ot~ects, assume a secondary significance compared to the effects of the
atmosphere during desc ent. Approach to the planet and descent into its
atmosphere occur with a speed several times higher than the flight of even
the fastest aircraft, and deceleration from that speed is accomplished by
aerodynamic braking in the atmosphere. The inertial forces acting on both
structure and crew dur ing such braking exceed their weights; heat-shielding -
of the vehicle protects its airtight cabin from powerful heat flows that
emanate from the shock layer of the gas, heated at the forward part o: the
vehicle to several thousand degrees. Besidea static loaus, the vehic7e
_ experiences vibration and ahock loads. The process of braking in the atmos-
phere is, as a rule, short-l~.ved; however, large gradients in load factor, `
~ altitude, flight velocity, heat flow, and temperature in the various parts
of the structure appear during braking. The extreme loads thus are the
first difference between descent vehicles and o~her types of space vehicles.
The enumerated charact eristics of the proc~ess of deacent through an atmos-
phere define yet another difference between de~Gent vehicles and other com- ,
ponent systems of spac e vehicles. These charact;eristics make considerable
- demands on the reliabi lity of all systems of the descent vehicle. Thus,
while it is possible to abort a flight and return to Earth to save crew and -
~ pavload, or to evacuate a crew with the sid of a supply space ship, it is -
almast impossible to r escue the crew from the airtight cabin of a descent
vehicle in an emergency during the braking phase of flight in the atmos-
phere. Let us note that rescue from orbit or from a trajectory of in~ection
into orbit depends on using the very eame deacent vehicle, and the proba-
bility of failure of the descent and landing systems thus determines the
safety of flight. Thus, the reliability of the ~'escent and landing systems
has to be higher than the reliability of other subsystems and ships of the
, space rocket system.
The third diffe'rentiating characteristic of descent vehicles is determined
by the rigid weight limitations placed on the entire space rocket system.
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_ As a rule, descent vehicles operate in the conc~uding phase of space flight
and each excess kg of weight is equivalent to several kg of weight in the
propulsion systems of the mission system or tens of kg in the initial weight
of the booster. For this reason, problems of optimizing the descent and
landing system by using the weight criterion are important. Also, problems
of minimizing the descent and landing system weight are com~lex and inter-
esting and require a detailed account of the interplay between the condi- ~
tions of motfon and the configuration of the vehicles. These relarionships
determine the fourth diff erence of descent vehicles from orbital and inter-
planetary ships and stages, namely, the complex dependence of weight charac-
- teristics on thoae gaometric parameters that determine the form of the
descent vehicle. These parameters determine, on one hand, the distribution
of the heat flows and mass of the thermal protection system and, on the
other hand, the lift and braking characteristics of the vehicle, which char-
acteristics, in turn, determine the dur~.tioa, thermal resistance, and load-
factor regimes of the descent trajectory, and, in the final accounting,
det~rmine the total mass of structure, heat shield, and fuel of the propul-
si..on system for bra.king and control.
~nly an integrated approach to the problem of optimizing the geom etric, com- -
ponent, and weight-propulsion parameters of descent vehicles can guarantee
the achievement of reliable results in choosing optimal design solutions.
Unfortunately, as frequently happens, erroneous c~nclusions are drawn about
the advantages of one or another vehicle configuratioa during the detailed
- analysis of nonrelated problems in ballistics, control, or aerodynamics. If ~
the optimizing of descent-vehicle parameters is based on a weight criterion,
then the computational model has to provide procedures for describing the
cor.figurations, for estimating aerodynamic and centering characteristics,
for modeling of the conditions of motion and heat exchange, and for estimat-
ing the weights of the structure, heat shield, ana landing system, while
also considering vehicle loading, detail features, and the requirements
placed on the landing site. For each type of flight vehicle, the computa-
tional model is constructed differently, and design criteria and limitations
are accounted for differently. Examples are shown in subsequent chapters of
this book of how allowances are made for limitatio:~s placed on vehicles -
designed for different missions. It is possible to point to several general
_ principles for constructing computational models and to attempt to unify the
general statement of the problem of op~imizing parameters of different types
of descent vehicles. It is assumed that use is made af numerical search
methods during optimization and that the choice ot design parameters is made
using a system for the automated synthesis of descent-vehicle configurations.
The typical algorithm for design or verification calculations used in a sys-
tem of automated search and selection of design solutions for descent vehi-
cles as a whole must necessarily include the following specialized computa-
tional models:
1. es *^UC111~.e or group of modules for describing the vehicle configuration,
for mathematical modeling of contours, and for computation of dimensional,
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volumet.ric-centering, and layout characteristics. For simply-shaped vehicles
formed by co_nbining several geometric shape~, a special computational routine
has to be provided for the calculation of dimensions, areas, and volumes.
As a rule,. the group of modules describing the configuration is combined wit~r-
a special procedure for generating the vehicle shape by using grzphic infor-
mation display devices (graphic displays, plotters, ~r digital printing -
devices).
2. A group of modules for the calculation o� aerodynamic characteristics.
Different modules are, as a rule, created for different classes of shapes and _
different regimes of flow.
3. Mcd~~les and routines for the computation of the conditions of motion.
These routines can differ by the degree of detail in design e3timates, they _
can use different principles and procedures of calculation, and they can form
part of the modules for computing aerodynamic and weight-power characteris-
tics.
4. Modules for cumulative calculation of weight-power characteristics.
These modules process and su~narize information for ~udging the appearance of
the vehicle as a whole. Depending on the formulation of the design search
problem, it is possible to include, as part of these modules or separately,
procedures for strength calculations, thermodynamic calculations, computa-
tions of dynamic processes, estimates of cost, synthesis of design layout,
and so forth.
All parameters determining the appearance of a typical descant vehicle can be
di~~ided into several groups. The first group ought to contain geometric
characteristics that determine the external contour of the body. Parameters
in this group are continuous and represent the starting point for calculating
aerodynamic characteristics and, therefore, largely determine the ballistic
design parameters and the thermal stress for the tra~ectory, as well as the
flow regime around the body and heat exchange in the different zones of the
vehicle's surface. The second group should contain parameters that describe
the control of motion. Among them are the coefficients that shape the pro-
grammed change in the ar.gle of attack or roll, or indicators that determine
the instant of transition from one to another type of control, for example,
motion along a constant altitude, constant load factor, or constant tempera-
ture trajectory.
The thir3 group of design parameters includes the weight-power and design-
- layout characteristics of the descent vehicle. These are prima.rily dimen-
sions of the structural elements and of the therma.l protection system, pewer
_ plant parameters, and data that determine the location of compartments and
assemblies onboard the vehicle. Among these parameters can be those that
change discretely or assume only integer values.
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The situation becomes more complicated when, combined with an automated
_ search for configuration and structure of the descent vehicle, different
internal layout schemes and various arrangements ~or the instruments and
assemblies forming part of the onboard syste~ are reviewed.
The variables in such a problem of optimization are the configuration param-
eters whose integer values correspond to certain variants of layout or sChe-
matic solutions. In this case, the task of an automated synthesis of the
structure of the vehicle is formulated as a problem of discrete progr~ntming
for whose solution s~ecialized search methods are u~ed as described in
Ref. 23, which also contains the formulation of the problem of synthesizing
the structure of a space vehicle. In the same book, examples are taken from
the practice of space-vehicle design in order to explore the solution to the
problem of opt _mal design thrc�?gh the use of a system of automated search
and selection of configurations, descent tra3ectories, design, and weight-
_ power parameters for different types of descent vehicles. For this reason,
we will limit ourselves here to a descriptior. of the general formulation for
- the problem of searching for design parameters of descent vehicles. We will
present several of the most interesting applications of design calculations
and r~sults from s uch a search and show the chosen conditions of motion and `
conf iguration of future descent vehicles.
The co~or., most typical problem of search for geometric and trajectory
parameters is usualZy solved at the initial design stage and determines the
appearance of the descent vehicle. Only continuous parameters that belong
to the first and jecond groups are varied and their values, corresponding to
the minimum of a weight index, are determined. 1'his index can be the mass
of the vehicle or the mass of the descent and landing equipment. The ques-
- tion of criteria that determir~e the goodn~ss of the vehicle in terms of
weight will be posed below as part of the discussion on descent-vehicle pay-
load. The simultaneous optimizing of vehicle shape and of the conditions of
motion is usually formulated as a problem in nonlinear programming and is
solved with tt~~e aid of known nvmerical sear.ch methods based on a atepwise -
p~ogression through the parameter space to arrive at the optimum variant.
The values of the design parametera for a siescent ~Jehicle determine the state
vector of the system to be optimized:
- X(Xl, X2, . . . , Xi, . . . , Xj.l) .
To eacn value X~k) of this vector corresponds ~ value F~k) of the ob~ ect
function ~(X) and the values
yl ~k~ ; y2 tk~ , . . . y~ ~k~ , . . . ym~k~ -
of the fu.^.ctional constraints Y1 `X~ , Y2 ~ , . , . , y (X) , , , , , ym(X) . Here
the vector X is def ined over the set R, which is ct~nstrained due to the
rang e of variation o~ each variable determined by inequalities of the type:
Bi S~ (constraints uf the first type) (l.l)
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and also due to the functional constraints, thzt is, due to inequalities of
*_he type
Y~ (X) < y~ ~ ~ (constraints of the second type) (1. 2)
The goal of the optimization is to find the best value of the cr3terion
_ Fop t for XER. The goal of the search is broader and can result in finding
a whole family of suboptimal solutions. It is among these solutions, which
are close to optimum but differ from it by a number of additional properties
and attributes t~iat cannot be formulated, that the developer will look for
the best variant for the vehicle design applicable to the design limitations
established. By choosing a design variant at the current level of the state
of the art, the designer must know to what degree he departs from the
strictly optimum solution according to each criterion.
The general goal of design research is broader, and it results in determin-
j^g the relationships between opttmal or suboptimal solutions and the basic
initial data and cons*_raints.
In choosing descent-vehicle design parameters, it is necessary to deal with
a number of criteria that in part are contradictory. As a rule, one of them
determines the achievement of weight goals, a second may target a specific
problem, for exa.mple, the achievement of maximum cross range, a decrease in
load factor, or a decrease in the flight velocity at the end of the descent
tra3ectory before landing. The problem of searching for the best values of
descent-vehicle parameters is thus divided into a number of partial problems
of optimization, for each of which one of the best solutions for a given
criterion is determined with other criteria being considered as constraints.
By letting an optimum "drift" along fundamental constraints, the bounding
sets of optimum solution are determined, that is, the bounding relationships
of the best solutions in terms of each criterion are found while values of
the other criteria are held constant.
Among the criteria or constraints in problems ot searching for the best val-
~ ues of descent-vehicle parameters can be indices of weight or mass for the
whole vehicle, expenditures of the working medium from systems of attitude
control and braking, the value of characteristic velocity of the propulsion
systems, speed of flight at a given altitude, or altitude of flight aC a
given speed, load factor, drag, effect of load factor on the crew, cross
range or trajectory turn angle, flight range, temperature, specific heat
flows, or erosion of the surface layers at the heat-stressed points of che
vehicle, thickness or strength of the heat shield, time of flight, or time -
- of flight through the atmosphere. The choice of criteria and cons.traints is
determined by the specific search problem and depends on the particular mis-
sion of the vehicle. The mission of the vehicle determines the method for
= controlling motion, the type of structure, and class of shapes to be consid-
, er,~d in searching for the vehicle configuration.
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I~s spite of the brevity of the development period c+f space eugineering,
there exists at present a variety of descent vehicles, and this variety pro-
duces the need for some kind of classification. The authors do not pretend
_ to have an aIl-inclusive and definitive system for classifying descent vehi- -
cle attributes and grant that there are other ways o,~ ~rouping descent vehi- `
cles into classes and groups; in the future, one ought to expect a more sub-
sCantiated approach to the classification of descent vehicles. G
It is possible, f.or example, to group descent vehicles arbitrarily according
to the character of tlieir payloads, and this will be our first attribute for
their classification. The presence of man oii board the vehicle, or, more
correctly, the presence of systems that provide for his ability to live,
work, and participate in the control of the mission, substantially influ-
_ ences the composition of the payload, which then is reflected in the appear-
ance of the vehicle and in the trajectories of descent and landing. Descent
vehicles developed for automatic flight can be designeu for substantial load
_ factors, as a result of which the requirement for controlled descent and
?Zigh lift capability of the aerodynamic shape is removed.
The mission of the vehicle should probably be the second classifcation
attribute. The shape of the vehicle depends on the conditions of entry into
the atmosphere, on the parameters of the atmosphere and gravitational field
_ of the pl~net, and on the schedule of operations to be perforr~ed during
descent and landing.
An additionaJ. subdivision according to specific purpose and flight condi-
tions j.s possible for vehicles intended for operations close to a particular
planet. Descent veh3cles can enter the atmosphere either from an approach
trajectory with hyperbolic velocity or from a low orbit around the planet.
In either case, different requirements on the lifting properties of the
vehicle frequently arise. In addition, descent vehicies can either pene-
trate the atmosphere and deliver a payload, or dive through it and return to
an orbit around the planet or to a departure tra~ectory. In the tirst case,
delivery is made to either the surface, in which case a special landing gear
is used, or to a predetermined altltude in the atmosphere of the planet.
The external appearance of the vehicle, its configuration, and conditions of
motion in the atmosphere depend to a great extent on the manner in which the
lift capability of the aerodynamic shape is used. Here we have in mind the
maximum attainable lift-to-drag ratio, that is, the available lift-to-drag
ratio of a given shape for the assLIIned layout and centering characteristics.
In flight, the vehicle can be trimmed at some angle of attack to produce a
lift-to-drag ratio less than the available one. -
The rela.tionship between the lifting and braking properties will serve as a
third attribute for classification, according to which descent vehicles can
- arbitrarily be divided into five categories: Vehicles for ballistic descent
(with zero lift-to-drag ratio); vehicles with low lift-to-drag ratios
(K < 0.4); vehicles with modest lift-to-drag ratios (K = 0.4-0.7); vehicles
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with significant lifting capabilities (K = 0.7~1); and gl~.ding vehicles with
- high lif t-Co-drag ratios.
= Ballistic descent vehicles, in turn, can have different braking capabilities
and two limiting configurations: A shape with a large aerodynamic braking
device (as a rule, this is a blunt heat shield) or a pointed shape with ~{.gh
= midsection load factor.
There are many methods of classifying aerodynamic shapes having a modest to
lar ge lift-to-dr,:o ratio. The classes are obtained by means of transforming
var ious initial configurations to combine several bodies or various
conic and spheric shapes. The variety of all conceivable shapes tasks one's
ability to count them.
Conparisons of different classes of aerodynamtc shapes are useful only in
terms of applications to vehicles for specific missions at a given level of
the required lift-to-drag ratio.
One of the widespread means for classifying descent vehicles according to
the method of flight in the a~mosphere is based on the differences in con-
trol used during the main segment of the descent trajectory. Descent vehi-
cles can perform a noncontrolled ballistic descent as used in automated
vehicles; they can vary the amount of braking either gradua.lly or in steps
dur ing the flight, thus controlling the descent trajectory without requiring
lif t; they can, for a given constant lift that corresponds to the trim angle
of attack, change its pro~ection onto the vertical, by controlling the roll
angle, that is, by turning about the velocity vector; and, finally, they can
change the magnitude and direction of the lifting force by controlling the
angles of roll and attack during the flight. Roll control can be achieved
re3atively simply for blunt-sha.ped vehicles with a low value of the avail-
able lift, inasmuch as rotation about the velocity vector can be achieved
with gas jets and without changing the balance of the vehicle. Such a
des cent, which has been called "slipping," was used for the Soyuz and Apollo `
spacecraft.
And finally, the various types of descent vehicles can also be divided
according to structural attributes, which determine the principle of opera-
tio n of individual descent and landing systems.
Dep ending on the manner in which structure and thermal protection of descent
vehicles are used, they can be of the single- or multiple-use varieties,
with multiple-use descent vehicles, in turn, subdivided according to thermal
pro tection method ("hot structure," ablative thermal protection, thermal
pro tection using heat-sink or radiating screen, and so forth).
The landing systems of descent vehicles also influence their appearance and _
_ can be used to differentiate one from another. Thus, we differentiate
between devices for prelanding braking and shock-absorbing devices for soft
laadings. For the latter, use can be made of landing gears with different
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energy absorbers in the support mechanism, infiatable balloons, or soft-
landing thrusters, wY:tch provide a verrical impulse at the moment of contact
with the surface. The prelanding maneuver devices and braking systems can
consist of wj.ngs, liquid-propellant rockets, turbo3et or turbofan power
plants, rotor systems with different methods of dr~.ving the blades, para-
chute and paraballoon braking devices, and collapsible soft-envelope wings
- with either an infla.tabl.e or a rigid afrframe.
lfiere are many other ways for differentiating among descent vehicles by
structur.al features of individual systems. Let us now try to define general
~riteria ror comparing descent vehl.cles intended for the same mission, that
is, let us make the transition from the attributes that separate different
descent vehicles to the attributes that unite them.
One of the basic nondimensional criteria used in comparing descent vehicle
characteristics is the weight criterion for descent and landing systems.
'I'he widely-used form of the weight criterion for the descent and landing
system, namely, the ratio of the weight (mass) of the descent and landing
system, G~~~, to the total weight (mass) of the vehicle, G~, is not correct,
inasmuch as the total weight of the vehicle already includes the weight of
rhe descent and landing system. Only the ratio of the weight of the descent
and ianding system to a reference weight, fixed for a11 variants of deacei;t
vehicl::s, can objectively describe the weight improvement o~ the v~~hicle.
Such a reference weight could be the weight (or mass) of the payload, G~.
Thus, if for two vehicles the weight efficiency, that is, the ratio G~~~/GO'
is 50 percenr and 70 percent, and at first glance the difference is only
20 percenfi of the total weight, in actuality, for given uniform constraints
placed on the weight of the descent vehicle in terms of total weight of the
space system, we have to reduce the payload weight in the second vehicle by
nearly a factor of 2 compared to the first vehicle. The ratio
~CC1I = GCCII~GO (1.3)
G~~ ~ - GCCII ~GO
is 100 percent i.n. the first case and 2.33 percent in the second case. Thus,
for an identical payload weight fraction, the true difference in actual pay-
load weigtit is 60 percent.
The payload for a typical descent vehicle can be conveniently thought of as
cc-~sisting not only of the scientific equipment to be returned to Earth or
delivered to a planet but also all onboard systems, which remain fixed dur-
ing changes of. design parameters of the vehicle, that is, in changing from
one variant of the vehicle to another.
In this fashion, the payload includes all that does not ~orm part of the
descent and landing system. As a rule, the descent and landing system
includes the structure of the airtight cabin, the structure of the gl~der
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or aerodyna;tic braking equipment, the thermal protection and thermal ~nsula-
tion, propu~~.sion equipment ~or att~,tude control and prelanding braking, and
systems that provide �or a soft landing. The other onboard systems, namely,
the collection of life~support systems, systems for thermal control, power `
generation, and radio cammunication, and the instrumentation portion of the
control system all form part of the payload of the descent vehicle.
The systems mentioned are arranged in the body of the vehicle, and the man-
ner of distributing them in the various sections of a given configuration
~ determines the volwne characteristics and mass properties. One of the non-
dimensional crit~ria, which indirectly reflects the increase in vehicle
weight with an increase in its lifting c:apabilities, is defined by the fol-
lowing well-known formula:
KS = (4.85/S) ~VP` (1.4)
1~
where V is the available voiume of the vehicle (in terms of the external
outline~; ~ is the total surface of the vehicle. This criterion represents
the ratio of the surface of a given vehicle to the minimum possibls surrace
for the given volume and can be thought of as a kind of efficiency in using
the vehicle surface.
For. vehicles with a fineness ratio close to unity and c~rlthout wings or otl-:er
protruding surfaces, tbis criterion is close to unity. The typical descent
vehicle shape with small or moderate lift~to-drag ratios con~ists of two
bodies: A forward volume and an aft volume (Figure 1.1). The forward vol-
ume can be a segment or a blunted cone, either circular or elliptical. The
aft volume can be spherical or conical. Both volumes are 3oined at a co~unon
surface, which can arbitrarily be considered to be a plane, and all aerody-
namic characteristics can be referred to it. This plane, called the lifting
plane, is perpendicular to the plane of symmetry of the vehicle, in which,
as a rule, lie the vehicle center of mass and the center of balance of aero-
dynamic forces.
Aerodynamic forces used for braking and controlling the descent trajectory .
are basically generated by the surface of the forward body. In the hyper-
sonic regime of flow, aerodynamic forces are generated by the pressure in
the stream as it is decelerated by the forward part of the vehicle. The
front lifting surface is coated with a thermal protection ].ayer and forms
the heat shield or screen. Side surfaces of the aft body lie in the base
region (region of blanketing), require less thermal protection, and farm the
fairing about the payload. The tnain portion of the required volume of the
descent vehicle is formed by the volume of the aft body and its surface is
larger than the surface of the heat shield, which results in an advantage of
weight, inasmuch as the density of a unit of surface of the heat shield is
two to three times higher than the density of a unit of surface of the base
- fairing.
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The centering character~,stics ~or this type o~ vehi,cle can he compaxed to
the layaut parameters by using the nondizqensional criterion Kv, whi.ch is
determined ~or the design tri~u condit~,on and ~or the corresponding location
~ of the resultant aerodynamic foxce. The criterion used is �the ratio between
the minimum possible volume that can ~ie cut from the vehicle by a plane in
which the resultant force lies and one-half of the total available volume. _
The minimum possible volume is bounded for this situation by the surface of
the vehicle on the u>>stream si,de and by a plane containing the resultant
force acting normal the surface of symmetry of the vehicle.
The vehicle's center of mass for the design balance condition 1.s located
along the resultant aerodynamic force. If, by using active devices for
stabil3zation an~ control, the attitude of the vehicle relative to the
- velocity vector can be controlled before entry into the atmosphere, ther_, as
a rule, there is no requirement on where the center of mass is along the
resultant. TY?e same has to hold in the case of descent vehicles whose
motion is controlled in accordance with the tasks set.
In any case, whether control is used for redLCing the cumulative effect of -
- deceleration to an allowable level (.descent to Earth), for shaping a shallow
trajectory for approaching the surface with tninimum velocity (descent to
Mars), ur the injection of the vehicle into the descent regime at the high-
est possible altitude (descent into the atmosphere o~ Venus), the design
conditions for the fli.ght are determined by the method chosen for control-
ling descent. Any deviation from the desired trajectory due to an off-
design entry into the atmosphere by the vehicle with an arbitrary angle of
attack has to be avoided, as this wi].1 lead to failure to achieve the basic
flight objectives. A direct consequence of such an approach in choosing the
conditions of motion is the main requirement on the location of the center
of gravity: Its location has to be along the resultant aerodynamic force
- for the design angle of attack, and for insignificant deviat~ons from this
angle of attack the vehicle has to be statically stab~e. Let us agree that
by narmal direction we will mean the direction pdra11e1 to the lifting sur- _
face of the vehicle and by longitudinal direction we will mean the direction
perpendic~ilar to the lifting surface.
Thus, the longitudinal movement of the center of gravity is, as a rule,
unlimited for angles of attack close to the design value. The matter stands
differently with displacement of the center of gravity in the normal direc-
tion, because here we have to relate the centering of the vehicle to the
location of the resultar~t aerodynamic force, which is almost perpendicular
to the lifting surface. For vehicles with a moderate lift-to-drag ratio,
such a displacement of the center can cause certain difficulties. If the
surface, which passes throci~h the resultant, divides the vehicle volume in
half, then the necessary centering can be achieved without any weight loss.
On the other har.d, depending on the degree to which the required location
for the resultant is moved away from the geometric center of the volume of
the vehicle, layout difficulties arise and the ballast weight goes up. The
criteria stated above indirectly reflect the possibilities that exist .for
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some forms of vehicles to use the internal volume to attain impzoved lift-to- _
drag ratios. This is parti.cularly important for those cases where the
required magni,tude of the lift~to~-drag ratio approaches 0.5 to 0.7. Achiev-
ing such a lift-to-drag ratio with traditional segmented conic forms, which
retain a good ratio of surface to volume and a relatively small surface of
the heat shield, is diff icult for a number of reasons, among them the dis-
placement of the resultant and the attendant poor use of the volume while
maintaining the necessary lateral centering.
To improve further the shape of controlled vehicles for entry into the atmos-
pheres of Earth, Venus, and Jupiter at hyperbolic velocities, let us recall
the relationsnip between convective and radiative heat flows and the require-
ment for more pointed heat shields. In such c3rc~rmstances, we have to con-
sider shapes with conic heat shields in a zeia angle-of-attack flow.
By replacing the se~-mented shield of traditional segmented-conic shapes in a
zero angle-of-attack flow (Figures 1.2, 1.3) by a shield shaped like an
obliquely-cut elliptical cone, a shape will rer~ult that has the same volume
and surface as the segmented cone but with a number of advantages. The main
advantages are: Increased lift-to-drag ratios for given angles of attack;
smaller angles of attack of the lifting surface for equal lift-to-drag
ratios; lower structural weight for the lifting envelope of the heat shield
and lower weight of the thermal protection system in those cases where the
radiative heat flows from the shock layer exceed the convective ones (for
vehicles entering the atmospheres of Earth, Venus, and Jupiter at hyperbolic
velocity); and a more advantageous l.ocat~!on of the resultant aerodynamic
force fra~ the point of view of the required lateral centering.
The latter circwnstance becomes especially important in those cases where it
is necessary to achieve a higher value of the lift-to-drag ratio (K = 0.7 to
0.8). Obta3ning such ratios in a shape with a segmented shield (Figure 1.4)
is difficult, duE to the rigid requirements on the locaGion of the center of
gravity, which h~s to follow the resultant in moving to the upstream side of
the generatrix of the payload fairing.
The centering advantages of the shapes considered can be seen in Figure 1.2,
where the initial and transformed shapes and the locatior. of the resultant ~
for various angles of attack, corresponding to lift~to-drag ratios of 0.3
and 0.4, are shown.
These advantages are even more apparent for shapes intended for achieving
higher values of the lif t-to-drag ratio (K = 0.7 to 0.8). One of the pos-
sible nonsymmetrical segmented-conic shapes ansi a bi-conic shape are shown
in Figure 1.4. They were generated by substituting an obliquely-cut ellip-
tical conic shield for the segmented heat shield.
The curves in Figures 1.3 and 1.5 show the aerodynamic characteristics for
different angles of attack of the initial and resultant shapes and demon-
strate the first two advantages of the new shapes. We must mention another "
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advantage of the bi.-conic form, in urhich the axis of the ~orward cone, which
forms the surface of the heat shield, is parallel to the velocity vector of -
- the impinging f1ow. The heat shield is in a flow at zerc+ angle of attack,
and for such a symmetrical flo~t it is fairly easy to calculate the aerody-
namic characteristics and the parameters that de~ermine the heat transfer
at the surface of the heat shield. For vehicles that have to fly in night
_ conditions and particularly for those entering the atmospheres of Earth and
planets at hyperbolic velocities, the latter advantage is considerable.
Zero angle-of-attack flow i.mplies that the forward cone has to be elliptical
with its oblique cut either elliptical or circular, as, for example, in the
shane shown in Figure 1.2. I~'airings in the base regions of vehicles are, as
a rule, circular cones. Thus, these difficultiea in production engineering
of the fuselage and applicarion of thermal prote~tion surfaces a�re peculiar
only to heat shie.lds w:ith the new shapes.
Neither can we remain silent about such a property of the above bi-conic
shapes as their ability to preserve centering characteristics while the
lizt-to-drag ratio is increased. The derivative~ dK/da and dIC/dYT for these
shapes are significantly larger tha.n for segmented-conic ones. Tf we con-
sider the influence of the scatter in centering properties on aerodynam3c
characteristics of the finished product, then the mentioned proper_ty has to
be classified as a shortcoming. If, however, we consider the system for
controlling centering before f light as part of the flight process, or if we
consider a property such as adaptation of shape and layout to changing
regimes of halancing (as, for example, for increasing the ava~lable lift-to-
drag ratio) as important to the design process, then the property indicated
has tn be considered an advantage.
To avoid repeating descent trajectory calculations when changing the mass or
dinensional properties of a vehicle without changing the vehicle shape, it
is useful to develop a method that permits the determination of the tra~ec-
tory characterist~.cs without having t~ integrate the equations of mot3on, if
the descent control is known only in general. Such estimates are also ~
neede.d in problems of optimizing vehicle shape when the type of descent tra-
jectory is known, that is, the variation of load factor is known as a func-
tion of time or as a function of flight velocity. Tn the above-mentioned
problens of optimization for each variant oi descent vehicle considered, a
design calculation is done that culminates in estima~ing th~ weight charac- `
teristics of the descent and landing system for different constraints. An
iterative procedure is used to determine the weight during design calcula-
tions, because, in order to determine weight, it is necessary to know the
descent tra~ectory, which, in turn, makes it necessary to know the loading
on the lifting surface, which, in turn again, depends on the weight of the
vehicle. Also, the computational cases far ttie majority of descent and
landing syste~m,~ consider t:~e e~trzme conditions of flight. The determination
of extreme conditions of load factor, heat flow, temperature, time, and range
of flight is dependent on a series of iterations, 3n each of which a specific -
descent tra~ectory is calculated.
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The material presented above di,ctates the basic need ~or ~nethods ta pe~�mit
economical use of machine ti~ne and approxi~uate determination o~ tra,jectory
parameters. Known methods (~tefs. 6, 20, 40~ are not conyenient .�or design
purposes, as they do not give values ~or thermal flow loading along the tra- ~
jectory but only at the end o~ the calculat3,on. This removes the basic
advantage of such methods for purposes of design search; specifically, the
~~pportunity is lost to proceed in these calculat~,nns directly to the ~xtreme
conditions of flight in terms vf load factor and heat transfer. In the pro- _
posed method of calculating motion, the type of trajectory is given by a
combination of control laws that include constant load-factor laws; programs
_ that, by followi~~g the given relationship between load factor and flight
velocity, can shape a constant-altitude trajectory or assure that the tra-
jectory will remain within gi~*en constraints of temperat�ire and flight
range.
The unified sequence of control laws defines a whole set of trajectories,
and each tr ajectury in the set differs only by one design-ballistic param- -
eter, namely, the loading on the lifting surface. The independent variable
_ is flight velocity and, by specifying the switching instants (on the veloc-
- ity scale) makes it possible to combine all possible trajcctories into a
single class and to establish general rangea for controlling interactions.
It is suff icient to obtain the variation of speed and altitude of flight as
a function of time in order to estimate the parameters of motion, loading,
heat flow, and temperatures, and, consequently, ablation and heating of the
. thermal pro tection surface.
By integrating the four basic equations of motion, we can determine the -
_ indicated relationships with a precision sufficient for design calculations.
These equations based on generally accepted assumptions (Refs. 1, 6) and
written in the velocity coordinates in a unified form suitable for any
planet are
dV Qv2 g3-gn SIA e~ ' ~1. J~
dt 2P=
dt 2Ps g3 ~ ga COS 9(g Rn 1; 6~
\ /
dN _V sin e; ~ (1. 7)
~ dt
dL _ V cas 6, h~ . 8)
df
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where V i~ the flight velocity; A is the ~light~path angle o;ith respect to
the horizon; g3 is the free-fall a c celeration at the Earth's surface; g is
the free-faJ.l acceleration ~t the altitude o~ ~light in tine atmosphere o~ a -
given planet; R i~ the distance from the center of the planet to the alti-
tude at which t~e vehic2e is moving; H is the altitude of flight; L ie the
range of flight; p is the density of the atmosphere, evalua*_ed at the given
altitude of flight; Px is the loading on the li~fting surface. Here
PX = GJ(CXS) or PX = mg3/(CXS1 where G is the total weight of the vehicle;
CX is the drag; S is the area of lif ting surface; and m is the mass of the
vehicle.
- It is assumed that the law according to wh~.ch density varies with altitude
of flight is known and that by determining the density we have an estimate
of altitude.
The magnitude of the effective lift- to-drag ratio K at any instant of time
determines the control of descent. For a constant angle of attack, the load-
ing PX of the vehicle practically do es not change along the trajectory and.
the magnitude of the effective lif t- to-drag ratio is determined by the roll
angle y and the available lif.t--to-drag ratio of the vehicle, ICp, namely,
K = Kp cos y (1.9)
As wi11 be shown in later chapters, the preferable trajectories of motion of
future descent vehicles entering the atmospheres of Earth, Mars, and Venus
will include leveling-o.ff regimes wi th constant altitude or constant load-
tactor plateaus. Transition to the leveling-off regime is achieved after
the atmospheric entry segment, that is, after achiev3ng maximum load factor
or heat-shield surface temperature. Control during the atmospheric entry
phase (permitting flexible res~onse to changes in entry c~nditions and the
attendant change in vehicle characteristics) uses control algorithms or a
collection of programmed relationships between load fa~tor and time or
be*_ween load factor and flight veloc ity, in order to bring the controlled `
object to within established load fa ctor and temperature bounds.
In the simplest case, the atmospheric entry phase can be modeled by a linear
relationship between lcad factor and time. By assuming the angle of entry
_ into the atmosphere to be small (controlled descent), we assume 6~ 0; `
cos 8 x 1 and sin 8~ 0 in the equations of motion and obtain the following
relationships for the flight phase with a constant rate of increase in load
factor:
' V=V,x- 1 CT2 1 ; (1.10)
2 jl1+KP
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nr m~x 2P~. ,C, - ~ 1. ~ ~ ~ -
~roY 1+KN v2
. n~max ~~.1~~
n~ = Qt t~
�
L= VBx,~ - ~ ( l. l 31
6v.~ +Kn
where VBX is the velocity of entry into tne atmosphere; c=(nE ,~~/aTB)g3
is the gradient of the change in acceleration during braking; n~ ~X is tFie
maximum load factor characteristic for the given entry corridor and the
available lift-to-drag ratio of the vehicle; OTB is the interval of time
from the instant of entering the atmosphere (load factor of 0.05 to 0.1 g's)
to the instant of experi encing maximum load factor. If, during the period
of atmospheric entry, a flexible algorithm is used that permits obtaining a
general relationship between the required accelerat3on and the apparent
velocity for all conditions of entry, then a maneuver will result that pro-
vides f,:r capture of the vehicle (for motion along the upper boundary of the
entry corridor) and inje ction of the vehicle into a permissible level in
terms of load factor or temperature. Let us assume that we are following a
linear relationship between the longitudinal acceleration VX and the appar-
ent velocity (VBX - V), that is, a relationship of the type
VX = a + b (VBX - V) (1.14) -
This relationship for the longitudinal acceleration is now substituted into
Equation 1.5.
The nominal, or reference, tra~ectory is common for all conditions of entry,
and in such a case is d e termined by the following equations, o~tained by
integrating the equations of motion with the condition of Equation 1.14 and
using the same assumptions as in the preceding case:
V~VB- b (e�~- (1. 15)
2Pxae~
Q= a 2 ; (1.16)
. ~3 r Vs= _ b~ e~t _ 1~ I �
L ~
n.=[a j-b (V~-V)] g3 ; (1'. 17)
Y~ +KP
t= Vu'C - b.~ - b~ ea~ b2 , ~ ( I. 18)
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Immediately after passing the maximuan load factor, that ~s, at the instant
the vehiclQ ie in a state to move iurther by skipping with a maximum effec-
tive lift-to-drag ratio, the tra~ectory augle approaches zero and then
changes sign and becomes positive.
The time interval from the moment of entry into the atmosphere to the
instant of passing peak load factor will be designated by ~TB, and flight
velocity at this same instant I~y VB. The constant altitude regime of level-
ing off should be started at this time, ds the control program generates an
effective lift-to-drag ratio starting wtth a zero value for the derivative
of the tra~ectory flight-path angle. Taking tliis into account, we can inte-
grate Equation 1.5 by assumin$ the density of the atmosphere to be constant.
We ob~ain the following expression for the velocity of flight during the
constant-altitude phase:
V= ~~1/VH) +~Pg3/2P~) CT ~ ~TB) l^1 C1.19) -
The density of the atmosphere during the constant-altitude phase of flight
equals
- p=(2PX'~e max~/ CVg2 1~ + I~`) (1.2U)
Load factor is determined from the expression
nE � CnE ~V2)/VB2 (1.21)
The increase in flight range along the tra~ectory is determined by using the
formula
oL =L - L- 2P in Qg + eg - ln 1
� Q83 ~Vw QPs � 2Px ) V~ ] (1.22)
Motion along a conatant-temperature tra~ �cto;-y usually begins when the lim-
iting temperature is attained. We will coi.sider that at the instant of -
transition to the isotherm the flight time is ~TB and tlie velocity is VB.
The condition that the equilibrium temperature re~main constant means that
along this trajectory there is preserved a constant heat flow in the ther-
mally stressed zone of the vehicle. A formula like the following one is
usually used to estimate the magnitude of the convective heat flow:
9 = ~qP~n (1.23)
We will assume that the magnitude of he.at flow will be maintained at the
attained value q~. With the control system providing for motion along the
isotherm, the density ~f the atmosphere in this case will depend or. flight
velocity in the following fashion:
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t
' e-(CJ)m in � (1.24)
9 vm
Integration of tize equation of ~otivn (Equation 1.5~, with the condition of
Equation 1.24 and assum~ng tt~t the ~li;gIit~patFi angle ~s zero, yields the
following relat~onsTiip betareen fl~gfit velocity and time:
R-m 1 m
m n- n: Q~ m ~3 a-m
V- VB ~(cQ ~ 2PX (~c- n~r~ . C1. 25)
Load factor as a function of time for a constant~temperature tra~ ectory var-
= ies in the following fashion:
i
11s=~]-~Kp( ~m Zp X ~
~ 9 x
n-rn 1 0~_~
n- m qq m 83 n-m
X VB - m ~CQ ) 2Fx - ~i,l � (1. 26)
The range for flight along a constant-temperature segment of the tra~ectory
can be determined approximately By tfie following formula:
eL~L-L,~- _ _ 1 _ X.
~ .
, � n-�: 4o m g3 in '
- rrc (Cq , 2P,~ Cn - rn + 11
/
R-^~ 1 n n
J m
X 1 V. -n mm f C� l~` p~.Y-~Ys) n m--Vs ~ l. 27
� ~ )
~ \ a 1 .
The constant-load~factor segment begins after the point of leveling off or
following flight in the constant-altitude regime. Departure from the iso-
therm can also be achieved using a constant load~factor tra~ectory. Assum-
- ing that the flight--patfl angle (faetween the velocity vector and the local
horizontal) is small and assuming a load factor much larger than the
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_ magnitude of g~ si;n 6~ ate ohtain the approximate expressi,on ~or the tra~ ec~
r_ory parameters for a constant~load~factor platesu;
V.~ Vx- ,/1 +K2 S3 ~~-~Ta)~ - (1..28)
Y p
' 2PXn~-
; (1. 29)
- yi +Kpv~ .
aL-L-L�.~ V�(~-~~c�)- ga~T-Otp)~, (1.30) -
1+Kp
where Lg, Vx, and QTg are the initial values o~ distance, velocity, and time
o� flight corresponding to the inatant of transition to constant~load~factor
' flight. '
All the above relationships are characteristic f,or different segments of con-
controlled descent, wlth the control system using the lift characteristics
of the vehicle, shaping a sl~llow tra~ectory, and thus impleznenting a law
for varying the load factor as a function of fligh~ velocity.
With the variation of velocity and tlight altitude known, it is possible to
estimate heat flow and, hence, the requtred mass of the ablat3ve thermal
_ protection layer.
During the preliminary design stage, when a number o� approximate models are
used, specific heat flows are computed by correlation formula~. Tn the gen-
eral case, correlation relationships for estimating heat flows have the form
~
~ 9.=k~Q=V mR"`; (1.31)
9K=kKQ�V�R", . (1.32)
where q and q~ are the specific rad iative and convective heat flows at the
stagnat~on point of the vehicle; R is the radius o# bluntness.
During entry into the atmospfieres of Venus and Earth at hyperbol~c veloci- -
ties, radtative heat flows are crnaparable to convective ones, while during
entrp into the atmosphere of Jupiter, radiative flows predominate. Under
such conditions, the esti~at~on of the optimum radius of bluntness is of
some interest at the stage of qualita tive analysis o~ braking in the atmos-
phere. Finding th,e opt~mum xadius o� ~ilun~ness is important ~or that class
of tra~ ectories tfiat fiave been opti~zed by using the criterion de~ining the
special problem whose solution determines the conditions of motion. For
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entry into Earth's atmosphere, this problem turns out to be (as will be shown
in the next chapter) the search ~or a minimum integral of load factor; for
entry into the atmosphere of Venus, it is the completion of braking in the
upper layers of the atmosphere. Tn this and other cases, trajectories with
constant-altitude plateaus are preferred, followed by further shallow desc:ent
along a constant-load-factor trajectory. Parameters of tra~ectories of this ~
- type can be determined without integrating the equations of motion; there-
- fore, for vehiclns with variable loading on the lifting surfaces, it is pos-
sible to estimate an optimum radius of bluntness for different conditions of
entry into an atmosphere.
The expression for the optimum radius of bluntness is derived from the condi-
tion
dR ~Qr+QK)=~� (1.33)
where Qr and QK are the integrated heat flows at the stagnation point.
By considering Equations 1.31 and 1.32, we can determine the optimum radius
of bluntness by using the unit integral heat flows, that is, heat flows at the
stagnation point with radius 1 m:
TN � m .
~ kKQ`� V ~ d~ n .
m
Ro~,c = - n _K . (1. 34)
f krQ~ V~, dt ~
- o
In Figure 1.6, the reader can see a typical variation of the optimum radius
of bluntness as a function of speed of entry into the Earth's atmosphere for
vehicles with different loading on a lifting surface with the available lift-
to-drag ratio equal to 0.6. The curves have been constructed for trajecto-
ries corresponding to the lower bounda of the entry corridor with the con-
straint that after passing through the maximum load factor the conditions of
motion correspond to the minimal integrated load factor effect (Chapter 2).
Let us note that the magnitude of the optimum radius of bluntness determined
using Equation 1.34 does in no way imply the need to make the vehicle with
that exact radius of bluntness but only allows us to assess the correlation
of the integral heat flows at the stagnation point for the given type of
descent tra~ectory.
The loss in the total heat flow depends on the optimum radius of bluntness,
which is borne out by the curve in Figure 1.7, showing in relative units the
magnitude of the loss in the total integrated heat flows for deviations from
the optimal radius.
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The true value of the optimal radiua o� bluntness has to be determined by
r.he masa of the thermal protection material to be ablated and heated at the
stagnation point of the hea~ ahield and its vicinity, while also consider-
ing the masa of the structure that supports the heat shield.
The ma$nitude of the radius of bluntness has,, as a rule, little influence on
- the mass of the vehicle. Here the deciding factors are the dimensions of
~he surface of the heat shield and the general curvature of its surface with
respect to the local angles of attack, inasmuch as they determine, on one
hand, the drag of the vehicle, its lifting characteristics, and, therefore,
the thermal stress of the descent tra~ectory, and, on the other hand, the
mass of the entire shield in ternLS of the mass-density of a square meter of
area.
R R
3
r~~ ~6~ ~ ~ 6-
~ w.~ ~
R,
. / ~ ~ 4.
2~ 5
Figure 1.1 General View of Descent Vehicle with Modest Lift-to-Drag Ratio
Key:
l. Heat shield
2. Forward volume
3. Payload fairing
4. Af t volume
5. Lifting surface
6. Center of mass
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~ - ~Z~ . -
~ N-0,5; a�~~1>�
~ V~ `,so X-0,3; aM,.~�19�
_ 60' (y) ,
- ( 2 ) axe~ _ _ .
1
. ' ` ~
; K~07i a,vec'1S~ ,
~ fl-O,Si a,vu"l9� ,
` 2so (2)
Figure 1.2. External Contours and the Location of the Resultant Aerodynamic
Force of Segmented-Conic and Bi-Conic Shapes of Vehicles with -
Small Lift-to-Drag Ratio
Key :
1. Center-of-mass locations
2. Lifting surface angle of attack
Vehicle with segmented heat shield
Vehicle with heat shield in the shape of an obliquely-cut cone
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C ,
~
OfS ~s
~ ~ ,K
- ~ /
05 ~ ~ ~
, ~ \ .
/
~~X
0,25 ~f ~
. ~ ~ D io z0 JO ~o ' ~ i )
_ Figure 1.3. Aerodynamic Characteristics of Segmented-Conic and Bi-Conic
Shapes at Various Angles of Attack for Vehicles with Small
Lift-to-Drag Ratio
Key :
' Vehicle with segmented heat shield
Vehicle with heat shield in the shape of an obliquely-cut cone
1. I,ifting surface angle of attack
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(2 ~ .
� K-0,97; aN~,.-50'
' K-0,5f; aNer-JO� K�0,78; aNaryO~ _
lf-0,71; aMrr-40�,~{i~ ~0,56; a,,,�-JO' ~Z ~
. /
/
SO' 45' ~o
_a \ j~ ,l~1~ - ` �
. \ 8U,
~Z ~ aNer \ 65'
~
Figure 1.4. External Contours and the Location of the Resultant Aerodynamic
Force of Segmented-Conic and Bi~-Conic Shapes of Veh3cles with a
Modest Lift-to-Drag Ratio -
Key:
1. Center-of-mass locations
2. Lift3ng surface angle of attack
Vehicle with segmented heat shield
Vehicle with heat shield in the shape of an obliquely-cut cone
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K Cf ~
\ .
~
'
~ ~K
~ r
~
/ ~ ~
s ~S / ~
~
�
ZS JO 35 40 45 Sa aM� (1 ~ ~
~ .
Figure 1.5. Aerodynami c Characteristics of Segmented-Conic and Bi-Conic
Shapes at Various Angles of A.tCack for Vehicles with Modest
Lif t-to-Drag Ratio
Key: .
Vehicle with s egmented heat shield
- Vehicle with heat shield in the shape of an obliquely-cut cone
1. Lifting surface angle of attack
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R, H . -
1,5
I 2 ~o . -
~ S ~ P,~ -10000N/nf ( t )
. ~ _
03 P'~6000N,yt
� ' (1
0
>1 12 /J >S~ IS !6 Vo, nn/c
~2)
Figure 1.6. Dependence of the Optimum Radius of $luntness on the Speed ~f ~
Entering into the Earth's Atmosphere
Key:
1. Loading P~ on the lifting surface (N/m2)
2. Vp (lan/s)
d~
0,8
- _
0,4
0,2 ~ ~
~ f 1 3 k k
. ,
Figure 1.7. Dependence of Total Heat F1ow Loss on the Radius of Bluntness
(Q is the heat flow in the bluntness region of radius R;
R = R/Ropt; AQ = (Q - Qopt)/Qopt)
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CHAPTER TWO
PROBLII~IS CHARACTERISTIC FOR TFiE RETURN TO EARTH FROM NEAR-EARTH
AND INTERPLANETARY SPACE FLIGHTS
Most questiona related to the development of future descent vehiclea deal `
with problems of descent through the Earth's atmosphere. It is in this
atmosphere and in near-Earth space tha t the majority of new engineering
solutions are verified.
. The return to Earth for interplanetary, piloted flights is usually related
to expeditions involving flight around Mars and Venus, with automated sound-
ing of thes~ planets or landing on the surface of Mars. The velocity of
entry into Farth's atmosphere exceeds escape velocity and reaches 16 to
17 km/s. By studying Figure 2.1, we can see how the required magnitude of
the lift-to-drag ratio I~ varies with changes in entry velocity (for a con-
stant maximum total load factor of 10 g's). We also see the variation in
magnitude of the maximum load factor n~X for the vehicle (for several con-
stant values of the lift-to-drag ratio K). Finally, we also see that the
lift capability of the vehicle pxovides a corridor of entry into the atmos-
phere with a height of 20 lan. The following conclusion can be drawn from
calculating the limiting descent tra3ec tories, as generalized in the graph:
If we use current estimates of navigation accuracy in the planetary
approach portion of the tra~ectory and if we want to use traditional descent
principles (vehicle of modest lift-to-drag ratio with roll control), then we
run into load factors during descent who se maximum value will be the same as
wa3 experienced during the firat flights into space (8-12 g's), with the
time of braking in the at~osphere increased by a factor of 2.
If we now turn to the relationship between maximum load factor and time, as,
for example, shown in Reference 12 and in Figure 2.2, we can determine by
using simple estimates the maximum load-factor effect that can be tolerated
by the crew for optimum posture of each crew member with respect to the -
acceleration vector (~78� to the spinal column in a chest-to-back direction).
The situation is aggrevated by the fact that the crew will ha.ve become decon-
ditioned due to prolonged flight under weightlessness. In such conditions,
the load-factor effect on the crew exer c ises an essential influence on the
choice of the regime for braking in the atmosphere, and to determine the cri-
teria that model the load-factor effect becomes one of the basic tasks of
ballistic design research.
Let us first consider those general cri t eria that describe the maximum per-
missible level of acceleration. Having seen the relationship between load
factor and the limiting time of action (shown in Figure 2.2), we note that
the limi~ing time that man can sustain a certain load factor increases
sharply with a decrease in load factor. For example, by reducing the load
factor by a factor of 2, from 10 down to 5 g's, the time of tolerance
increases from 115 to 650 s, that is, approximately by a factor of 5 to 6.
An even more striking change is obtained for a further reduction of the load -
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factor below the level characteristic for the segqent of entry into the
atmosphere. It i,s logical to assume that, by having reduced the load factor
during the basic stage o� braking yet having pxeseryed its waxi~uum value at
the moment of entry into the atmosphere, that ts, hy maintaining the speci-
fied corridor of entry for the vehicle with modest lift-to-drag ratio, we
will lower the cumulative load-factor effect on the crew.
The quantitative side of the cumulative load-factor effect can be conven-
iently estimated by introducing the concept of a maximum (permissible)
effect impulse J-~P, which is determined as the product of the load factor
acting and the limiting time corresponding to this load factor. Determining
th~ effect impulse makes it opossible to construct a system of nondimen-
sional cumu].ative load-factor criteria. -
Let us turn to Figure 2.3, which shows two curves for the maximum-effect
impulse as a function of load factor, with the lower curve corresponding to
a body that has been weakened by the effect of weightl~essness. These curves
show an interesting relationship, namely: As the velocity of entry into the
atmosphere increases, the maximum permissible load factor, that is, the load
factor that can be sustained by a body, decreases. Indeed, the dashed hori-
zontal lines in the figure define levels of different required effect
impulses that correspond to velocities of entry into the atmosphere of 12,
15, and 17 km/s. The points where the horizontal lines cross the curves of
maximum impulse give the maximum load factor for the condition that braking
is occurring with a constant load factor of this magnitude, that is, for the
~ co:~dition of isoload-factor braking. It is easy to establish that the maxi-
mum load tactor for a pilot in top condition drops in this case from 10 g's
for a velocity of entry into the atmosphere of 12 km/s down to 8 g's for an
entry velocity of 17 km/s (for vehicles that have an available lift-to-drag
ratio of 0.5).
The example considered once again shows that in estimating load-factor
effects it is not enough to operate solely on the basis of load-factor mag-
nitude but that it is necessary to introduced criteria that, based on the
maximum impulse, determine the magnitude of the cumulative load-factor
effect.
The relationship between maximum impulse and load-factor magnitude for the _
- optimum posture of the pilot can be approximated by the following formulas:
1. For 5< n< 8.67
JnP = 14,000/n - 180 (n - 8) (2,1)
2. For 8.67 < n< 16
JnP = 14,000/n - 120 (2.2)
- 3. For16 1, the effect of the load factor will
lead to irrev~rsible changes in the body.
By summing the fractions m of load-factor effects for the separate segments
of the trajectory and then integrating in the limit the ratio between load
factor and the corresponding maximum permissibl~ impulse, we obtain a nondi-
mensional criterion which is termed the "measure of the cumulative load-fac-
tor effect":
`SK n
M- J ~~~P (n~ dt. (2.5)
0
If, for the given method of braking in the atmosphere, M< 1, then we assume
that the cumulative effect is less than the limiting one. This criterion is
described in Ref~rence 11. The criterion proposed in a number of works that
does not use the concept of a maximum permissible impulse (Ref. 18) has the
same physical mear~ing.
The concept of an effect impulse is methodologically justified inasmuch as
~ tied to it is a construction of more complex criteria that consider the pre-
history of the loading process. The matter rests on the fact that the meas-
ure of the cumulative effect mechanistically describes the pxocess of accu-
mulating the physiological changes in the human body, as if summing the
power of the effect of the loading on the body organs, which are under
stress from reacting to the forces of inertia and from external pressure
applied by the seat lodgments. Also, this criterion reflects the phenomenon _
of accumulation of load-factor effects onesidedly, without accounting for
the prehistory of the loading process nor for the effect of other load fac-
tors on the body. The acceleration effect process is distributed over time,
and it is quite significant during which period of loading the body is sub-
jected ~o the effect of the maximum load factor. We can assume that the
form of the load-factor envelope as a function of time influences the
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cumulative ~oad-factor effect. Peak loads at the end of the descent ~ra~ec-
tory or repeated peak loads will produce a more severe re~ction than the
initial loa3-f actor effects at the moment of entry into the atmosphere. The
prehistory of the loading process can, to some degree, be accounted for with
the aid of a nondimensional criterion that represents the maximum of the
ratio between the integral of load factor and the maximum permissible effect
impulse corres ponding to that load factor. The maximum value of thie ratio
is chosen from all loads computed in the process. Let us call this crite-
rion the "degree of loading due to load factor impulse," and we.will deter-
mine it by using the formula
^K .
~ nd~G
CH~~=1T[8X ~n n) . (2.6)
P~
For ~J~ > 1 we shonld assume that fihe body has been overstressed, and
that~or CH~J~ < 1 the degree of residual viability C~~J~ as a r~~ult of
act~Qn of the load-factor impulse is logically estimated as
~O( ) = 1 - CgCJ).
We should note that the criterion considered only incompleteiy accounts for
the loading prehistory, because no information on load-factor gradients nor
on how the maximum permissible load factors were reached is used. There-
fore, in choos ing the descent tra~ectory by using the cumulative load-fac-
tor criterion, the degree of loading is esimated simultaneously with the
checking of the maximum permissible load-factor levels, which supplements
the basic criterion, which is the measure of the cumulative effect. Deter-
mining trajectories that are optimal in terms of load-factor effect is car-
ried out by using numerical search methods and with the optimization problem
formulated as a problem in nonlinear programming.
Constraints for this search are: The limiting values of load factor and its
gradient, heat flows, ablation of the thermal protection layer, altitude,
and speed of f light. These parameters are varied in the course of the
search and det ermine (using apparent flight velocity as the independent
variable) the instants of switching from one ro11 angle to another or the
instants of swi tching to a different load-factor profile, to be followed by
the control system.
'Itao trajectories of atmospheric entry at hyperbotic velocity are shown in
Figure 2.4 using altitude and nondimensional flight velocity as the coordi-
nates. One of the trajectories (indicated by the dashed line), which was
the initial one in the search, has a short constant-load-factor segment of
braking at the maximum permissible load factor, which equals the maximum
during atmospheric entry. The figure shows how the cumulative effect grows
during the flight. The initial trajectory is optimal in terms of ablative
mass and mass of the heat-sink layer. The search computations yielded the
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trajectory shown by solid lines, that is, the traj ectory with an extended
constant-altitude regime of leveling off, which p ermits lowering the cumu-
lative load-factor effect without skipping out of the atmosphere, that is,
without a substantial increase in range and thermal protection mass, while
retaining their maximum values at the instant of entry into the atmosphere.
Results of moving the optimum along one of the constraints are shown in
Figure 2.5, namely, along the constraint placed on the angle of atmospheric
entry for a descent trajectory for returns from lunar flights or from high-
altitude near-Earth orbits. The functional relati onship shown represents a
family of optimal solutions, and it is possible to determine the limiting
angle of entry for a given vehicle with optimal control, that is, the maxi-
mum entry angle corresponding to the maximum permissible value of the cumu-
lative-effect measure.
We have thus considered the following criteria: The measure of the cumula-
tive load-factor effect as used in choosing the trajectory only under con-
ditions that the constraints on load factor itself and on its gradient are
known. Information on nearing the limiting values of load factor and its
time derivative supplement the integral criteria and permit the construction
of a system of complex indicators for describing the viability of a body
subjected to the effect of changing accelerations.
From a physiological point of view, we can subdivide acceleration effects
arbitrarily as follows: (1) Effect on bone-muscle fibers; (2) effect on
liquid-filled organs (brain, liver, stomach); (3) effect on blood vessels
(changes in hydrostatic pressure in the blood); (4) effect on rhythmic proc-
esses in the body (frequency rhythms of heart, breathing); (S) decrease in
the working volume of the lungs due to decreased muscle activity; (6) degra-
dation in ideomotor reaction, limitation of motion functions; and (7) -
decrease in the solid angle of view.
The body reacts differently to the various load~fa ctor effects. Massive
inertial forces and external concentrated reactions lead to stresses in the
connective, bone, and muscle tissues, thereby limi ting mobility. The gradi-
ent in load factor affects the rhythmic processes in the body, changing the
rhythm of the heartbeat and the frequency of breathing, and acting on the
vestibular apparatus. The load-factor impulse aff ects blood circulation,
complicating blood supply to the brain, and causes congestion, increased
pressure, and local hemorrhaging. In order to judge the viability of the
body to be sub~ected to the extreme results of all load-factor effects, it
is necessary to consider the degree of remaining viability following the
action of each eff ect. Analogously to the load-factor impulse, for which
the degree of loading is determined by Equation 2.6, the degree gf loading
CH~n~ due to acceleration itself and the degree of loading,,C~~dn/dT) due to
the gradient of acceleration are determined by _
Cd"~=max-" � C~an,dt~=max dn/d~
, x
n~P ~dR/dT)~~p
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The degree of remaining viability after the action of the load-factor
effects mentioned is determined the same way as for the load-factor effect
impulse:
C~~n~ = 1 - CH~n~ (2.7)
~O(dn/dT) = 1 - C~~dn/dT) (2.8)
In calculating the remaining viability sequentially after the action of each
factor, care must be taken in approaching the maximum permissib'.E values of
impulse, load factor, and gradient. The limiting value of load factor is
that value which, if a body is exposed to it for even a short period, pro-
duces irreversible changes in that body. According to several sets of data,
this load-factor value is in the range of 30 to 40 g's. Using this crite-
rion for evaluat~ng the ability of pilots to work, we have to reduce this
limit and consider as a threshold value of acceleration the one at which a ~
temporary disturbance of separate functions of the body occurs under condi-
tions such that after removal of the load these functions are restored.
The limiting value of the ac~:eleration gradient is even more vaguely known
and, apparently, lie.s in the range of 5 to 10 m/s3.
We must mentioned that the ma~ority of inedical researchers consider the
cumulative load-factor effect to be the key one, as provision of normal
blood flow to the body is the determining condition for supporting life.
In Reference 12, in which research resu~ts on sustaining of load factors by
pilots in good condition and out of condition are generalized, it is assured
that "one of the leading limiting circumstances of the capability of man to
tolerate acceleration is the disturbance of the oxygen balance of the body." .
The problem of tolerating load factors is still being investigated and a
full definition of a number of questions does not yet exist. This concerns, _
in first or.der, determining the relative significance of some of the factors
of the load-factor eff ect, the level of reliability, and questions of pre-
cision in estimating the limiting quantities of load factor, gradient, and
- load-factor impulse.
If we ignore questions of the pilot's ability to function, then we can take
_ the probability for a successful outcome of the loading process for the
entire crew as a single criterion for this evaluation of the cumulative
load-factor effect. To each component of the load-factor effect, namely,
its gradient, the magnitude of inertial forces, and the cumulative effect of
acceleration, there corresponda a scatter of values in the neighborhood of
the threshold magnitude. This scatter is assumed as given, with the distri-
bution of limiting load-factor values known. In such a case, the determina-
t~.on of the probability Pi (of ex~eeding for ~ust one member of the crew _
*_t~~ limiting value for a given braking .profile with a kr+own load-factor
et:��elope given as a function of time) does not representi much work, if such
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probability is determined for only a single component of the load-factor -
effect. A calculation made in this fashion wi11 become the basis of an
optimistic load-factor-effect model based on the assumption of independence
between the effects of load factor, its gradient, and laad-factor impulse.
The probability P~ for a successful outcome of the loading is determined by
using auch a model in the f~llowing fashion:
P,~=[(1 -P,l (1 -Pz)~1 -P3))"~ (2.9)
where P1, P2, and P3 are the probabilities of exceeding the maximum permis-
~ sible values of load factor, gradient, and load-factor impulse, based on
data for an out-of-condition crew; and n is the number of crew members.
Models that account for the dependence of load factors on one another and on
the prehistory of the loading process are closer to reality. The basis for
such a maximwn-permissible-value model is a computational scheme for deter-
mining the probability of a successful outcome through use of "pessimistic"
" estimates. Such estimates deal with the correlation of load-factor effects
and are made for each of the limiting effects with the condition that the
degree of loading is determined by considering the degree af remaining via-
bility for a body sub~ected earlier to other load-factor effects. Modeling
~ of the cumulative load-factor effect is done by statistical testing.
Each series of tests is conducted for a single, specific descent trajectory,
and the results are compared to the goal of choosing the best conditions of
motion. In each test, the degree of loading due to the gradient of acceler-
_ ation during the beginning phasa of braking is determi.ned for the given ran-
dom magnitude of the maximum permissible gradient obtained by using the law
for the distribution of threshold values of this gradient.
The measure of the remaining viability becomes the new level corresponding
to which there is a decrease in the magnitude of the maximum permissible
load factor as obtained from a sequential random test. Having determined
the magnitude of the maximum load factor according to the distribution furic-
tion for the threshold value of this loading and having reduced it in
accordance with the measure of remaining viability, we can calculate the
degree of loading due to load factor. For this we compare the maximum value
of load factor along the tra~ectory to the decreased value of maximum per-
missible loading. The newly computed measure of remaining viability will,
in this fashion, account for two load-factor effects. In computing each of
the tra~ectories being considered, one ahould take care to determine ahead
of time the maximum value of the relationship between load-factor integral
and the maximum permissible impulse correaponding to this load factor. Not-
ing the magnitude of load factor corresponding to this maximum, a random
selection is made based on the distribution of the maximum permissible
impulse, after which the threshold value obtained for impulse is corrected
and reduced in accordance with the measure of remaining viability computed
earlier and based on the effect of two other load-factor effects.
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The degree of loading due to .he impulse accounts for the prehistory uf
braking and, in the present case, looks like the ratio between load-factor
integral, characteristic for the initial degree of loading calculated for
the trajectory, and the newly corrected threshold value of impulse.
The probability for a successful outcome of ~he loading process for one
member of the crew as well as for the crew as a whole can be estimated with
a given reliability by repeating the indicated operation for the given
descent tra~ectc~ry and modeling the time of appearance of threshold values ~
of gradient, load factor, and impulse as a random parameter.
If an estimate of pilots' fitness levels, which also affect the successful
outcome of *_he braking process, is required, it is necessary to consider
the following parameters, which describe pilots as links in a control sys-
tem for a flight vehicle in the descent phase:
1. Inf ormation lag in the process of operating on signals from control-cir-
cuit sensing elements;
2. Pilot lag and lag in the control link connecting pilot and control
device.
Nominal values and scatter of these parameters depend on load-factor
effects, and in general the probability for a successful outcome of the
braking process includes consideration of control system reliability.
Let us consider one more aspect of the problem of returning to Earth from
interplanetary expeditions, and in particular let us attempt to determine
the possible outline of descent vehicles for the condition that the princi-
ples of descent through the atmosphere remain constant.
One of the conditions for choosing the shape of a vehicle will be assumed ~
to be the requirement for using existing methods for braking in the atmos-
phere and contY.,lling descent. We will consider vehicles of blunted and
moderately-poii.~ed shapes, which have minimal available lift-to-drag ratios
approximately equal to the required value (K,~ = 0.65 to 0.78). Descent
trajectories are chosen using the criterion of minimal cumulative load-fac-
tor eff ect. The maximunl load factor following entry into the atmosphere is
approximately 10 g's for entry speeds of 17 km/s. Figure 2.6 illustrates
one possible path in searching for the best shape of a vehicle intended for
entry into the atmo~phere with a speed of 17 km/s and carrying a crew of
six. Subsequent transition to better shapes is achieved by transforming
individual portions of the shapes under consideration while holding the fol-
lowing parameters constant: Volume required, payload weight, maximurn per- _
missible load factor, available lift-to-drag ratio, and type of descent tra- -
jectory, that i~, the set of isoaltitude and isoload-factor segments follow-
ing occurrence of peak load.
The initial shape, close in shape to existing descent vehicles and with
acceptable volume characteristics, does not provide the required
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lift-to-drag ratio and therefore the first stage of transformation of the ~
shape is a change in configuration of the aft volume, having changed the ~
angle of attack of the lifting surface. ~
'Itao bounding _:aths of transformation of the aft volume are shown in Figure
2.6, and they lead to shapes I and 2. One of these paths involves an ~
obiique intersection between body and lifting surface (shape 1); the other
involves an increase in the apex angle of the aft body and a consequt.t
increzse in lifting surface for a symmetrical shape (shape 2). Next to each
shape are noted the values for the coefficient of usable volume utilization
(Y:~), as well as the total mass of the thermal protection system and of the
descent and Ianding gear. Mass characteristics are shown in relative units
and are referred to the mass of the payload of the descent vehicles, which -
includes crew, exploration results, life-support systems, thermal control
~ systems, power su lies control s stems hardware ortion
PP ~ Y ( p r.adio systems,
and equipmer_t for carrying on operations during the period of approach to
~ the atmosphere and aftez landing.
Included in the descent and landing gear mass are those remaining descent
vehicle systems whose masses either directly or indirectly depend on the
vehicle's dimensions and shape, r_amely: Thermal protection caith insulation,
structure of the airtight cabin, base fairing, and the forward section with
the heat shield, parachute-rocket system for alighting, actuators of the
attitude control system, and systems of centering through using th~ reserves
of working fluid.
The path that leads to shapes with significant asy~etry due to the oblique
. cut leads to configurations of the shape 1 type with a small area for the
heat shield and relatively sruall total surface but with extremely inconven-
ient centering constraints, which lead to an unacceptable mass for the cen-
tering system and to a large tot21 mass of the descent vehicle, while
resulting in a. small thermal protection and vehicle structure mass.
The compromise solution shown in Figure 2.6 by shane 3 has b2tter centering
characteristics.
- The use of a symmetrical shape with low fineness ratio (shape 2) gives sat-
isfactory conditions while providing satisfactory lateral centering
(K~ = 0.75), but due to the enlarged area of the heat shield, substantial
total surface, and increase of available volume beyond what is needed, the
required m3ss of the descent and landing gear (which includes the mass of
the thermal protection system) is not acceptable. -
The characteristics of a11 three vehicle shapes mentioned are improved if
in the next stag~ of transformation t11e configuration of the forward part
is shaped as an obliquely-cut elliptic cone in a zero angle-of-attack flow.
The result of such a change manifests itself in the mass of the centering
~ control sy~;tpm due to the improved volume and centering control character-
istics (Lhapter 1). In Figure 2.6, we can see what changes are introduced
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under these circumstances in the centering criteria, in the mass of the
thermal protection system, and in the descent and landing gear as a whole.
The compromise solution (shape 3a) is attractive not only because of satis-
factory mass and centering characteristics but also due to the circumstance
that not only the initial segmented-conic but also the resulting bi-conic
shapes preserve accentable and simple outlines for the payload section,
which is shaped like a circular cone. Shapes 3 and 3a are shown in Figure
2.6. The landing section, containing the hermetically sealed cabin, which
is part of such forms, is shown in Figure 2.6 and can be used as part of
other desc~nt vehicles that have the traditional symmetric segmented-cone
shape and are intended for entry into the atmosphere from near-Earth orbits
and from lunar trajectories.
One of the advantages of vehicles witr. moderately pointed shapes which enter
- the atmosphere ae hyperbolic speeds (in addition to the volume and weight
centering advan~ages (Chapter 1)) is the lower mass of the thermal protec-
tion system and the simple arrangement of the flow past the forward cone,
whose angle of attack is close to zero. The optimum radius of bluntness
(when radiative heat flows predominate at the stagnation point) is consid-
erably smaller than 1 m. Estimates of the amount of a~,~stion of the surface
have to be made considering feedback in the mechanism of breakdown of the
thermal protection layer, that is, by considering the blocking of convective
flows and the screening of radiative heat flows by the boundary layer of the
pyrolytic gas being formed by the products of breakdown in the thermal pro-
tection material.
Figure 2.7 shows the change in ablation rate of the thermal protection mate-
rial during entry into the atmosphere along the lower boundary of the entry
corridor for three typical points on the heat shield of a vehicle with shape
3a (Figure 2.6). We can also see how the mass of the ablative coating
changes at the stagnation point as a function of time per m2 of surface.
Let us now consider questions of control during motion along a tra~ectory
that is characterized by constraints on temperature, load factor, or heat
flow.
For motion along isotemperature trajectories, there occur heat flows with
values at or less than the maximum permissible ones; here, a constant value
of heat flow indicates a constant temperature equiZibrium at the stagnation
point (Refs. 20, 40). Motion along an isote~uperature trajectory (along an
isotherm) can be obtained by changing roll attitude without changing the
regime of balance. The nominal value of the required roll angle is deter-
mined as a function of the value of the effective lift-to-drag ratio.
The equations of motion in velocity coordinates were presented in Chapter 1,
from which, by assuming that along the segment of the isotherm of interest,
the flight path angle is close to zero, and because~ sin 6 p 0 and
cos A~ 1, we obtain the following relationships for the conditions of
motion in the Earth's atmosphere:
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dt = - (2PX/PgV2) dV (2.10)
d8 pkV~ QV2 V2
-V dV 2P,~ =Kg 2PX g+ R~ (2.11)
Let us express d6/dV in terms of p and V. By suhstituting Equation 2.10 in
Equation 1.7, we obtain dH/dV =-(2PX/pgV) 6. The equation relating den-
sity to altitude is differentiated with respect to V from which we obtain ~
the expression
- dH/dV = - (dP/dV) C1/sP)� (2.12)
i By equating the right-hand sides of the expressions for dH/dV, we obtain
~ 8 = ~g~/ZPXR) ~dP/dV). (2.13)
By differentiating 6 witn respect to V, we get
d8 g d2Q dQ
dV 2P,~9 (dV2 V +dV ~ ~ (2.14 )
Substituting Equation 2.14 in Equation 2.11, we obtain the initial expres-
sion for estimating the effective lift-to-drag ratio required for motion in
the leveling-off region:
IC =~V 2 1 VZ =-�QR'V3 rd 2Q V dQ
21'z + 8R 4Pz~ 1dV2 + dV (2.15)
- From Equation 2.15, we obtain the relationship for isoaltitude (Chapter 4),
isoload factor, and isotemperature conditions of motion. For the iatter, we
can obtain the derivative of density with respect to velocity by using the
condition of constant heat flow determined by the formula
q = Cqp~n . (2.16)
By taking the derivative of q with respect to V and setting it equsl to
~ero, we find the expression for the first derivative of density with
respect to velocity:
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dP/dV = - (n/m) A ~1/V) (2.17)
Differentiating Equation 2.17 again with respect to V, we get
~2A/dV2 = (n/m) ~P/V2) C1.+ n/m) (2.18)
Substituting Equations 2.17 and 2.18 into Equation 2.15 and noting that
nX ~ pV2/2PX, we obtain the following formula for the effective lift-to-drag
ratio
K= 1 (1- V? - g"x RZ 1
' nx \ SR ~ m2 V~�) ~ (2.19)
The first two terms inside the parentheses in the expression for determining
the required magzitude of the effective lift-to-drag ratio determine the
, condition of flight along an isoaltitude (at a constant flight-path angle of
the trajectory), and the third term expresses the change in lift-to-drag
ratio (negative) required for descent into the dense layers of the atmos-
phere in accordance with the decreased flight speed. In particular, descent
along an isoload-factor trajectory corresponds to the condition m= 0.5 and
n= 2, while a descent along isotherms is given by the conditions m= 0.5;
n = 3 to 3.75.
Analysis of Equation 2.19 shows that for descent fro~u low near-Earth orbits
the magnitude of the effective lift-to-drag ratio cannot be negative for
motion along an isotherm. At the initial moment of passing the point of
leveling off, the angle of roll is close to 60 to 70� and subsequently,
according to the degree of penetration into the atiaosphere, it will, as a
rule, increase further so that later on, when it again decreases, it can
provide an increase of the effective lift-to-drag ratio for flight veloci-
ties of less than 6 km/s. During flight along an isotherm, the load factor
increases, as can be seen, for example, from Equation 1.26.
We have touched on design-ballistic questions related to descent but have
not touched on questions of the prelanding maneuver and touchdown, although
problems of automated landing at an airport are no less complex and inter-
esting thar. problems of entry into the atmosphere and the overcoming of
temperature and load-factor effects for motion at hypersonic speeds.
Although analysis of the problem of landing a vehicle with a high lift-to-
drag ratio and controllable pitch, yaw, and roll is more familiar to air-
craft designers, it is difficult to find in the near future any analogies
between problems of a gliding landing at an airport and the: problema of
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landing on planets. Therefore, we will touch on the question familiar to
designers of traditional descent vehicles but also of interest to
researchers for reusable rocket~space systems and future heavy descent vehi-
cles. We will discuss well-known landing systems that use parachutes and
rocket motors. Systems of this type will doubtless be used for a long time
for solving descent and landing problems, and it would be naive to assume
that a11 operations of returning from space will be concluded by a gliding
landing at an airport in the near tuture. Descent vehicles entering the
atmosphere from interplanetary trajectories and rocket boosters will land
_ using parachute-rocket systems, which are best from the point of view of
weight. Such parachute-rocket systems will permit full use of power plants
with onboard propellant storage and will enable landing in different parts
of the Earth.
It would be tempting to continue using traditional methods of terminal brak-
ing and, in particular, to use parachute systems of existing dimensions for
the descent modules of interplanetary vehicles or for reusable rocket assem-
blies. Let us examine the optimal characteristics of parachute-reaction
systems from this point of view. TY:e masa M~.R of the landing gear, as
expressed in relative units (that is, referred to the mass of the landing
load M~), will consist of the following components
Mc.n � Mc.n~MO � Mn.c + M R.Y + MaBT (2.20)
where M~.~ is the relative mass of the parachute device; M R.y is the rela-
tive mass of the propulsion-system fuel supply to be used for terminal brak-
ing; and FiaBT is the rElative mass of the automated equipment and apparatus
for controlling the operation of the landing gear. Optimal velocity for
parachuting corresponding to the minimum mass of the landing gear is quite
high and is measured in tens of m/s. Let us convince ourselves of this
fact. The parachute descent velocity, V~, corresponding to the minimum
mass, is determined from the condition
~C.II~dV~ = 0. (2.21)
Here .M~ BT is on the order of 0.01 for large values of M0. We will assume
that MaBT is independent of V~. From the condition that the parachute
descent has achieved equilibrium, we obtain an expression for the relative
~ mass of the parachute systems:
2t (1 + K~ + K2) ~
Mn.c= CxQpCOSBV~ ~ ~2.22
where d is the specific weight of a square m of the parachute canopy; K1 is
the rati,o between the mass of the shroud lines with the strands and the mass
of the canopies; K2 is the ratio between the mass o~ the attachment elements
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and the mear,s for deploying the parachute and the mass of the canopies; CX
is the parachute coefficient of drag; p0 is the density of the atmosphere in
the area o� landing; 6 is the angle of deviation of the parachute axis from
the velocity vector (for multicanopy systems). The values of all listed
coefficients and parameters will be assumed to be independent of the veloc-
ity of parachuting. The relative mass of the liquid propulsion rocket-land-
ing systems is determined according to the following equation:
. r ~Xl
M,~.y aT.o) `1 - e ~a 1-~-' 1',~.yno� (2. 23)
Equation 2.23 is explained in Chapter 3. The relative mass of a solid-pro-
pellant rocket can be determined by using the following expression:
v~ �
M,c.y = MKOpn l 1- e f g, ~ ( 2. 24 )
where MK~Pn is the relative mass of the rocket-motor casing. The thrust-to-
weight ratio at landing will be assumed to be much larger than unity, which,
in turn, yields VX x V~.
For liquid-propellant rockets, which, as a rule, are used for other stages
of flight, the values of aT.~, Y A.y , and np are independent of V~, with
aT,p � 1. For a solid-propellant rocket, the value of MK~gn depends on VX,
and for large motors is equal to (0.1 to 0.2)x M,A,,y . We will assume that
MKppn = constant. The magnitude of V~ does not exceed 100 m/s, and there-
fore we can assume that
Vx
1-~.~ ~x . (2.25)
. B
is accurate to within 2 to 4 percent. From the condition given in Equation
2.21 and taking into account Equations 2.22, 2.24, and 2.25, we obtain
3 4(1+K,+K2)a~~ .
V p (2.26)
CO t~ CsQo cos 6 ~ �
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It is easy to convince ourselves that for any set of values of the quanti-
ties that make up Equation 2.26, the optim~ parachuting velocity will be in
the range of 20 to 30 m/s.
The opt3mum mass for parachute-rocket systems for such descent velocities
will be 5 to 6 percent of the mass of the landing 1oad, that is, about 2
times smaller than for the majority of systesns currently used, which are
- designed for parachute velocities of 7 to 8 m/s. The above considerations
on the optimum parachuting velocity are nothing new to persons who deal with
landing systems. Implementation of systems with nonoptimal parachute veloc-
ities is explained by the fact that it is necessary to allow for the possi-
bility of an emergency landing without propulsive system as well as for di`-
ficulties of providing accuracy in a system for controlling the landing for
higher descent velocities. Errors in sensors (as a rule, altimeeers) depend
~ mainly on the descent velocity, impulse, and altitude at which the propul-
sion system is turned on.
One of the possible methods for decreasing the scatter of landing velocities
is the use of a staircase thrust profile for operating the deceleration
roclcet, thus achieving a gradual decrease of velocifiy. The task during the
first stage of decreasing velocity is the lowering of the parachute descent
velocity to a magnitude on the order of 10 to 15 m/s. The first-stage motor
= has a thrust-to-weight ratio of lees than unity and operates underneath the
parachute canopy, thus effectively reducing the landing weight. Errors in
switching on the propulsion system (in altitude) are compensated for by
increasing its duty cycle, that is, by adding a reserve of fuel for the pre-
landing auxiliary braking.
The second phase of braking is accomplished traditionally, that is, by
imparting a F:cwerful impulse to the whole system at a low altitude above the
~ Earth's surface. .
The main difficulties in implementing a staircase braking profile are con-
nected with the operation of auxiliary braking during the first stage of
decreasing velocity, �or which it is necessary to formulate a program of
rocket motor operation with ~ gradual increase in thrust, to prevent dynamic
effects on the parachute-load system.
Figure 2.8 shows the variation in thrust of solid-propellant rockets for a
load weighing 5 tons (force). Also shown is one of several possible shapes
for the first-stage motor propellant charge, which provides a progressive
burn with transition to constant thrust.
From the thrust curves for the first-stage (T1) and second-stage (T2) motors,
it is apparent that there is a parallel-3equential switching on of the
motors. tn Figure 2.9, we can see how the nominal values of descent veloc-
ity and altitude vary as a function of time for the case of programmed brak-
ing shown in Figure 2.8.
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In the exami~le given, only one of several possi,bilities is explored of pre-
touchdown braking of parachute-rocket systems, which provide nearly minimal
mass characteristics for landing gear. �
Xn /7max dNti
/ ~ ~ ifM
8
~ k,0~4S ~ ~ ~ 10
06 � D,SS ~ ' ~ K:~;
~ 075~ ~
04 4 ~ i~ dN~ OS . 10
_ ' ' 0, J
Z 'X^ I I
0,1 ,
~ _d'Ir
R�c? ~ I ~~ar '
~'f 12~ ;J 14. ;5 ;5 ve�x.a/c (i~
Figure 2.1. The Required Lift-to-Drag Ratio, I~; Maximum Load Factor
and Entry Corridor Height, ~H~; as a Function of the VPloci~ty~~
of Entry into the Earth's Atmosphere, VBX
Key:
1. VBX, in km/s -
- n ~
za i i ~
~5 t ' ~ ~
; , .
~ i i
~o
~2 i ~
5
( .
0 SO ~00 !s0 Z00 130 t, c (3 )
Figure 2.2. Maximum Sustainable Load Factor, n, As a Function of Time, t,
During Which the Load Factor Acts
Key:
. 1. Body in condition
2. Body not in condition
3. Time during which maximum load fact is sustained (s)
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) Cz)
(a )
?OOO _ _ _ �yM/c K�o,s
Vv-/S~rn/t _
fS00 - ~ y,�11Kn/c _
1000 ~
ts00 ~~P~ C 3,
7 B 9 10 11 ~
Figure 2.3. Maximum Permissible Impulse as a Function of Load Factor
Key: -
1. Impulse, J (s)
2. Maximum permissible impulse, J.~P
3. Maximum permissible impulse for body weakened by weightlessness,
J,~p
0
4. Velocity of entry (km/s)
N, N
� MM~~.6~`
M ~
60 1,1
40 Q8 ~ ~ ~ -
i
10 4 " H ~
. /
QZ 0,4 0,6 O~B V-Yo/Y
Figure 2.4. Measure M of the Cumulative Effect for Two Descent Trajectories
During Atmospheric Entry at Hyperbolic Velocity
Key :
- - - Initial tra3 ectory
- Optimal tra~ectory
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~4)
Ma,m -
~0
~ x
~6 2 -
' ~.2
Q8 , '
,r 3
0,4 , '
,~,0 6~0 ~0 ~$0 B~, tpaB C S)
Figure 2.5. Maximwn Permissible Values of the Cwnulative Effect Measure for
Different Angles of Entry into the Atmosphere, 9BX
Key :
1. Bounding set of optimal solutions
2. Reginn of possible trajectories
- 3. Boundary of permissible load-factor effect on crew ~
5. ABXt(deg.)
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0 -
~ ~
- R ~
! 2 R
v ~ -
1
R I'
� K.,,,=0,4J ~ H,.,,=0,75
G,,,,-0,7~+5 3 C�p=q785
~cc,a' l, 550 ~un' 470 .
~ R~ ~
' ~ ~.,,,~0,58 b
~ /1� Cnn'0,760 ~ Ia
/ ~ccn'!, 450 R .
H,,~= 0, 65 ~ ~ Rf /l
Cnn'07J5 Kv~' 1
~crn' ~,405 ~ J� j ' Gnn'0,765
\ If,~�0,867 ~tc"~,435 '
llotadovHe,u amceK ~~~�-0,745
~i ~ ~ccn'I,350
Figure 2.6. Shape Transformation in Searching for the Best Descent Vehicle
Configuration
~ Key: ~
1. Landing module
2. GT3~: Relative ma.ss of thermal protection system
3. G~~~: Relative mass of descent and landing gear
4. IC~ : Coefficient of usable volume util ization
II
5. G: Mass expressed as a fraction of des cent-vehicle payload
Shape K~ ~4) GT3II ~2) ~CCII(3)
II
0 - - _
1 0.41 0.745 1.550
2 0.75 0.785 1.470
3 0.58 0.760 1.450
la 0.65 0.735 1.405
2a 1 0.765 1.435
3a Q.867 0.745 1.350
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(4) (5) C6 ) _
G~ ~xrJn= ~e'
~o
a ~ : ~
- Z ,ao ~ ~ '
is ~f ' G z
,
f 0 s~ ~ i
~ / ~ G .
O S 1S 2:
,
3
i .
~ 70 40 . 60 fD0 t,c ~
Figure 2.7. Ablation Rate of Thermal Protection Layer as a Function of
Flight Time
Key :
Entry along lower boundary of the entry corridor
Entry along the upper boundary of the entry corridor
1, 2, 3. Typical points on the heat shield~ as shown in sketch
4. Ablation rate per unit of area, G(kg/m /s)
5. Cumulative ablation per uait of area, (kg/m2)
6. G without gas injection (at point 1)
7. Time (s) n
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!2)
T,m r,m~T~ �
SO S
40 4 ~
~ T ~i
JO 3
10 1 T=
!0 !
~ 1 2 d 4 5 6 Z,c
( 3)
Figure 2.8. Variation in Thrust of Lander Rocket Motors and Typical Charge
Shape of the First-Stage Motor
Key:
1. T1: First-stage thrust ~tons (force)) -
2. T~: Second-stage thrust (tons (force))
3. T: Time (s)
(1) ~2) ~
l~n Y, M/~ ,
_ 60 30
H
40 10
V
?0 f0
. 0 ? 4 6 Z,c
(3)
Figure 2.9. Parameters of Motion of a Parachute-Rocket System Using a
Staircase Prelar.ding Braking Profile
Key :
- 1. H: p.ltitude (m) ~
2. V: Velocity (m/s)
3 . T : Time (s ) `
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CHAPTER 3
CHOICE OF DESIGN-BALLISTIC PARAMETERS FOR LANDING VEHICLES OF
MARTIAN PLANETARY SYSTEMS
The basic operational feature of a Mars landing is the rarefaction of the
Martian atmosphere. Therefore, during analysis of the problem of descent to
the planet, it is necessary to determine the best method for braking, con-
sidering both the use of deployable aerodynamic braking devices and the
suitability of the vehicle for active prelanding braking by rocket motors.
It is to be expected that heavy descent vehicles, which wi11 make controlled
descents to the Martian surface with the aim of delivering large planetary
systems, will be developed in the future.
Let us review methods for the practical calculation of optimal design-bal-
listic parameters for such descent vehicles.
The choice of optimum dimensions and component layout for a Martian descent
vehicle is determined, as a rule, through a compromise bei:ween vehicle lift-
ing-surface area and propellant reserve for braking before touchdown.
During parametric calculations in the preliminary design stage, it is fre-
quently necessary to determine analytically the optimum ballistic parameters
and dimensions of the vehicle for conditions that motion in the atmosphere
will be along an optimum-energy tra~ectory. The problem of optimizing is
solved by parts, with a class of optimal control laws for descent determined
first. For the set of optimum tra3 ectories, design-ballistic parametera
are then determined, namely, load on the midsection and available lift-to-
drag ratio. Following such calculations, it is possible to go on to the
determina.tion of the brake shroud dimensions for specific configurations of
the aerodynamic braking device.
Let us consider the methods for analytical estimation of optimal design-bal-
listic parameters, solving separately the problem of choosing the optimal
control law fur descent and the problem of searching for design-baliistic
parameters and, in particular, load on the vehicle midsection. For this we
will assulile that, in the design of the vehicle, use is made of simple and
accepted solutions, that is, for aerodynamic braking at hypersonic veloci-
ties, a forward heat shield will be used. The heat shield is at an angle of
attack sufficient for obtaining a small lift-to-drag ratio, wi~h the angle
of attack and magnitude of lift-to-drag ratio, as determined by balance
weights or asymmetric shap~ of the shield, held constant during the descent
process. Motion control is achieved by rolling about the veloc3,ty vector.
It is assumed that all prelanding braking is accomplished by a rocket motor,
which is subsequently used for the vernier tnaneuver and soft landing on the
~ surface.
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moment of switching from one value o~ effecCi,ve lift,to-drag ratio to
~ another or from one load factor prof ile to another.
Four characteristic descent trajectories in altitude versus flight velocity
coordinates are shown in Figure 3.1 for a vehicle that has a hEat shield `
_ with a half-cone angle of 70� and a fineness ratio of 0.9. It is assumed
that the vehicle is trimmed at an an~le of attack corresponding to a lift-2
to-drag ratio of K= 0.4 and that it has a midsection loading of 5,900 N/m
(600 kg-force/m2) . Calculations were made for a model atmosphere approxi-
mating the nominal one (Ref. 24) and for a velocity of entry into the atmos-
phere of 4.6 lan/s, with the condition that, at the moment of initial entry
into the atmosphere, flight with maximum effective lift-to-drag ratio and
aith maximum permissible load factor of 5 g's is attained.
Tra,jectory l, as seen in the figure, is a glancing trajectory; this is the
limiting skip trajectory, because the maximum effective lift-to-drag ratio
is ma:tntained (angle of roll is constant at zer~). -
Trajectory 2 is the liTaiting steepest trajectnry, inasmuch as along this
traj ectory maxl.mum load factor is maintained during the entire time of
flight, as a result of which the vehicle does not have enough time for brak-
ing and at the surface has a flight velocity corresponding to this load fac-
tor (1.7 to 1.9 lan/s).
In the first case, the flight velocity at the surface is also large, inas-
mucr as the second dive into the atmosphere occurs a~ong a steep traj ectory.
Flight along a sloping trajectory following the initial dive is preferable,
because the vehicle has considerable time for braking in the dense layers of
the atmosphere and approaches the surface of the planet along a shallow
slope. ~ao such traj ectories are shown in Figure 3.1. In both cases, the
vehicle approaches the surface with maximum effective lift-to-drag ratio, in
- the same manner as is done along the entire length of traj ectory 1. How- ~
ever, after passing maximum load factor immediately after atmospheric entry
and in contrast to trajectory 1, the effective lift-tc.-drag ratio, that is,
the proj ection of the lift force on the vertical, is decreased . The shallow _
skip of tra~ ectory 3 is designed in such a mar_ner that during the period of
flight, starting with a velocity of 3.2 km/s and ending with a velocity of
1.2 km/s, the vehicle moves with a roll angle of 10� corresponding to tiie
value of effective lift-to-drag ratio of 0.135. Trajectory 4 has a. flight
segment of isoaltitude leveling-off during the same perir~d. The instantane-
ous value of roll angle during this segment is determined from the condition
that the derivative of the flight-path angle be zero.
For both trajectories 3 and 4, a near-minimum value of flight velocity is
attained at altitudes of 2 to 4 km, with the value of this velocity somewhat
lower for the shallow-skip tra~ectory. However, the trajectory with the
isoaltitude plateau is preferable, because, due to the shallower approach
to altitudes of 3 to 4 km, the losses in propellant for the prelanding brak-
ing will be somewhat lower than for the skipping trajectory. .
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~
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Let us now investigate how the speed of approach to the surface v^~~,es for
, such level trajectories when lifting and braking characteristi.:., ~f the
vehicl? are varied. The terminal velocity of flight at different altitudes
during descent with an isoaltttude segment ls shown in Fig.:re 3.2 as a func-
tion of the loading on the midsection of the vehicle for different values of
the available lift-to-drag ratio for a nominal model of the atmosphere. We
note a sharp increase in velocity at ~ow load levels, increasing f rom 200 to
500 N/m2 (20 to 50 percent kg-force/m ) up to 2,000 to 3,000 N/m2 (200 to
300 kg-force/m2) and a further gradual growth for PX greater than 4,000 N/m2
to 5,000 N/m2 (4~0 to 500 kg-force/m2).
As we can sep from the figure (lower curves), rhe sensitivity of the trajec-
tory in terrus of the terminal velocity criterion as a function of available
lift-to-drag ratio is nonlinear. This nonlinearity is mor~ ;raphically rep-
resented in Figure 3.3, where it is shown how, at a 2-km altitude, flight
velocity of a trajectory with leveiing-off varies with increasing available
lift-to-drag ratio for a constant PX (PX = 7,850 N/m2 or 800 kg-force/m2),
and also as a function of increase in loading on the midsection for a con-
stant lift-to-drag ratio (K = 0.526). We see once again that in the area of
large PX the terminal velocity increases with PX but quite sluggishly. The
gradient of the variation of velocity as a function of available lift-to-
drag ratio is clearly not constant, and the velocity decreases most when the
Iift-to-drag ratio increases zr~m 0.2 to 0.6.
Generally speaking, we cannot consider the relationship among terminal
velocity, Px, and K in isolation, inasmuch as these characteristics are
closely related to one another by the lift-drag polar of the vehicle. For
one and the same vehicle having the same aerodynamic shape and dimensions,
an increase in lift capability is accompanied by worsening braking charac-
teristics, and it would be more correct to consider the effect of the angle
of attack on the magnitude of the terminal velocity. It turns out that for
each vehicle shape there exists a zone of best trim angles of attack, and
Figure 3.4 shows this zone for a vehicle with a shield in the shape of a
symmetric blunted cone with a semicone angle of 70�. Two curves are shown,
relating the speed of flight at an altitude of 2 km to the angle of attack
for an isoaltitude trajectory, with one curve characterizing a vehicle with
an extensive heat shield and the other describing a vehicle of relatively
small diameter and fineness ratio close to unity.
It is useful to approximate the value of terminal flight velocity considered
above through formulas, which can be used to determine that velocity for the
optimal descent trajectory as a function of flight altitude and loading on
the midsection of the vehicle for the range of the best regions of trim,
that is, for those values of the available lift-to-drag ratio (KP) that are
characteristic for blunted and for moderately pointed shapes.
For the majority of atmosphere models, the structure of the formulas can be
developed on the basis of steady-state motion, as a result of which these
foL-mulas will have the form -
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We will also assume that the flow field about the vehicle and the method of
generating aerodynamic forces used for braking and control of the descent
trajectory are traditional, that is, vehicles with simple aerodynamic shapes
having small available lift-to--drag ratios are used. Such a vehicle has a
deployed aerodynamic braking shr~ud carrying the payload enclosed by a
lightweight fairing in the shac~ow zone. The shroud can project beyond the
payload, and the shield itself has a blunted or moderately pointed shape.
?et us consider the type of descent trajectory for which the best design-
ballistic vehicle parameters are to be determined.
The optimal descent control law for a Martian descent vehicle is chosen in
accordance with criteria that detine the terminal descent phase at the
_ instant the vehicle touches the planet's surface. In general, such a crite-
rion can be the mass of the propulsion system for given dimensions of the
_ aerodynamic braking shroud. By making a number of simplifications permissi- -
ble for parametric estimates, the weight or mass criterion can be isolated, -
as it is not necessary to consider the thrst-to-weight ratio and the mass of
the combustion chamber, but to work only with the fuel supply for the pre-
touchdown braking and landing maneuvers. The choice of optimum thrust-to-
weight ratio has to be made considering changes in flight velocity at the
various altitudes at which the rocket is ignited and the thrust-control pro-
gram during braking. Several results of calculations of the optimum thrust-
to-weight ratio will be discussed at the end of the chapter. The speed of
entry into the atmosphere of Mars is significantly lower than the speed of
entry into the atmospheres of EarCh or Venus; it varjes from 3.6 km/s
(descent from low near-planetary orbits) up to 4.~ to 6 km/s (descent from
highly elliptical areocentric orbits or approach txaje~�toriPS). Along such
descent trajectories, the maximum cunvective heat flows ~w~.~ch predominate
in the heat transfer) do not, as a rule, exceed 102 kW/m and thus the con-
straints of heating and load-factor effects are not the decisive ones for
Martian planetary descent vehicles.
If we ignore the thrust-to-weight ratio and take the characteristic velocity
of the descent rocket motor as the criterion for searching for the regime of
descent control, we can make additional simplifications in the problem, -
reducing the problem of optimizing the weight-energy characteristics to a
problem of minimizing the characteristics of the descent trajectory.
Results of solving such problems appear mosr graph~c if the minimum of just ~
a single, basic characteristic of the trajectory is determined, namely, the
velocity at the desired final altitude of passive flight at the moment that
the landing rocket motor is ignited.
Strictly speaking, we are not in a position (having optimized the conditions
of descent motion by using the criterion of a minimum characteristic veloc-
ity) to ignore the peculiarities of the regime of active braking and, in
particular, the flight-path angle, which is the angle of inclination of the
velocity vector to the horizontal at the end of the aerodynamic braking seg-
ment. The characteristic landing velocity VX has the following components:
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- l. VH is the velocity of flight at altitude H~ at the end of aerodynamic -
0
braking and before ignition of the rocket motor for prelanding braking (the
magnitude of V~0 depends on the type of descent tra~ectory chosen and on the
ballistic characteristics af the vehicle.
2. aVr~ is the gravitational loss in velocity during the active brakin~
segment (determined by the thrust vectory control profile and the length of
the br~king segment, that is, by the thrust-to-~aeight ratio of the rocket
motor, which, in its turn, depends on the altitude H~ and the initial flight
path angle A~). The magnitude of evrn is small and amounts to from 2 to 4
percent of VHO; thus, in parametric studies of the optimal design-ballistic
parameters, it is possible to neglect the effect of characteristics of the
final segment of the descent trajectory on the value of dVrn,
3. ~VY~P is the characteristic velocity loss due to cont~ol during the seg-
ment of active prelanding braking. The total loss due to control amounts to
2 to 5 percent of VH .
0
. The three components mentioned can be combined, inasmuch as they determine
the propellant reserves needed for the pre].anding braking. Thus, ~VX for _
the braking is ~VXTOPM - V~O + ~Vr~ + ~VY~, and ~VXTOPM - KV Vg~~ Where
K~ = 1.04 to 1.10. The remaining components of the characteristic velocity
determine propellant losses during the landing segment (~V ) and the
xIIOC
amoui~t of unexpended propeilant in the tanks (~Vr~). For all practical
purposes, ~VxIIOC is independent of the descent trajectory characteristics:
~VxIIOC = OV~ + ~V ~CT ~3.1]
where aV~ is the velocity reserve for the prelanding maneuver. aV~
depends on the time alloted for the maneuver. For a maneuver time of 60 s,
the magnitude of GV~ is 270 m/s; OV ~ ~T are the velocity losses during
vernier braking for a soft landing. The magnitude of ~V ~~T, as a rule, is
small and amounts to between 20 to 60 m/s. ~Vr~,p is the component of the
characteristic velocity, determined by the guaranteed reserve of propellant
in the tanks. The magnitude of ~Vr~ is usually determined by the tank
coefficient and, in practice, does not depend on the parameters of the
descent trajectory.
Let us observe how the conditions of motion on the aerodynamic-braking seg-
ment influence the terminal phase of such braking. Parameters of the tra-
jectory to be varied during optimizat~~n are the trim angle of attack for
the vehicle of� a particular configura_~on and a set of control actions that
determine the variation of the roll angle and, thus, the magnitude of the
~ effective lift-to-drag ratio. These control actions can be in the form of a
series of turns, by defining the instants at which the attitude control
thrusters are turned on or, which is even more graphic, by defining the
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VHo-Ky Y g31H f~KP~~ ~3~2~
where K5, is the coefficient that determines the control method on the descent
segment and the peculiariti,es of the atmosphere model being used; gn is the
rree-fall acceleration close to the surface of the planet; g3 is the free-
fall acceleration in Earth's at~nosphere; pH is the density of the atmaaphere
at the terminal altitude of flight; the function f(~), which defines veloc-
ity, depends on the magnitude of the lift~to-drag ra io in the region of
optimal values of K.P.
For an isoaltitude control law, for segmented or conic blunted shapes, and
for a nominal atmoephere model, Equation 3.2 has the form ~
V,~o = 0,49qt,�~SP�~5 (1, 26Kp - 0, ~ 7KP 0,45)�~5, ( 3. 3) -
.
which desc~ibes the magnitude of flight velocity at altitudes from 0 to 4
km; for 980 N/m2 (100 kg-force/m2) 5 PX S 6,800 N/m2 (700 ltg-force/m2); and
for 0.2 S ICp 0.6, with an accuracy on the order of 10 percent.
The proposed method for determining the optima.l parameters of descent and
landing ~;ears is based on parametric calculations, in which widely known
terms are used: MidsecCion loading, specific mass/m2 of structure, specific _
mass for the rocket--motor casings, relative weight of the propulsion system,
thrust-to-weight ratio, and others. The authors decided not to use indica-
tors based on mass but to preserve commonly accepted weight criteria,
because this is permitted by the GOST [All-union state standard].
The weight of the descent and landing gear, G~~n, consists of the following
components:
~CCII ~ G~ + GAY + G060PY~1,,
where G~ is the weight of the structure and thermal protection layer of the
aerodynamic braking device; GAy is the weight of the propulsion system
including the propellant supply; and G oG o P~Q is the equipment weight.
The third itemized comp~nent does not, in practice, depend on the design-
ballistic parameters of the vehicle, and for thfs reason we will determine
the optimal loading on the midsection starting with the minimu~ weight of
the structure, thermal protection, and propulsion system.
For heavy descent vehicles intended for missions to Mars, the weight char-
acteristics of the vehicle structure conflict with analogous characteristics _
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of the landing propulsion system. Because of this, the parameter usea, and
- the one that relates these two, is the flight speed at tht moment the rocket -
is ignited before landing. The task of optimization is thus the determina-
tion of the optimum loading on the midsection. We will determine the opti-
mum weight of the descent and landing gear in.relative units, without tying
- the optimum loading on the midsection to concrete dimensions of the vehicle
or relating component weights of the descent and landing gear to the initial
weight G0, which is characteristic for the vehicle before atmospheric entry.
The weight of str.ucture and tYiermal pr~tection i,s usually estimated during
parametric design calculations by using specific weight o� the lifting or
wetted surfaces. We, too, will use the specific weight, defining it as the
weight of structure and thermal protection per each m2 of lifting surface:
G~1A � GKT/S (3.4)
_ By letting PX = G~/CXS, we obtain as the weight of structure and thermal
protection:
G~
GK.T -~yA ~xPx � (3. 5)
It should not be assumed, even for parametric design estimates, that G yA
will remain constant for vehicles of different configuration and size.
Increase in vehicle dimensions and lifting surface area for the same total
weight will necessarily lead to a decrease in the weight of each m2 of lift-
ing surface due to the circumstances listed below. First, for decreased
loading on the midsection there will be a decrease in the thermal stress for
the descent trajectory, the vehicle will be braked at higher altitudes, and
the weight of the thermal protection coating per m2 of heat-shield surface
and bottom fairing will decrease. This decrease is also aided by the addi-
tional reduction in convective heat flows due to the increased bluntness
radius with an increase in vehicle dimension.
Second, with an increase in dimensions of the lifting surface, there is a
decrease in the fraction of structural weight per unit area of the surface.
This structure consists of load-carrying elements of the vehicle fi~selage, -
and payload support components a:id assemblies of airlocks, docking a~echa-
nisms, and so forth. Some increase in weight of the lifting envelope of the
heat shield affects stability due to increased blun~ness radius and counter-
acts the general trend to reducing the weight per m of lifting surface with
an increase in its area.
In the limit, as the dimensions of the shield increase (Px 0), the struc-
tural weight per m2of such a shield will tend toward very small values.
For vehicles with the customary relationship between diameter and length and
with modest dimensions and values of loading on the midsection
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(PX = 4,000 to 8,000 N/mZ or 400 to 800 kg-force/m2), the specific weight of
the structure and thermal protection of the body for each m~ of lifting sur-
face changes little with changes in the diameter of the yehicle.
From the material presented, it follows that first, the specific weight of
structure and thermal protection per m2 of l~,fting surface is not a constant
valuQ for vehicles with lifting surfaces of different dimensions, and, sec-
ond, this specific weight is a nonlinear function of the midsection loading,
increasing with increases in this loading, and trrith the steepest increase in
the specific weight to be expected when midsection loads are small.
In the general case, the relationship between specific weight G~1~ and the
midsection load of bluntly-shaped vehicles can be described using the equa-
tion:
C
G yq - ~lPX 2, (3. 6)
where C2 is an index of degree and lies in the range of 0 to 1.
Typical relationships between specific weight and midsection load are shown
in Figure 3.5 and were obtained by approximating the results of design-lay-
out work on vehicles that have different ratios between heat-shield area and
total wetted surface. This relationship has the form
~ YQ _ KKTPXO . 5 ~ ( 3 . 7 )
where Y.KT is the coefficient that defines the rate of change in G JR for a
structure with thermal protection. Substituting Equation 3.6 in Equation
3.4 and transforming to relative units, we obtain
GKT ~'KT 10.5 . (3.8)
~ ~xPx
Undoubtedly, the relationships presented reflect more the qualitative than
the quantitative side of the relationships between the structural weight and
the heat-shield dimensions of a descent vehicle. In each concrete case, the
weight characteristics will depend on the power plant of the vehicle, the
parfiiculars of loading, the computational cases, the layout of structural
components and assemblies of the heat shield, and the materials used for
thermal protection and structure. However, for parametric studies, it is
usually necessary to neglect specifics of the structure so that, having once
determined the range of optimal values of the design parameters, it is pos-
sible to estimate their values more precisely through the use of add~tional,
more detailed computational models.
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Let us consider the weight model of the propulsion system for prelanding
braking and touchdown. The weight of the power plant has the following com-
ponents:
~n.y - ~Ka N~ Grni~n ~ G6aK+ ~ 3. 9~
where G~ is the weight of the combustion chamber; GTO~/~ is the weight ~f -
the propellant required for landing; and G 6~ is the weight of the tanks .
with fittings.
By switching to relative units in Equation 3.9 and referring the weight of
the components to the total weight of the vehicle, we obtain
~A,y=1'~3,~no~' GG~~~, ~l-{-ar.o)~ (3.10)
0
where y~ is the specific weight of the combustion chamber (kg-force of
weight per kg-force of thrust); n~ is the initial thrust-to-weight ratio of
the landing propulsion system; aT~ is the tank coefficient, which is the
- ratio of the propellant tank system weight to the weight of the propellant.
Since ~
vX
GO - G ronn _ C ~ 3.1.1 ~
Go
we have
. r Vxl ,
GA,y = ( ~ [L.r.~) \ 1 - e~e 1 + Y~�no. (3 .12 )
At the beginning of this chapter, we considered the components of character-
istic velocity and divided VX as follows:
V X= V ya T- OV r.n T~v~�up ~~~MBK T~~IOC7 ~~V roP -
_ ~V x.rupnc ~ ~V Y.noc ~ ~V rap�
The magnitude of DVr~ is determined by the tank coefficient and is included
in aT0'
By considering Equation 3.1, we get the relationship between VX and VH
0
VX = KVV~O + ~VxROC , (3.13)
where ~VxIIOC is independent of VH .
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The magnitude of VX does not exceed 1,500 to 2,000 m/s, and therefore it is
possible to transform Equation 3.12 into the following expression with only
a small error:
VZ `
~~.y=~~+ar.o) ( - 2~z~` I-}-Y~ano� (3.14)
/
By considering the expression of Equation 3.3 for V~ , we obtain
0 -
KvVyaPXs-{-~AVs.na ~KvVyoPs'S-{-~VX.noc~2
- ~,~.y=(I-~-a:.o)C. 2B2J2 ~-}-Y,~y~o. (3.15)
where
VH = 0.96pg 0.5 ~1.26K~2 - 0.77RP + 0.45)-~'S (3.16)
0
If the thrust-to-weight ratio of the landing power plant is assumed constant
during parametric studies of different descent vehicle variants, then the
altitude at which the power plant is switched on for the prelanding braking
will also remain constant for vehicles with arbitrary loads on the midsec-
tion. By choosing a value of Kp close to the optimum one for the class of
shapes considered, we can determine the optimum value for PX from the con-
dition
aP ~QK.T+QA.Y)=0. (3.17)
x
Substituting Equations 3.8 and 3.15 into Equation 3.17 and assuming that
OV~pC is a small quantity, we obtain the following equation for the opti-
mum loading on the midsection:
/ J~~ K2 V 2
l 1 T at.o~ KvV
yo y_ KK.r y3 1-f (lr.o~ v N~ _ ~ 3. I8 ~
_ 2g! 2Cx 2g2~2
where
y = 1/ X
For realistic values of Cx, J, H~ and aT, the optimum value of PX is found
to be in the range from 1,500 N/m~ to 3,550 N/m2 (150 kg-force/m2 to 350 kg-
force/m2).
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Figure 3.6 ~~hows the typical variation in the weight constituents of the
descent and landing gear as a function of vehicle lifting surface for one of
the possible configur.ations based on the initial data given.
As already mentioned in Chapter 1, the reiative weight of the descent and
landing gear cannot be used as an objective criterion for comparison, and
Figuze 3.7 substantiates this. The figure illustrates the dependence of the
_ ratio of the weight of the descent and landing gear to the weight of the
payload of the descent vehicle as a function of the load on the midsection,
that is, as a f�,zction of ~he dimensions of the lifting surface for a con-
stant total vehicle weight. The opt:imum lies in the range of
PX = 1,500 N/m2 to 3,000 N/m2 (150 kg-force/m2 to 300 kg-force/m2), which is
indicated broadly in the curves of Figure 3.6 but more distinctly i:~ Figure
3.7.
In conclusion, let us consider the question of the best thrust-to-weight
ratio of the landing rocket motor. In choosing the optimum thrust for the
landing phase of a Martian landing vehicle, allowances should be made for
g�ravitational losses during the prelanding braking and for the increase in
characteristic velocity to account for c.hanges in initial conditions at the
moment the rocket motor is ignited. Inasmuch as the vehicle is actively
braked by aerodynamic drag during its approach to the surface, it is more
advantageous to ignite the landing rocket motor at the lowest possible alti-
tude, tnus resulting in a thrust-to-weight ratio of about 2. A typical
relationship of the relative mass of the propellant section (GT~), the pro-
� pul.sion system G R~1 , and the flight velocity at the instant the rocket
motor is ignited is shown in Figure 3.8 as a function of the initial thrust-
to-weight ratio of the landing vehicle.
H, Kn
~6 T 1
- JZ 3
_ ~ 4
1
4 -
G
/D00 ?000 ,~000 V, M/c
Figure 3.1. T~pical Descent Trajectories Through the Martian Atmosphere.
Altitude H in lan, velocity V in m/s
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~OC 3 4 �
Kp-Q4; H-fO~M
((p=Q4; H~7,SKn
d00 '
lfo=0; N -0
400 ~ !
; ~ ,r,-q,2; h o i
' ' 100 ~0�45; y0 Kc-Q4; H�0
j ~
iooo zoaa ~
ao 4 0o P,~, ~;/M t!~~
1
0
0 ?
0
0 J
G
O 4 0
0 P,~, x z c, . ~ Z
, R
Figure 3.2. Terminal Flight Velocity as a Function of Vehicle Midsection
Loading
Key:
1. Velocity, V (m/s) ,
2. Midsection loading, Px, in N/m2 (upper scale) and in kg-force/m2
(lower scale) .
3. Available lift-to-drag ratio,
4. Altitude, H, in km, for which terminal velocity is shown
~i) .
V,n/c
900
BSO P.~(K "D.516)
B00
Ifp(~,~ -8000) .
750
0 0~4 06 8 1,0 Ifo
10A7 8000 9000 1p000 P,~, H/nt ~ Z~
Figure 3.3. Terminal Flight Velocity at 2-km Altitude as a Function of Mid- -
section Loading (for Constant Lift-to-Drag Ratio) and as a
Function of Lift-to-Drag Ratio for Constant Midsection Loading
Key:
~ 1. V (m/s)
2. PX (N/m2)
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~1) y.H/c
800
600
400 -
100
0 10 20 30 40 SO ad ~
(2)
Figure 3.4. Terminal Flight Velocity as a Function of Trim Angle of Attack
Key :
1. V (m/s)
2. Trim a (deg. )
.
~ 1 ~ ~ya, x~c/n t �
~ 3~ ~Mr=Z,1
Il,~,=7~0
60
_ ~ -;8
40 ~
i
ZO
~ 10D0 4000 6000 8~00 JOOOO P, H/MZ
100 400 600 B00 f000 P,,~rtc/n~ ~2
Figure 3.5. Specif ic Weight of m2 of Lifting Surface as a Function of Mid-
section Load
Key :
1. Specific weight, GyA (kg-force/m2)
2. Midsection load, PX, in N/m2 (upper scale) and in kg-force/m2
(lower scale)
3. Coefficient K~, defining rate of change in specific weight ~~R
as a function of load PX = 1.8, 2.0, 2.2)
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(1) ~�rt .
s
,M'=Q~ ~ i
n'~`~ 1 ~ ~ ~
y
,,4 X J H,= S ~p=0~4
v4 04
Q3
t ~d9 Gya-70KZC/nt
QZ l � ~ -SOMZ~~~t c 3 ~
, c,~�~o~r~iM.
Cya=30Krc/n:
~ G~ (G ~�X,o~,~)
Q C5 ~ fNm (Gyd -cons t) '
i
f000 1000 3000 4000 SOOOP,,N/n
0 f00 f00 J00 400 S00 P,~,Krt/nt ~2 ~
~
Figure 3.6. Mass Changes in the Components of the Descent and Landing Gear
as a Function of Vehicle Lifting Surface
Key:
1. G~Cn, relative mass of descent and landing gear
2. Midsection loading PX in N/m2 (upper scale) and in kg-force/m2
(lower scale)
3. Specific weight G y~y (kg-force/m2)
4. G~y , relative mass of power plant
5. GgT, relative mass of thermal protection and structure
(Equations 3.4 and 3.7)
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~ ~ ~ KI �Gttn~GnM
~A~~ /
~ i /fp�Qi
. \ /
0,8 " ~
. ~v'0.~ /
~
06 X0l~~4 \ / ~�-QS ,
_ ,
f000 Z000 J000 4000 P�N/ro2
0 f00 100 J00 400 P,,KZC/nt
� ~2~
_ Figure 3.7. Mass of the Descent and Landing Gear as a Function of Payload
for Vehicles of Different Lifting Area
Key:
1. KB = GC~~/G~H, ratio of descent and landing gear mass to the pay-
load mass
2. Midsection loading PX in N/m2 (upper scale) and in kg-force/m2 -
(lower scale)
(1)
C ~M, rr/c ' ~ ;
t~ /00J
~e~
09 900
08 dOD ~ ~`,o
O1 ~OO V� . .
~ 1 2 n,
Figvre 3.8. Relative Mass of Rocket Motor and Flight Velocity at the
Instant of Motor Tgnition as a Function of Initial Thrust-to-
Weight Ratio
Key :
1. VH (m/s) -
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CHAPTER 4
ANALYSIS OF TRAJECTORY CHAR.ACTERISTICS FOR CONTROLLID ENTRY -
AND DESCENT INTO THE ATMOSPHERE OF VENUS
The great density of the Venusian atmosphere determines the shape of descent ~
vel-iicles and the schedule of operations to be conducted during the descent
~-nd after the landing on the surface. Contradictory requir ements, which the
vehicle has to satisfy at the different stages of flight, make it necessary
to change the vehicle configuration as it penetrates the atmosphere. Entry
into the atmosphere is made with great speed, reaching 12 lmi/s for direct
descent from approach trajectories, 10.2 km/s for descent from highly ellip-
tical orbits, and 7.5 km/s for descent from a near-planetary orbit. The
vehicle is braked intensively in the upper layers of the atmosphere and, for
= a shallow entry, the conditions of motion and heat exchange resemble proc-
esses in the Earth`s atmosphere. The shape of a descent vehicle designed
_ for braking during entry into the atmosphere does not differ from the tradi-
tioaal shape of descent vehicles. However, the process of intensive braking
is accomplished, as a rule, at altitudes of 30 to 40 lan, af t er which the
vehicle begins to descend gradually and to penetrate the dense, i00-degree-C
atmosphere.
Vehicles for descent into or drift in an atmo~;phere whose p r essure reaches
100 kg-force/m2 resemble more a deep-diving ba.thyscaph than a flight vehicle
in their structure and external appearance.
There are two groups of problems that had to be solved in d e signing ballis-
tic descent vehicles and that will have to be met in the future in the ~
design of heavy controllable vehicles for the exploration of Venus. _
- The first group of problems is connected with selecting the best means of
braking in the upper layers of the atmosphere and with decreasing th~ levels -
of load factors and thermal stresses fnr the descent trajec t ory.
- The secon.d group of problems is defined by the operating conditions at high
external nressures and temperatures, and the problems of des cent and landing
under these conditions.
Une way of solving problems of landing on Venus is a functional separation
of the planetary landing system into descent and landing vehicles, with the -
latter being the payload of the former. Conditions on the p lanet are
apparently not suitable for conduct-!ng landin~ operations o r for the design
in the near future of planetary systems equipped with means of returning to
Earth or to a near-planetary orbit. However, the design of neavy control-
lable desr_ent vehicles capable of effectively braking in the upper layers
of the planet's atmosphere, deli~~ering ther e their landing or drift vehicles
for flight in the denser layers of the atmosphere, appears to be one of the
- tasks of space flight for the near future. Let us consider the design-bal-
listic questions applicable to such vehicles, dealing with d escent and con-
- trc,~llPd flight at ~he boundary of the dense atmospheric layers.
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~ In contrast to a descent into the atmosphere of Mars (Chapter 3), the prob-
lem of absori~ing energy of the vehicle and obtaining acceptable landing
velocities for Venus is achieved without special changes in the structure
and without having to perform special control operations during the descent
segment. A ballistic descent vehicle entering the atmosphere within a wide
range of entry angles (from -30 to -90�) will decelerate to a speed required
for Che operation of landing systems. The speed of descent to the surface
of the planet is determined by the specific loading on the vehicle's midsec-
tion and depends on the conditions of entry into the atmosphere and, for
practical purposes, is independenC of the lift-to-drag ratio.
The region of attainable descent velocities that are characteristic for dif-
ferent braki:~g characteristics of the vehicle is shown in Figure 4.1 for
variations in atmospheric parameters. It is apparent that acceptable land-
- ing veloci~_ies can be achieved without special auxiliary braking systems.
- The instant of terminating the intensive braking phase during atmospheric
entr~ aad Lransition t~ slow descent should be fixed, becaus~ it is pre-
cisel~ at this instant that a change is made in the method of controlling the
trajectory, operation of descent vehicle systems is terminated, and landing
gears or paraballoon systems intended for aerostatic braking of the landing
vehicle begin functionin~. Apparently, it is the transition to subsonic
flight speeds which indicates that deployment of landing systems can be ini-
tiated. We can determine when such a moment arrives for specific conditions
of atmospheric entry (entry speed of 11 km/s, trajectory flight-path angle
of 30� at a 130-km altitude) by examining Figure 4.2. The achievement of a
flight speed of 200 m/s is taken as the transition point. The curve showing
the terminal flight altitude (Hk) at this moment (for a nominal model of the
atmosphere ror a descent trajectory wi~h maximum effective lift-to-drag
ratio, and for vehicles with different lifting characteristics (KP) and spe-
cific loading (PX) on the lifting surface) indicates that for acceptable
- vehicle dimensions, the increase in the available lift-to-drag ratio makes
it possible to start the descent phase at higher altitudes (and simultane-
ously increasing the vehicle's loiter time in the upper layers of the atmos-
phere and the effectiver~ess of operations for sounding the at~ospher.e of the
planet). It is characteristic that even for ballistic descent the initial
descent altitude is between 20 and 40 km. However, a ballistic descent in
conjunction with the steep angle of entry into the atmosphere is accompanied
by high load factors and heat flows. Using a small or moderate lift-to-drag
ratio makes shallow entries into the atmosphere possible (entry angles less
than 30�), thus reducing both loading and thermal stress. The graph of Fig-
ure 4.3 illustrates the results of calculations of limiting entry trajecto-
ries attainable for motion near the upper boundary of the entry corridor
(condition of capture by the atmosphere) and close to the lower boundary of
the entry ~orridor (determined by the maximum load factor). The required
dimension of the entry corridor is determined on the basis of the differ-
ence in the altitude of the theoretical perigees, and for a vehicle with
braking characteristics close to conventional ones (Px = 2,000 to 5,000 N/m2
(200 to 500 kg-�orce/m2)), this results in a 40-1~-higri entry corridor. The
figure shows the relationship between the maximum load factors reached during
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entry along the iower boundary of the corridor (entry angles of 5 to 7� at
a 130-1an altitude) and the availab].e lift-to-drag ratio. 'This relationship
is valid when, for motion along the upper boundary of the corridor of pre-
scribed dimension, capture is pr^vided for by the value of the lift-to-drag
ratio considered (for negative lift). The region between the two curves in
the fi~ure is characterlstic for most approach trajectories, with a speed
of entry into the atmosphere between 11 km/s (lower curve) anri 12 km/s
(upper curve). Analogously to returns to Earth from tnterplanetary approach
trajectnries, an increase in l.ift--to-drag ratio and use of a shallow atmos-
pheric entry result in a sharp reduction in specific loading. For example,
].oading is decreased by 2-1/2 to 3 times by changing from a ballistic
vehicle to a vehicle with a moderate lift-to-drag ratio of about 0.6 to 0.8.
An even greater reduction in l~ad factor is expected for descents from near.-
planetary orbits. The high additional energy losses of such a mode of
sounding the atmosphere are justified by the added opportunities For making
a circuinspect choice of the region to be studied in the atmosphere and then
directing the descent vehicle to that location. Figure 4.4 shows how the
magnitude of the maximum load factor for vehicles with such braking
characteristics changes with available lift-to-drag ratio for the case of
entry from near-planetary orbits.
The cross hatch regions show what is attainable for descent vehicles enter-
ing the atmosphere from highly elliptical orbits with the lower boundaries
- of each zone corresponding to entry speeds into the atmosphere of 9 km/s and
upper ones to entry speeds of 10 km/s. The various regions shown in Figure
4.4 are characterized by varying .requ~_rements on the control during descent
- and landing operations of the planetary sounder. Region l, whose bounda- -
ries are shown by a dot-dash line, corresponds to low control requirPment~
_ for descent from orbit, but entails high load factors for vehicles of tra-
ditional form, with lower load factors for vehicles with high lift-to-drag
ratios. The dimensions of the required entry corr.idor in this case is 80 ~
km. Region 2, bounded by dashed lines, corresponds to a 40-km entry corri-
dor, and, finally, region 3, whose boundary is shown by solid lines, corre-
sponds to a 20-km entry corridor. In the latter case, it is possible to
reduce load factors to acceptable 1evels for traditionally shaped vehicles.
Proper choice of the conditions of motion during the descent phase also
makes it possible for vehicles with moderate available lift-to-drag ratios
to brake at higher altitudes than for constant-lift flight (Figure 4.2).
Use of levelin~-off regions after passing the maximum load factor makes it ~
possible to avoid, on one side, skipping and leaving t-he atmosphere, and on
the ott~er side, to avoid dives with a loss of altitude. One of the means
for leve]i.ng off is control based on an isoaltitude law for changing the
effective lift-to-drag ratio. For the condition that the derivative of the
flight-path angle equals zero, there exists a completely defined law in the
equations of motion for the variat3.on of the available lift-to-drag ratio
and drag of the vehicle. It is possible to achieve flight with a constant
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flight-path angle and constant loading on the midsection without changing
the angle o~ attack by controlling the roll attitude about the velocity vec- -
tor for small initial flight-path angles. This has to be initiated iimnedi-
ately after the instant of passing maximum load factor and before the begin-
ning of the skip. If the flight-path angle is assumed to be small, then the
programmed relationship for the roll angle of the vehicle has the form
YKP=arccos ( ~ X ~1 ~ 1 g-1
P n r gnRn I g3
where
/L - n~ma: IV~.
xT -
. Yl + Kp.
Here n~ is the programmed instantaneous value of load factor determined by
flight speed;gR, RR are the acceleration of gravity and the radius of the
planet; g3 is the acceleration of gravity on Earth; and n is the magni-
Emax
tude of the maximum load factor before entering the leveling-off region.
Results of the search for the best region of braking show that for the main
segment of the descent traj ectory it is best to use leveled-off motion or
shallow descent. Figure 4.5 shows an example of an optimal trajectory for
entry along the lower boundary of a 40-1~ corridor. Speed of atmospheric
entry is 11 1~/s, the loading on the midsection is 5,000 N/m2, and the
available lif t-to-drag ratio is 0.5.
The isoaltitude segment of leveling off is provided by cha.nging the effec-
tive lift-to-drag ratio, whose required magnitude is shown by the dashed
line in the same figure.
The isoaltitude segment of the trajectory and the isoload-factor segment
that follows it are more graphically represented in Figure 4.6, where the
same trajectory is shown in the altitude-vs-flight-speed coordinate system.
The dashed line shows the required nominal value of the roll angle deter-
mined from the condition of implementing the design control laws. Comple-
tion of the segment of intensive braking in the atmosphere is accomplished
at the maximum lift force wi~h a zero roll angle of the vehicle.
The result of the described program of braking is the injection of the
descending vehicle into the final phase of braking in a relatively short
time and without significant loss of altitude, during which the landing
apparatus begins to operate.
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(t) va,Mi~
10
I
10 - -
~ I i ~
1000 4000 6000 P,, N/~ r~21
7 100 400 600P,ntc n / ~
Figure 4.1. Region of Attainable Descent Velocities for Landings on the
Surface of Venus
I:ey :
1. V~ (m/s)
2. PX in N/m2 (upper scale) and in kg-force/m2 (lower scale)
~L~ N,~,~'y '(2) -
~ P-3000H/r+' i
~ ~ 5000 ~ l000-
0-I ~
~ j
' 0 Q2 04 f~6 qB~a �
Figure 4.2. Initial Descent Altitude for Different Descent Vehicle Charac-
teristics
Key:
1. Hk (km)
2. Px (N/m2)
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nr ma,
10
>6 �
f
8
z
0 ~ p.4 QB �;1 ~P
Figure 4.3. Maximum Load Factor as a Function of Available Lift-to-Drag
Ratio for a 40-1~ Entry Corridor and Direct Descent from an
Approach Tra~ectory. (For Vehicles with a Midsection Loading
of 2,000-5,000 N/m2)
Key:
1. Speed of entry into the atmosphere: 12 km/s
2. Speed of entry into the atmosphere: 11 km/s
nE,~Q.
-�-dH,~ -BOKn
10./ ~ ----dN.?-40Kn
%%e~ --dH,,=10~rn
;6 T-~- ' .
~ ~
_ 8
4 ~:i.
~ � 1
0 QZ Q,4 Q6 I(p
Figure 4.4. Maximum Load Factor as a Function of Available Lift-to-Drag
Ratio for Descent from Near-Planetary Orbits. (For PX of
2,000-5,000 N/m2 and V of 9-10 lcm/s)
Key :
1. Entry corridor of 80 lan
2. Entry corridor of 40 km
3. Entry corridor of 20 lam
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, Ci) C2 ) ~3 )
K,~, nr V, N,KM . ~
i50
14 I ~ ~
/1
1,0 /0 !0 ~Q~
8 V~ H ~
05 5 6 50 - I
_ ,
2 ~ ~ R ~ K'm ~
I 60 ~ ~ .
~ ~r0 �110 16D 700 t, c r4~
Figure 4.5. Parameters of a Descent Tra~ectory with Leveled-off Segment
Key :
I. Segment of entry into 1. Effective lift-to-drag ratio
, the atmosphere 2. V (1~/s)
II. Isoaltitude segment 3. H(km)
III. Isoload-factor segment 4. t(s)
(1) (2)
Y,~ N,,rn
140 j~ d
100 J0p -
160 I ~
~ZO
. 80 f~ 1
~r0 I
2 4 6 8 10 V,r~n/C (3~
Figure 4.6. Trajectory of Descent with Isoaltitude Plateau Shown in the
Altitude/Speed-of-Flight Coordinates
Key :
I. Segment of entry into 1. Roll angle
the atmosphere Z, g
II. Isoaltitude segzaent 3. V(lan/s)
III. Isoload-factor segment
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~
CHAPTER 5
DETERMINING DESIGN PARAMETERS FOR DESCENT VEIiICLES
TO SOUND THE ATMOSPHERE OF JUPITER
The distinguishing feature of entry into the atmosphere of Jupiter is the
great speed of approach to the planet. This speed is mainly dependent on
the value of the parabolic escape velocity, whose magnitude is determined
by the powerful gravitational field of the planet and exceeds 60 km/s.
In analyzing the conditions of entry into the Jovian atmosphere, it is nec-
essary to consider the planet's rotation. This makes it possible to reduce
the speed of atmospheric entry to a value of 49 to SO km/s, having arranged
entry along a shallow trajectory in an easterly direction in the region of
the equator and in the same direction as the motion of the rotating atmos-
phere. The main portion of the vehicle's kinetic energy is reduced during
the process of intensive braking in the upper layers of the atmosphere and
before reaching the cloud layer. The distance from the planet's center to
the cloud layer is usua.lly assumed as the base or reference level of calcu-
lations, and, as in Earth conditions, the pressure at this level is said to -
be 1 atmosphere.
Most current models of the Jovian atmosphere assume an insignificant varia-
tion in temperature with altitude in the upper layers of the atmosphere, and
the composition of the atmosphere is assumed to be basically hydrogen (70 to
85 percent) and helium (30 to 15 percent). _
Implementation of descent, whose goal is the delivery of instrument contain-
ers for measuring atmospheric parameters, is complicated by the high speed
of entry into the atmosphere, with attendant intensive thermal and load-fac-
tor effects on the vehicle.
In aerodynamic heating, radiative heat flows from the shock layer predomi-
nate, and, as the vehicle is braked very hard with load factors reaching
hundreds of g's, the thermal effect on the surface coating has the character
of a thermal shock.
In contrast to descent vehicles entering the atmospheres of Earth, Mars, and
Venus, a descent vehicle for sounding Jupiter and other giant planets is
subjected to a much more powerful thermal effect, and thermal protection of
such a vehicle is one of the basic problems in ~_qigning it. Choice of a
descent-vehicle shape constrains the task of minimizing L::e weights of
structure and of thermal protection.
Let us imagine what the flight regimes and the shape of such a vehicle will
be and in what respects this vehicle must differ from traditional descent
vehicles that land on the nearby planets. It is logical to assume that for
the first experiments in sounding the Jovian atmosphere nreference will be
given to simple and proved designs and to reliable technical solutions. We
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can also assume that the vehicles that will be initially designed will per-
form ballistic, uncontrolled descents analogous to the initial Martian,
Earth, and Venusian descent vehicles.
Among the systems necessary for descent and which will affect the choice of
the vehicle's shape, structure and thermal protection are basic. The struc- "
ture includes the envelope, the primary structure of the body, fittings for
the payload, devices for attaching apparatus, hatches, and structural
frames. The vehicle's envelope can be nonhermetic, but it must be rigid in
order to withstand the dynamic loading during oscillation af the ~~ehicle at
the moment of atmospheric entry. Instruments and equipment, however, have
to be located in hermetically sealed sections. The lifting portion of the
envelope (the heat shield) affects stability due to external pressure, which
attains magnitudes on the order of tens of atmospheres at the maximum
dynamic pressures. Thermal protection includes the ablative surface, the
- heated thermal protection coating, and thermal insulation. Powerful thermal
flows require the use of thermal protection coatings possessing high break-
down enthalpy, a significant working temperature of sublimation, and good
solid-phase reflectivity at the exterior breakdown front of the coating and
high absorption characteristics of the products of the breakdown of the
coating that enter the boundary layer. The thermal-protection work process
is accompanied by a significant expenditure of the coating mass as a result
of which there forms in the boundary layer a bubble of pyr~lytic gas, con-
sisting mainly of products of the breakdown of the thermal protection coat-
ing
~ The influx of this gas, which is cooler than the hydrogen-helium mixture in
the shock layer, blocks convective heat flows, and a sheet of the products
of the breakdown absorbs a part of the energy that radiates from the shock
layer in accordance with the integral spectral coefficient of ab~orpt~.on.
The linear rate of breakdown of the thermal protection material at this
instant is so great that the propagation of the breakdown front into the
- depth of the thermal protection surface precedes the motion of the "heating
wave," that :.s, a movement of the layer with the limiting admissible temper-
ature for the structure. -
Flow about the vehicle during the entire basic segment of the descent tra-
jectory is hypersonic, and the aerodynamic braking force is determined by
the pressure distribution over the surfaces of the vehicle, where this force
is generated by a spherical or conical heat shield of either blunted or mod-
erately pointed shape. The aft fairing of the vehicle located in the dead-
flow zone can carry spoilers, which provide passive stabilization of the
vehicle for a misoriented entry into the atmosphere aft end first. ~
In choosing design parameters, one should assume that the vehicle should be
capable of moving along any trajectory within the design corridor of entry
into the atmosphere. In the present case, the upper boundary of the corri-
dor of entry is determined by the capture by the atmosphere, while the lower
boundary corresponds to maximum loading. For vehicles with different shapes
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tt~at have d~fferent ballistic characteristics, capture by the atmosphere can
be determin~=d by the minimum value of the specific load for a given choice
of model atmosphere.
In searching for the design parameters of the descent vehicle, it is neces-
sary to estimate the mass of structure and of thermal protection for vehicles
with diff erent configurations. In accordance with a general formulation of
the design-search problems for descent vehicles, each such design estimste
has to be carrierl out through a comparison of all possible conditions of -
motion with the object of eliminating those that result in the maximum values
n for the mass of structure and thermal protection material. It can be
- assumed, and the results of calculations for specific vehicle shapes support
such an assurnption, that the maximum mass of ablated thermal protection mate-
rial will occur for a trajectory of deep immersion into the atmosphere that
differs from flight during entry along the upper boundary of the corridor.
The greatest mass of thermal insulation will be required for the longest tra-
jectory, which occurs for motion close to the upper boundary of the entry
corridor. Maximum external pressure~on the lifting envelope of the heat
siiield, which is the controlling case for design of the shell, will occur at
the highest dynamic pressure, that is, at the moment of maximum load factor
for the steepest allowable trajectory. Let us formulate the assumptions
made in the design calculations for the case of searching for best parameter
values that will determine the shape, design elements, and conditions of
motion of the descent vehicle. Basic among these assumptions are the fo~-
lowing:
1. In calculating volumetric and centering characteristics, it is enough to
use the dimensions of the vehicle's outline. Geometric parameters that can
, be varied during selection of the shape are the outlines of the external
thermal protection layer before entry into the atmosphere.
2. In calculating the aerodynamic characteristics, we can assume that the t!
shape of the vehicle remains constant, that is, the linear ablation and ero-
sion of the surface can be neglected in calculating vehicle drag.
3. Ballistic characteristics and, in particular, loads on the midsection
are determined along the descent trajectory with consideration for the
dimensional changes of the lifting profile due to the linear ablation.
Simultaneously, in calculating the load on the midsection, consideration is
given to the decreasing weight of the vehicle during flight due to the
decreasing mass of the thermal protection coating.
4. Tn estimating heat flows, it is necessary to account for the increase in
the radius ~f bluntness due to the varying ablation rates of the mass of the
coating at the stagnation point and along the periphery of the heat shield.
This circumstance is especially important for small initial radii of blunt-
ness for conic frontal shields. Optimum initial radius of bluntnessy deter-
mined from the relationship of radiative and convective heat flows, is on
the order of 0.10 to 0.20 ~m for vehicles with a midsection loading of from
2,000 to 5,000 N/m2 (200 to 500 kg--force/m2).
5. Inasmuch as the basic part of the descent trajectory (during which the
vehicle passes through maximum loading and thermal flows, and the kinetic
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energy is reduced Ly more than 99.$ percent) occurs at hypersonic flight
velocities (for Mach numbers above the 4-to-6 range), it is adequate to use
Newton's hypothesis for determining the aerodynamic characteristics. The
ratio of static pressure to dynamic pressure at specific points on the sur-
face is assumed to be constant along the descent tra~ectory under this con-
dition.
6. Ablation of the thermal protection coating has to be determined consid-
ering the emission of radiation and the blocking of convective thermal flows
by the layer of gas formed by the breakdown products of the thermal protec-
tion material.
The assumptions listed are, in our opinion, acceptable for estimating the
conditions of motion and aerodynamic and weight characteristics of the
vehicle during design calculations, i~o-re complex models of the flow and
heat exchange are used for verification calculations during this stage of
design.
In Figur.e 5.1 are shown the parameters of a descent trajectory for a vehicle
with conic heat shield, which, b efore entry into the atmosphere, has a load
_ of 1,720 N/m2 (175 kg-force/m2) on the midsection, a bluntness radius of
0.2 m at the stagnation point, and a cone half-angle of 70�. Also shown in
the figure is the variation in load on the midsection for such a vehicle due
to abla~ion of the thermal protection coating during flight.
The search for the best values of parameters that determine the shape of the
vehicle is broken down into a series of design calculations, during each of
which ~aass characteristics are determined. Each of these design calcula-
tions is iterative, because in the algorithm for computing weight or mass
~_t is necessary to assune the ballistic parameter, which, in turn, is deter-
mined by the weight or mass of the vehicle. Also, as shown earlier, deter-
mining the mass of the system should consider the worst case for each sys-
tem, that is, the worst of all possible descent trajectories. One should
also consider that for each new configuration of geometric parameters of the
vehicle it is necessary to search for the limiting allowable trajectories -
corr.esponding to the upper and lower boundaries of the entry corridor. All
the material presented above establishes the need for repeated tra~ectory
calculations during a single pass through the design calculations. It is
possible to reduce the labor in such an operation and the time lost if one
foregoes accuracy in estimating descent trajectory characteristics and does
not integrate the equat~.ons of motion. For estimating the mass character-
istics, it is usually necessary to know the distribution af the loads on the
vehicle (structural mass), flight speed, and density of the atmosphere at a
given altitude (thermal flows and mass of the thermal protection layer) as
functions of time. For design esti?nates it is desirable to determine rap-
idly the limiting trajectories, that is, ~ra~ectories with the largest and
smallest values of maximum load factor. For such purposes, a method can be
proposed for estimating trajectory parameters that uses the thermodynamic
relationships involving load factor and flight time.
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During ballistic descQnt, the load factor initially increases sharply and
then, having passed a maximum, decreases gradually. As a rule, for maximum
permissible load factors in excess of 100 g's, a single load-factor maximum
occurs, and a typical expression for it (as a function of flight time) is
conveniently given by a formula of the type
n = aT2e-cT [5.1]
where n is the load factor; T is the time of ~light; and a and c are coeffi-
cients that determine the shape of the load factor curve, which depends on
- the assumed model for the atmosphere and the conditions of entry into the
atmosphere.
If in the equations of motion we assume a load factor much greater than
unity and a flight-path angle during the main segment of braking in the
atmosphere as close to zero, integration of the equations of motion in
velocity coordinates yields the following expressions for velocity and
atmospheric density as functions of time:
V l~'Bx - 9,8a f e~T ( Tz - 2~� ? l - ?
~ ; .
~ c c2 c3 ~ c3 (5. 2)
~ e ^ Ps~Pn 1 ~
[v (T)j2 ' (5. 3)
where
c = 2/TB; a = 7.4n~X/TB
VBX is the speed of entry into the atmosphere; tB is the instant of passing -
the maximum load factor; and n~~ is the maximum load factor. This load
factor is less than n~~ g, the limiting one for the vehicle (for the lower
boundary of the entry corridor), and greater than n~X the minimum per- -
missible one for the condition of capture of the vehicle by the atmosphere �
(upper boundary of the entry corridor); and Px ~p is the mean value of load-
ing on the midsection for this trajectory. As the minimum permissible load
factor, nmaX B, we will assume 100 to 150 g's.
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Figure 5.2 shows the results of a precise calculation of trajectory, namely:
Maximum load factor as a function of the angle of entry into the atmosphere
for a vehicle with a midsection load of 1,720 N/m2 (175 kg-force/m2) and for
a nominal atmosphere model. Reaching an altitude of 400 km above the cloud
layer is assumed to be the instant of entry into the atmosphere. The same
figure shows the period for braking required to reduce velocity to 1 km/s _
and the relative mass of the ablated thermal layer for different conditions
of atmospheric entry for a vehicle that has a conic heat shield with a cone
half-angle of 70� and a bluntness radius of 0.2 m.
In the general case, the levels of maximum load factor are related by an
equation that, for design calculation, is assumed to be linear. The maximum
load factor within the corridor of entry is given by the expression
nmax - nmax B ~1 - ~~8) (5.4)
where ~8 is the increase in the angle of atmospheric entry and d~termines
the location of the given tra~ ectory with respect to the upper boundary.
The coefficient Kn depends on the assumed atmosphere model and, for practi-
cal purposes, does not depend on the vehicle's baZlistic parameters. For
design calculations, it can be assumed that K~ = 0.19 + O.U3. The instant
of passing the ma~imum load factor is determined within the bounds of total -
braking time Te, which in turn depends on the velocity impulse and maximum
' load factor, as follows:
~ =1C~T., ,c, _ vaX K
� ~a~ (5.5)
nmaz 9, 8
where K~P is the coefficient determining the mean load factor or degree of
"fill" of the load-factor diagram; K~P = 0.3 to 0.36. The coefficient K.~
depends on the assumed atmosphere model and lies in the range from 0.31 to
0.34.
Methods of design estimation of the mass of the ablative and heated thermal
protecti~n ldyers are based on recommendations in references from homt and
abroad (Refs. 1, 61, 62, 63).
The main difficulty in calculating the ablation of the thermal protection
layer is the need to account for weakened convective and radiative heat
transfer in the layer ef pyrolytic gas formed by the products of the break-
down of the coating. Estimating the race of decrease in the mass of thermal
protection, under conditions when convective heat transfer is blocked by
ablation and the radiation is frequently absorbed in accordance with the
integral spectrum of absorption of the breakdown products of the thermal
protection layer, can be done by using expressions for the attenuation of
heat transfer computed for different surfaces. These expressions take into
account the feedback in the mechanism of coating breakdown and show how the
relationship between thermal-protection-mass ablation and the rate of mass
change in the incoming flow affects the relationship between heat transfer
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through the vapor barrier of the coati.ng and heat transfer due to radiation
from the shock layer or due to tran5port by the gas in the boundary layer.
Figure 5.3 shows how the convective (qk) and radiative (qr) heat transfer
and ~ass of the ablating thermal protection layer vary as a function of time
for a trajectory with a maximum load factor of 200 g's for a vehicle of mod-
erately pointed shape with a cone half-angle of 50� and midsection loading
of 2,940 N/m2 (300 kg-force/m2).
To draw any conc:lusions about the advantages of one or another aerodynamic
form solely on the basis of calculations of thermal protection layer abla-
- tion (as is ur.fortunately done in a number of references) is premature, in
that the total mass of the vehicle depends on the thickness of the heat-sink
insulation, on the thickness of the envelope, and on the dimensions of the _
structural components of the vehicle body, as well as on the dimensions of
the trim tabs of the skirt, or the ballast mass. Only by integrated
accounting for all indicated factors and by considering volime-centering
character~stics and heat-shield strength is it possible to judge competently
the advantages of one or another aerodynamic shape.
Calculations show that for substantial pointing of the nose cone the volu-
metric characteristics of the vehicle worsen as the available volume and,
~herefore, surface of the vehicle increase. At the same time, the surface
use eff ectiveness decreases, that is, the total surface and the heat-shield
- surface are increased, which can lead to a sharp increase in the mass of the
structure and of thermal protection. However, in this case, the specific
weight or mass of one square meter of heat-shield structure is reduced due
to the increa~e in the curvature of the envelope, which provides resistance
to external pressure. Radiative and convective heat transfer on the side
surfaces of a slender cone are substantially smaller than in the stagnation
region, but as the shape becomes slenderer, heat transfer in the stagna-
tion region increases sharply because the vehicle, having small resistance
and cross section (that is, a greater midsection load) moves along a steeper
trajectory with higher thermal stresses at a given level of maximum permis-
sible load factor. Also, it is necessary to account for ballast mass and
means for passive stabilization (trim tabs and skirts) that increase for a
slenderer nose cone.
Apparently, if no account is taken of other circumstances (descent vehicle -
- layout as part of the spacecraft, presence of experimental and engineering
sections, degree of reliability in design estimates, and so forth) and
weight or mass criteria are used in choosing vehicle shape, then the optimum
shape for a vehicle with a mass of 500 to 2,000 kg is the shape with a mod-
erately slender ccnic heat shield having a cone half-angle of 45 to 55� and
a bluntness radius of 0.15 to 0.25 m. The choice of diameter or slenderness
ratio of the vehicle depends on the method used to provide static stability
for noncontrolled entry into the atmosphere.
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One should exp~ct that the mass oF structure and of thermal protection will _
be 110 to 140 percent of the payload mass for a descent sounder vehicle
delivering measuring instruments into the Jovian atmosphere from an approach
tra~e~tory, with the mass of thermal protection coating ablated being 20 to
30 percent of the total vehicle mass.
~)uestions of reliability and precision of design estimates are quite impor-
tant for vehicles such as Jupiter sounders being designed for flights under
_ new 3^.d unknown conditions. The uncertainty of operating conditions, due to _
]ack oE knowledge of many paramete:s af the atmosphere and planet as well as
to unusual, extreme conditions of entry into the atmospher e, makes caution -
necessary in using the results of such design calculations based on any
metho~s, no matter ho~~ complex, because all methods have co use initial data
that have substantial scatter in their values. Under such conditions, it is
necessary to supplement the traditional and known design c riteria by risk
criteria, which directly or indirectly detennine the possib le variation in
values of design characteristics. These risk criteria als o define the -
de;ree to whicn the scatter in specific parameter values influences the
scatter of design criteria values and the level of reliabi lity of the rela-
_ tive estimates based on comparisons of possible values of s uch design cri-
- teria to the standardized limits established for them.
- In the present case, ~he scatter in the values of mass or ~,Teight character-
istics is primarily due to the scatter in data that determine both the
atmosphere model and the parameters defining the conditions of atmospheric
entry. The variations in the values of the characteristics of structural
and ther~al protection materials (effective enthalpy, breakdown temperature,
degree of surface blackness, and so forth) and the scatter in the aerody-
namic characteristics of the vehicle have a definite effec t on the scatter
in the mass c~iteria.
The distribution histogram shown in Figure 5.4 was obtained from statistical
modeling of the scatter in the mass of structure and therma.l protection for
one type of descent vehicle. Relative mass indicators are shown, with the
mass of the d~scent gear given as a fraction of payload mass.
One criterion that can be used to estimate risk in making decisions based on
_ information from design calculations =s the probability of coming up with
design estimates that would lead to poor or simply impossib le vehicle
designs. Let us assume that there exists a certain limit on the total
vehici~~ mass expressed as a fraction of the interplanetary system mass. On
the other hand, there also exists a well-defined minimum ma.ss for the -
instruments needed for scientific exploration. The limiting boundary for
the mass of the structure with thermal protection is shown in Figure 5.4,
- and this boundary is determined by the minimum mass of the payload.
A situation can occur in which the mass of both the structure and the ther-
- mal pratection of the vehicle exceeds the boundary establis hed for it by the
_ payload. In our example, the probability of such an event is quite high
(P = 0.08).
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A ctiange in ttie vehicle parameters or the establishment of better substanti-
_ ated and ad~~antageous weight reserves on part of the enti-re space complex
will make it possible to reduce the decision-making risks during the prelim-
inary d~sign stage. Criteria for estimating risk when design estimates are
uncertain and examples that illustrate the means of applying such criteria
are described, for example, in Reference 23. -
(1) (z) C3) '
f; P� H/ro r .
_ NM~'(' 7M
~ I
_ 1B?0 L I
J50 P' ~
I
J~ J00 ~500 ' - - I
S 1
?u^ jp~ 1400 ~ H i- ~
' I ~
1G !pp Jt'00 ~ i
50 i n ~
I �
0 10 40 60 80 Z,c
diPm, 8i > E2,
i
The stabilization power plant thrust takes on the value Ti = Ki for t< t~;
and Ti = 0 for t>_ ti , when d2 > 0(i = 1, 2, 3), where ti is the time
of cut-�off for i-th p~wer plant.
The sliding speed of the i-th landing-gear leg along the surface is deter- ~
mined by the expression
_ Xi = X+ YY - Si sin (Y - e)� (6.28)
Depending on the character of motion of the landing-gear pads, the compdnent
forces FN and FT are determined for specific cases that frequen*_ly arise,
i i
as follows: (1) The energy absorber collapses and the i--th landing-gear
leg slides; (2) the energy absorber collapses, but there is no sliding;
(3} the landing-gear leg slides, but there is no collnpse.
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Characteristically, ~Xi~ > el; ~dil < e2; Fi (n - 1) ~ 0. In this case, we
obtain a system of diff erential equations for each landing-gear leg; for
example, for leg 1:
mXl~. -FL, sec;~l cos �1=F,v,-~F,v,-W� cos 6-{-
-~-"1YYz-(T~~Ta~'Ta) cas(Y-8),
?WY-}-F~, sec 3, (Xl cos �l -Y sin �l)- (6.29)
_ ~FT. -I- FT. ) ~-i- FN,�Y_ - F,v,X3 T 1L1- TzL2 T3L3. -
By solving the corresponding system, we determine Y and F If F > 0,
L1~ Li
then the value calculated for FL is used for the inte~ration of the system.
i
If FLl < 0, then we should assume FL = 0. The angle ui is computed by
i
using the formula
Xi -
Ui = -u~ ,
(Xi~
Using the found values of ui and FL1, we evaluate si = y- 8+ ui; Fi = FL
i
sec si; FN1 = Fi cos ui; FTl = Fi sin ui. (4) The landing-gear leg neither
collapses nor slides. This is characterized by
~Xi~ ~ E1' ~Si~ ~ E2~ Fi (n - 1) ~ 0.
F'or each landing-gear leg, we obtain a system of equations, for example, for
leg 1:
FT~-}-mYy= -FT.-FT,-~-IY/~ sin 6-mXly~-
-~Ti~Tz~'Ta) sin (y--8); -
~ .
FN, - mX lY F~-~ - F,v, W ~ cas 9- mYy2
-~-(T~ ~-?-Tz-I-Ta)cos(Y-8); (6.30)
?cvY -f - F~v,X i - Fr,Y = (Fr. FT,) 3'
r
~ -f-~`'N.1Cs-FN,X,-f-T~L~-T~2~T3L3�
By solving the corresponding system, we determine y, FTi and FNi. The angle
ui is determined by using the formula ui = arc tan (FT /FN For
i i
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F~ > 0, we have (1) if lui~ < up; then the computed values for FN ar~d FT
1 1 1
are used in integrating Equation 6.26; (2) iF luil > u~; the solution is not
valid because the landing gear must s lide (and case 3 above applies). For
FN. < 0, we should ass~e FNi = 0 and FTi = 0.
i
In this chapter, we previously consid ered the question of determining the
stable attitude of the space vehicle following a soft landing on a planet. -
In solving the planar problem, the cr iteria of capsizing of the space land-
ing vehicle wil]. appear slightly diff erent from the case of three-dimen-
sional vehicle motion. Tndeed, the space vehicle moving in a vertical
plane loses stability during soft landing if:
al > 0 or a2 < 0 (6.31)
where
- ai = arc tan (Xi/Y) + 6(for i= 1, 2) (6.32)
and where al(a2) is the angle between the local vertical and the line pass-
ing through the center of the mass of the vehicle and the point of contact
for leg 1(2j with the surface; Xi is the distance from the center of mass of
the vehicle to the line of action of the force FNi (Xi > 0 when the refer-
ence direction coincides with the X-axis).
The space vehicl~ will re~ain in a stable attitude following landing on the
surface of a planet if the follAwing conditions are met simultaneously:
ai0
B
0 u T~ d"=0 ~ `
/ 'z y~syy,
' ~ J 4 z
c ~
1 _
a ~
X
B
Onopa 4 X
(2) .
� . ,J Vc~ 4
Onopa J 1 ' ` U X On,~pa f
2 V~ ~ ~ Z X (Z)
~2~ 4 3 '
R
D� ,
. i .
- Z~Z' W
Onopa 2
~2) ~
Figure 6.1. Space Landing Vehicle Geometry
Key :
_ 1. Center of gravity
2. Landing-gear leg
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Bb/4l/CpPN[!t JNQVINlIII . B6/4/LCA2N!!t
napanempo6np~r cu~ 6 cr~cmene
' t-0 t+nGt ~2~
P~!!ltNUt ,SlpQ6M2NlLlL p'6[lN/PN((A
~r_~+~ d/IAOpQf(IfAfNlLB JMQ4BH(l/l
nap�nempoa np~
t+(~+f)�dt (3)
~ .
~ Figure 6.2. Block Diagram for the Integration Stages
Key :
1. Computation of parameter values at t= 0
2. Computation of forces in the system at t+ nAt
3. Solution of equations of motion to determine parameter values at
~ t + (n + 1)~t
f�.~ ~ . .
~ f .
~ ~
F . .
1
Fj -fj-r ~ 1 KJ .
, ~ 1 - t- f
I t' :
~ i F, _
~ ~ ~ .
, _f.a,K. ~ I ,
I K,
~ I
fo
d.r d~~j.,~ d d~lJ ~Cma~nr~t d pacm~
Nue 2.
. ~
Figure 6.3. Staircase-shaped Force-atrain Characteristics
Key:
1. Compression 2. Tension
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rok oHFrcr.nr. ~rsL oNLr
0
- x
~ Fl1
_ fii.( /ni~s -
- ~ - ~E-
frtl
~ t
Z f mi I,t .
r
fh~
0
~
~foir ~
Figure 6.4. Forces Acting on the Landing Gear Pads of a Space Landing
Vehicle. (mi is the equivalent mass of the pad of the i-th ~
landing gear)
~ ~
ioa
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. (1) 4
J' ~F N' 4G77N0~
u ~
3 f ~
~
z w ~ z
w .
3 (2) J
N-Mevrmyx
U U
, Z ~
4�
. ~ w ~ . t W
a ~ d ~
Figure 6.5. Determination of the Value of N
Key:
1. For even values of N
2. For odd values of N
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~ .
1 v' y ' . .
X, ~ .
.r xz
~ ~u.rt~ (z)
L~ t0~([JVNm
~ d'~0
1 Tr
~ ~ W~
~ , ~ ~ -
a
~ / N ~ az
x'
' \
. F' X, 1 ~ L`` ~d1 1 B
y4 a ~ at m ~ . n.
x~ v d d d d a
~ 1O ~ " ~ ~ (
= o~ o~ c~ ~
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Key:
1. Vefiicle name
2. Date of soft landing (day, month, year)
3. Space landing vehicle mass (kg)
4. Landing velocity
5. Vertical (horizontal) ~
6. Surface slope. (deg. )
7. Landing location
8. Landing-gear characteristics
9. Luna-9
10. Luna-13
11. Luna-16
12. Luna-17
13. Luna-20
14. Luna-21
15. Luna-24
16. Surveyor-1
17. Surveyor-3
18. Surveyor-5
19. Surveyor-6
20. Surveyor-7
- 21. Apollo-11
22. Apollo-12
23. Apollo-14
24. Apollo-15
25. Apollo-16 ~
26. Apollo-17
27. Mars-3
28. Mars-6
29. Viking-1
30. Viking-2
- 31. Venera-7
32. Venera-8
33. Venera-9
34. Venera-10
- 35. Data not determined
36. Ocean of Storms, lat. 7.1� N; long. 64.4� W
37. Ocean of Storms, lat. 18.9� N; long. 62.0� W
38. Sea of Plenty, lat. 0.7�S; long. 56.3� E
39. Sea of Rains, lat. 38.3� N; lon~. 35� W
40. Crater Apollonius C, lat. 3.5� N; long. 56.5� E
41. Sea of Serenity, lat. 25.8� N; long. 30.5� E
42. Sea of Crises, lat. 12� 45' N; long. 62� 12' E
43. Ocean of Storms, lat. 2.5� S; long. 43.2� W
44. Ocean of Storms, lat. 3.0� S; long. 23.3� W ,
45. Sea of Tranquility, lat. 1.4� N; long. 23.2� E
46. Central Gulf, lat. 0.5� N; long. 1.4� W
47. Crater ~cho, lat. 40.9� S; long. 11.5� W
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48. Lat. 0� 41' 15" N; long. 23� 26' E
49. Lat. 3.036� S; long. 23.416� W
50. Lat. 3� 40' 27" S; long. 17� 27' S8" W
51. Lat. 26.10� S; long. 3.60� E
52. Lat. 9� 0' S; long. 15' 35' W
53. Lat. 20.2� N; long. 30.7� E
54. Electris, lat. 45� S; long. 158� W
55. Erythraeum Mare, lat. 24� S; long. 19.5� W
56. Chrysc, lat. 19.5� N; long. 34� W
57. Utopia, lat. 44.3� N; long. 10� W
58. Lat. ~ 2� S(night side} and - 2,000 km from the morning
- terminator
59. Lat. ~ 10� S(day side) and ~ 600 lmi from the morning terminator
60. Lat. 31� 42' N; long. 290� 50'
61. Lat. 16� 02' N; 1ong. 291�
62. Balloons with compressed gas (spherical)
63. Four-legged, strut construction, tripod ehape
64. Three-legged with hydraulic shock absorber, tripod shape
65. Four-legged (deployable) cantilevered strut construction with
honeycomb shock absorber
66. Soft-landing engine and shock-absorbing structural elements
� 67. Three-legged (tripod shape)
68. Shock-absorbing structural elements
69. Toroidal envelope
70. (Calculated value)
there is no sensible force of gravity would be very difficult without
"clamping" engines).
We now present somP es~imates of the effectiveness of stabilizing engines
for the most crit~cal landing conditions ~downward along a steep slope,
a= 0; 6 X=-~~'). "Clamping" engines are installed on each of the vehi-
cle's lan~ing-gear legs (Figure 6.6). The duration of engine operation is
limited. It is obvious (Figure 10.7) that using stabilizing engines with a
total thrust of
4
T~ _ ~ Ti = 10,000 N
i=1
enlarges the stability zone of the space landin~ vehicle twofold, and using
engines with T~ = 20,000 N increases iC threefold.
Calculations show tha.t no stability zone exists for landings on surfaces
with a slope of -30� (Figure 10.7) unless stabilizing engines are used (in
which case, the static stability of the vehicle is preserved). Using sta-
bilizing engines, a stable landing on a surface with a slope of -30� and
within the required range of velocities can be achieved. -
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The change as a function of time of some parameters that characterize the
dynamics of soft landing of a vehicle with a four-legged landing gear on a
slope (8 =-20�, landing con�igurution 2--2, a= 0�) is shown in Figure 10.13
(for landing velocities Vvert - 3 m/s and Vhqr= X' - 1 m/s) and in
Figure 10.14 (landing velocities Vvert � Y' - 1 m/s and Vhor = 1 m/s).
In these figures, solid lines represent parameters during the landing proc-
ess when no stabilizing engines are used, while dot-dash lines show param-
eters when engines with a total thrust of 10,000 N are used. The character-
~stic moments during the soft landing are slzown in tt_e lower portion of the
Figures.
The variation of parameters for unstable landing when no stabilizing engines
are used (loss of stability at the instant t= 1.65 s, with a2 < 0) is
shown in Figure 10.13. A"clamping" engine assures stability of landing (at
t= 1.35 s; a2 = 0.26 rad; E= 58 J).
The change in parameters for the case of a stable landing even in the
absence of stabilizing engines is shown in Figure 10.14. The use of stabi-
lizing engines leads to a faster completion of the soft-landing process.
Let us consider the main events of this process in more detail.
l. Landing-gear leg no. 1(the doubled leg) touches the landing surface and
is defo rmed (t = 0 to 0.1 s). As a consequence of deformation of landing-
gear leg no. 1 and the presence of friction with the surface, the kinetic
energy of the vehicle decrzases (from 5,800 to 4,500 J). Deformation of
landing-gear leg no. 1 equals dlm = 0.007 m.
2. The space landing vehicle rotates about the pad of landing-gear leg
no. 1(t = 0.1 to 0.7 s). The center of mass of the vehicle continues to
drop until the instant when deformation of landing-gear leg no. 1 ceases.
Because of this, the kinetic energy of the vehicle increases (from 4,500 to
9,210 J).
3. I,anding-gear leg no. 2(also doubled) touches the landing surface and
is deformed (t = 0.7 to 0.8s). As a result of the deformation of landing-
gear Ieg no. 2 and because of the presence of friction with the surface, the
kinetic energy of the vehicle again decreases (from 9,210 to 1,400 J).
Deformation of landing-gear leg no. 2 equals d2m = 0.17. When the pad of
landing-gear leg no. 2 touches the surfare, the pad of landing-gear. leg
no. 1 lifts off from the landing surface.
- In accordanc_e with the assumed schedule for starting the stabilizing engine,
the engine is started at the instant the surface is touched by the pad of
landing-gear leg no. 2. The change in the process parameters that results
from starting the stabilizing engine until the instant the stability crite-
ria (Equation 6.33) are satisfied is shown (by the dot--dash line) in Figure
10.14.
4. The vehicle rotates about the pad of landing-gear leg no. 2(t = 0.8 to
- 1.0 s, when stabilizing engines are used; and t= 0.8 to 1.85 s, when
- 175
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stabilizing engines are not used). During such rotation, the vehicle center _
of mass rise~ and, consequently, its kinetic energy once dgain decreases
(from 1,400 to 50 J).
7n accordance with the stability criteria (Equation 6.33), when al < 0,
a2 > 0, and E< n= 50 J, the pxocess of landing the vehicle is considered
completed at the time t= 1.85 s(when no stahilizing engines are used) or
at t= 1.0 s(when such an engine is used},
The changes in srability zones for landings o~ variant B space landing vehi- _
~le on inclined :;ur�aces are shown in F~gures 10.1 to 10.3 ~s a funetion of
the force of resistance to the cozapressxon o� the landing~gear legs = 10
to 30 kN). The figures show tha.t for landings on inclined surfaces (6 =-5
- to -15�, a= 0), the vehicle's stability zane narrows significantly for
increases in the f.orce of compression on the landing--gear legs (especially
ror large values of landing velocity). The r.eason for this is the increase
in the toppling moment for the vehicle, which accompanies increases in
stiffness of the energy absorbers in the landing-gear legs. _
In gresenting this material, we have covered quite a broad range of landing
velocities ~~vert ~ 0 to 8 m/s and Vhor - 0 to 5 m/s) that is significantly
greater than the one that can be practically provided by the control systems
for normal ~andings of vehicles on planetary surfaces. Also, as previously
discussed (Chapter 6), the deformation of landing-gear legs was assumed to
be unlimited (in order to sinplify calculations).
I,et us no*_e that choosing a magnitude for the force of compression acting on
the landing-gear-leg energy absorbers is appropriate not only for determin-
ing the stability zones for vehicles during landing, but also in considering
the permissible load factors for the payload and for defining realistic mag-
nitudes of. deformation in the landing-gear legs.
The relative simplicity of the theoretical method presented in Chapter 6 for
studying the dynamics of soft landing (and, consequently, the small expendi-
ture of machine time when making calculations), and the ability to account
tor the effect of structural parameters of the vehicle and of initial land-
_ ing conditions ~n the dynamics of landing as well as a good agreement of
results with experimental data make the method fully acceptable for engi-
- neering calculations in design offices for tt~e purpose of refining the
detai.led designs o.f, space landing vehicles.
Let us now turn to the specifics of deformation of the landing-gear legs for
a space vehicle with different landing conditions. In reviewing the compu-
tational and e~perimental data on the deformation of landing-gear legs, we
note the following:
1. Deformation of all landing-gear legs of a vehicle ~.ncreases with an
~ increase in vertical landing veloci.ty.
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2. For vehicle Iandinbs downward along a slope (a = 0�, configuration
1-2-1), the largest deformation occurs for the leading landing-gear leg
(no. 3) and the smallest for the trailing lzg (no. 1). The deformation of
leg no. 3 is two- to 2.5-fold greater than tr.e deformation of leg no. 1.
Leg nos. 2 and 4 have approximately the same deformation, which is 1.5- to
tworold smaller than the deformation in leg no. 3. The deformation of leg
nos. 1 and 3 increases for increasing angles of the surface slope. -
3. In landing a vehicle up a slope (a = 180�, configuration 1-2-1), the
largest deformation is experienced by the landing�-gear leg (no. 1) upperntost
on the slope, and the least is experi.enced by the lowest Ieg (no. 3).
Deformation of landing-gear leg no. 1 is two~ to 2.5-fold greater than the
deformation of leg no. 3 and 1.2- to 1.5-fold greater than the deformation
of leg nos. 2 and 4. For increasing surtace slope angles, the deformation
of leg no. 1 increases and that of leg no. 3 decreases. This case of motion
represents the greatest danger 3uring a vehicle landing, from the viewpoint
of the possible failure of the leading leg no. 1 caused by the large defor-
mation of its energy absorber (Figure 10.15).
4. For landing a vehicle across the slope (a = 90�), the contact configura-
tion oE the supports is 1-1-1-1. For this, the greatest deform~tion occurs
in the leading landing-gear leg, no. 4, and the smallest deformation in the
trailing leg, no. 2. This is explained by the direction of the horizontal
velocity vector in the direction of leg no. 4. The deformation of leg no. 4
is 2.5- to threefold greater than the deformation of leg no. 2 and 1.2- to
1.5-fold greater than the deformation of leg no. 1. For increased surface
- slope angles, the deformation of the leading leg (no. 4) and the leg upper-
most on the slope (no. 1) increases, while the deformation of leg nos. 2 and
3 decreases.
S. For landings of the vehicle in configuration 2-2 downward along the
slope and for different values of the component of landing velocity, the
- deiormations of the leading landing-gear legs (nos. 3 and 4) are 2.5- to
threefold larger than the deformations of the trailing landing-gear legs
(nos. 1 and 2).
6. The character of the landing-gear-leg deforma.tions of various vehicles
is identical. However, the absolute magnitude of deforn~ation of the legs
decreases with increasing stiffness in their energy absorbers.
A summary of experimental values of acceleration and deiormation of landing-
gear legs, recorded during tests o� the dynamically simi.lar DPM-B model for
N= 3 and 4 during landings on a slope of 6=-5� (a = 0�), is shown in
Table 10.2. _
It is interesting to note that the total time for the landing process of the
model (including rebounds following initial contact with the ground) depends
to a large degree on the ma.gnitude and direction of the landing velocity and
equals from 0.55 tc 0.75 s, where the upper limit applies to the four-legged
, landing gear. The maximum acceleration at the ~odel's center of mass is
achieved approxiznately 0.1 s(recalculated for the actual space landing
vehicle, this value is 0.6 s) after the first contact with the ground. Of
course, the indicated time spans apply to the landir.g of the dynamically
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Table 10.2
Ncnw- Cxea~a N ycnoeNS naxaaax~ix AaryH-
TaHNE ~ HCiibITdHHH KOB yCKOpeHN~ AY rIPNblE93HN2
~1) ~2~ N II2p2?(~3)HNN ~R (4)
B i(. M. ~ 9~ $ OROpbi sanohceew
G,.=6 g (!O) co~r~ sHeprce~xorrba a
~ (5) G.a4,5g (il) ~2 pa3a 63nbiuex, qeH
~ Ha onopax npH nocneuy~o?nxx xcnw-
GII=7 g T2HHAX. AYIII t~ ,IIY1V
GIII=4 g noxaadsaaT 803HNKHOSC-
GIV =3,6 g(12~ xxe xone6aHx~i acne,qcr-
30 ~Qp~_ bII=85 MH Bxe cyxoro Tpexxa
~_t~'JoCE~� aIII�45 M41
i BIV~~ M31 ~i~j)
vr�~~7 M~C ~~J~
V.a2 M/c ~6)
- MOAP,lI6 yCTON9H08 ~rf~
B rL M� (9) I AIIII noxass~aer xa-
~ G~o I 1 g (IO) paKrepxoe nouHaTHe nop�
~ V�`5) .Ha r9ropax(it) mttx sallxek onopm
31 ~ - GII@'1,T5 g ~14~
GIII~5,5 g
_ v~=a,a ~5) a
ia4s'~~ (i2)
e= So M~~ ~6~ a~~tm~~ MK
Siv=7S MM
MOJ~EJI6 H2YCTOFt4N88/8`
\ J
B u. C9) I'a~ea~xo~ceTp~ se 3a �
G,.a4,S g (10 ) ~xaccKpoaaax M~!~exr xa-
8 I G, =4 g(1 caxxa onopaMx rpyxra
~(g) Ha ortopax acnel[rraxe xapy:uexxx
GI=3,6g ranbaasxqecxo~ CSA3N
GII~3,6g e~e~xuy 07RE'JlbHbiMH yac-
4~ Jjf Q ~III �II,6 Q~ ?AMN blOAenH -
Gr ~ =~i,s g (t2) (15 )
vr=3,26 rt/e (S) bi=75 ecx
Vs=2,2 ht/c (6~ dII=77 MK " -
6-~50 81I1�28 MM
~ MOACJib YC70}tqNB2 (?1 dIVa
~ ?
B~� M� ( 9) IIYII - ccrepesopor�
!P G~~S B (10) cxrHana xapaKTepkcrH� _
111 f V~~S) G~=4g ~11) xoH AarYaxa r
_y Ha onopax
Gc=Sg ~16~
42 Il Gii = 3,6 g
Giu~,1,7 g
V~=3~26 ~r/c (g) G~v=3,6g ~12~
V�~2,0 u/c (Q) bt=83 uw
0=-5� bic=50 ~tH
I MoAenb YCTOIiqNH2 ~rf~ bIV-`~ MM
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Key:
l. Test number
2. Configuration and test conditions
3. Data from accelerometers and posi.tion transducers
4. Comments
5. Vhor ~m/s)
6. Vvert ~m/s)
7. Model is stable
8. Model is unstable
5'. At the center of mass
10. Horizontal acceleration (g's)
11. Vertical acceleration (g's)
12. On the landing-gear legs: Accelerations, G, in g's, and compres-
sions, d, in mm ~ .
13. The landing-gear legs are equipped with honeycombs having an
energy capacity about twice as large as those in subsequent
tests. Accelerometers TTT and IV show oscillations due to
Coulomb friction
14. Position transducer IT indi.cates the characteristic lifting of
the piston in the trailing landing-gear leg
15. The galvanometers did not register the instant of ground contact
by the landing-gear legs due to interruptions in t::e galvanometer
leads between different parts of the model
16. The accelerometer II signal was "inverted" due to the transducer
' characteristic
similar variant B model on a surface with a slope of -5�, but the order of
magnitude of these time spans remains the same even when the slope of the
surface is changed by + 15�.
The results of the tests have shown that the largest load factors occur for
- landing-gear legs that lead when landing downward along the slope; they are
two- to threefold greater than those for trailing legs. For the case of
~a,~u~ng with a single leg leading (configurations 1-2-1 und 2-1), this dif-
ference becomes even greater. However, let u~ note that for a four-legged
landing gear, this effect is less pronounced than in a three-legged landing
gear, due to the larger number of legs. In other words, for a four-legged
vehicle, the choice of landing configurations is less critical. For a vehi-
cle with a three-legged landing gear, and a 2-1 landing configuration, the
condition a= 0� (one leg leading downward along the slope), which other-
wise has advantages from the point of view of stability in terms of toppling
when compared with a 1-2 configuration, is more dangerous in terms of the ~
larger load factors on the leading leg.
-i Results of landing tests downward along the slope confir- the iarge compres-
sion of the energy absorbers in the leading landing-gear ~egs. During such
compress,ion, the landing-gear legs reached the maximurn values of deformation
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possible with the model, namely, 0.083 to 0.085 m(which is equivalen:: to
about 0.5 m for the full-scale lunar vehicle). This represents a compres-
sion of the honeycomb blocks of moze than 70 perce*~t.
Based on the material presented, in conclusion we will give one of the pos-
sible development scenarios (~~11-Union State Standard 2.103-68) for a land-
ing gear to be used with a future space vehicle for soft landing on a planet,
its satellites, asteroids, or other celestial bodies.
1. During the phase for developing the technical proposal (All-Union State
Standard 2.118-73):
a. Collection of data on the profile and properties of the ground at the
proposed landing site;
b. Pr213minary choice of design and dimensions for the landing gear as a
function of mission, layout, and mass of the space landing vehicle and con-
trol system options;
c. Performance of theoretical studies ~n the vehicle landing dynamics for
the selected configurations of its landing gear (for limiting values of
landing conditions).
2. During the phase of developing the preli:minary design (All-Union State
Standard 2.119-73):
a. Collection of more-detailed information on the surface layer of the
celestial body using hard landing and fly-by sounders, artificial satellites
of the planet, and so forth;
b. Performance of a detailed theoretical analysis of the landing dynamics
for the most likely landing-gear design for the space landing vehicle;
c. Design, manufacture, and test of dynamically similar modeis of the space
la.iding vehicle with the most likely landing-gear design.
3. During the phase of engineering development (All-Union State Standard
2.120-73):
a. Design, manufacture, and dynamic testing of full-scale test-stand mock-
ups of space landing vehicles (both simplified and maneuvering);
b. Structural testing of the full-scale landi.ng gear at design load factors
associated with landing;
c. Design, manufacture, and development of full-scale flying mockups of the
space landing vehicle; part of the development of a piloted vehicle consists
of training the cosmonauts to perfo~s soft landings in the flying mockup and
_ in the definition of the program of activities to be performed on the planet
(crew training in a mockup is also possible in a test stand with sufficient
room for maneuvering);
~ d. ~'jnal terrestrial work to assess the effectiveness of the vehicle's
landing gear duririg emergency (unstable) landing situations by using full-
scale test-stand mockups.
The landing gear, having successfully completed the cycle of terrestrial
development, is now ready for the ultizaate test, flight testing as part of
the space vehicle.
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, - .
Vi,N/c r-,,J r�~ !'�v,s
4 -1
opbo ~ I
/
~ _
' i. i i
itopua~
� -~-ISwN
/ ~--f ^10nN
1 -�-~-LS~H
( / 3nMtpunenT: i
I , / I o-ytmOQ~uEo
~ . 1 �-Mryrnaivua0
o i�~~'0,5-11
1 1 J 4 V~, n/r.
C2)
Figure 10.1. Stability Zones During Landing of the Variant B Space Landing
Vehicle (downward along the slope, a= 0�; 6=-5�; N= 4;
configuration 2-2)
Key:
Vvert ~m/s)
2. Vhor ~m/s) -
Theoretical results:
for ~ = 15 kN
- - - for ~ = 20 kN
- - - for ~ = 25 kN
Experimental results:
o stable
� unstable
fexperim. - 0.6 to 1.2
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,
- (i)
3 )
3
f-,fO~N ~�13aN j�10~H E�ISeN
( 3) ( 3)
.
4
, ~ /
/
' ~ ~�,oRH (3)
J ~
/ ~ le~ouA ,
- ~ ---f-0,4 -1�I
I , , _.._!-ir -�-i�r,1
1 ~ J.Ktn[puntem~
/ o -yt~cuvuao
�-Mrytmuuva0o
~i+cn'0,6+J,1
f 1 � J 4 v,, n/c
~2~
Figure 10.2. Stability Zones During Landing of the Variant B Space Landing
Vehicle (downward along the slope, a= 0�; 8=-10�; N= 4;
configuration 2-2)
Key :
l. Vvert ~m/s)
2 . Vhor ~m~S ~
3. ~ (kN)
Theoretical results:
- for f = 0. 9
for f = 1.0
for f = 1.1
- � - � for f = 1. 3
Experimental results :
o s table
~ unstable
fexperim. � 0.6 to 1.2
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(i)
J �N~~ ~3) 3)
E'.10.+/ f'1D~N g'/S4X ~'IOeH
�
~ ! / ~
~
J ~
I itcrufr:C- �-r'J,9 ~
/ F'f -�-f';1
j ntJ. rram.
o-y:arJU~~,C f :dtM4
~ �-rai.rcuv:?o ,
.]-y:i~,uvuEa '
I ! ?-~arc^cw~u3c J r~ena 1-z-1
' I i�:,~''J,5+1,1 I
i 1 3 4 V~, n/,;
~Z~
Figure 10.3. Stability Zones During Landing of the Variant B Space Landing
' Vehicle (downward along the slope, a= 0�; 0=-15�; N= 4;
configuration 202)
Key:
~vert ~m/s)
h (m/s)
3. ~ g~CN~
Theoretical results:
forf=0.9
for f = 1.0
~ for f = 1.1
Experimental results:
o stable ror the 2-2 conf.iguration
� unstabl~.
a stable
� unstable for the 1-2-1 configuration
fexperim. - 0.6 to I.~
183
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' (5)
~f 2 �1 4 Vr,f?/c
ye~~~l p y3 ~(3)~�~~~ (4)
~ J~ ~s -
-d ~2~ / ,e ~2 3 4~
d' ~2 3 4 so
-J6
Figure 10.4. Zones of Landing Stability for the Variant B Space Landing
Vehicle (plotted in the coordinates Vhor and 6, with the
r.umerals 1, 2, 3, 4, and 5 next to the data points indicating
jverticalJ velocities of 1, 2, 3, 4, and 5 m/s, respectively.
1.5 x 104 N; f= 1.0; N= 4; configuration 2-2)
Key :
1. Vvert - 1 m/s
2' Uvert - 2 m/s
3. Vvert - 3 m/s
4. Vvert � 4 m/s
5. Vhor ~m/s)
Theoretical results
- - - Experimental results -
fexperim. - 0.6 to 1.2
184
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~1i~,~Ai.~~,~~~ ~ ~ M S
;i
+ $s.?; .
~ow.... ~ ~ . . ;~~~~~~t...
i ! I~
4 ~j, w
~ ~ c:r . Y ~ . " n
_ ~ ~rr. ~ . ~
~ 4 `
~r~..-.:,., '~:~7 ~ .
~ .
-
. , ~~"iik' . .
oi
V
~ ' tk r
J~~~ nx'A;~..
. aS+vL~�' ~~n . -~f
~C
+ ~ y.+,. i . . ~ . `~~H~ .
u VC.}~;� i � �
~
x ~ 'r.: ~ ' 3'r�i~?
, qsit _ .
I ~ . . _ _ _.:`aa
...~y.s~ ~viVw~.
: ' s~~~., ; ~ r ~ L`:;. .
#~Ci.!%K ~ : .c~. , - ~ ` , . '
i~_ _ r t~:.:. e~ ~ . ,
r ~ . . . f
Fi.gure 10.5. Stable Landing of the DPM-B Model (downward along the slope;
6= 15�; velocities Vyert - 2�~ m/s and Vhor - 3.7 m/s; con-
figuration 1-2-1; rigid ground; frame interval ~ 0.05 s)
185
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(i) .
_ ?Q,
8
- eo ~VO ~`Op
B= f0� B=-S�
s
e=-~ ~
~
4
- ' \ B-0� �
~ y,.5
0
. ~~p ;
Z
~ +1p�
8=+15�
0 1 Z d 4 V~, M/C
(2 )
1 ~
Figure 10.6. Stability Zones During Landing of the Variant A Space Landing
Vehicle = 3 x 104 N; f= 0.8; N= 4)
Key:
1. Vvert ~m/s)
2. Vhor C~/s)
Configuratiion 2-2:
Theoretical results
- - - Experimental results
Can�iguration 1-2-1:
- � Theoretical results
186
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~e~ y/~ .
8 8=-10� ,i
~ B= 1
6
Teop~ra:
Tf-10KH
4 ---Tf-fOKN
-Tr-O
� ~ 3NCnepuncMm:
p o-Tf-?O~rN; 9--ZS
~-T -177H 1g~3~
r ycmorivcr4o .
O=-JO ~ ~ aT�-36rrHJ
�-Tr-10KH; B� J7'xtycmodyrrEiv
0 1 1 3 4 n/C
Figure 10.7. Stability Zones During Landing of the Variant A Space Landing
Vehicle (using "clamping" engines; 3 x 104 N; a= 0�;
N= 4; f= 0.8; configuration 2-2)
Key:
1. Vvert ~m/s)
2. Vhor ~m~s~
Theoretical results:
for T~ = 20 kN
- - - for TE.= 10 kN
� for TE = 0
Experimental results:
Stable landings:
o for TE = 20 kN and A=-25�
~ for TE = 27 kN and 6=-30�
? for TE = 36 kN and A=-30�
Unstable landing
� for TE = 20 kN and 6=-37�
187
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(5)
Vs-Sn/c
� g� 6 7
y6=8n/c -mevpr~A
~g~ 3NCO~QQHCN?~
8 ~7~
~ f 7 3~~Z 4r
~8~ ~ V~~6 V~iy/c
~a'~'~~~ 4 (9)
-8 ~g)
_16 Vc'S Vc' y~.1rr/c~i~ .
~4) C3) Yr~ )
2
Figure 10.8. Zones of Landing Stability for the Variant C Space Landing
Vehicle (plotted in the coordinates ~hor and 6; for an expla-
nation of labeling of points see Figure 10.4; 42.5 x
104 N; N= 4; configur.ation 2-2)
Key:
1. Vvert - 1 m/s
2� Vvert - 2 m/s
3. Vvert - 3 m/s
4. Vvert - 4 m/s
5. Vvert - 5 m/s
6. Vvert - 6 m/s
7. Vvert = 7 m/s
8. Vvert = 8 m/s
9. Vhor ~m/s) ~
Theoretical results
-o- Experimental results
188
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~3)
~1~ ~e M~~i~i=y 0,5 0,55Q60 ~
6 ~
/ / ~iii O,d6
4 / I~~ ~4~
\ ~ \
~ \ \
2
~ \
' 0,J7 0,4
~ " 1 2 v~.M/~
~2)
Figure 10.9. Lines of Equal Landing Stability for the Variant A Space Land-
ing Vehicle (a = 0�; 3 x 104 N; 6=-15�; f= 0.8; N= 4;
configuration 1-2-1)
Key :
1. Vvert ~m/s)
2. Vhor ~m~S)
3. 0
4. Ocritical
189
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' ~ ~ _ _~1 - i =r
~T

~ t~ ~ . ~ . ~~~~l~i~~~ ~'i. ~ . ~ ~i ~ ~
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Vt, n/c
B
N�J, c.rena /-Z (3)
/ N-4, r,~ena 1-Z (4) N-J, c~ena 1-1 ( 3)
/
6 N� 4, c,~ena 1-1 (4 )
N-S,c~ena; 21 (5)
4 N�6,c,~ena17Z (6)
~
~ \
Z
\ ~
0
1 1 3 4 V~, a/c
(2)
Figure 10.10. Aependence of the Size of a Space Landing Vehicle's (Variant
A) Stability Zone on the Number of Landing-Gear Legs
3 x104N; 0.8; a=0�)
Key:
e = o�
e = -~o�
= 1. Vvert ~m/s)
2. Vhor Cm/s)
3. N= 3, conf iguration 1-2
4. N= 4, conf iguration 2--2
5. N= 5, conf iguration 1-2-2
6. N= 6, configuration 2-2-2
NPIS B--JO'
3
1,5
? Bf~
1,5
/ ~
~ J 4 S N
9
Figure 10.11. Optimizing of a Space Landing Vehicle's Landing Gear in
Terms of Mass (Variant A; 3 x 104 N; f= 0.8; a= 0�)
i9o
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r~ui: ~~~~i~ic~nr. UtiL: U~II.i
)
VQ, M/c
8
~ R-R/N�I,f9
6
R�;Z7 '
~ 4
~R-f,4
/ /
T
/
~ f 2 3 4. V~, y/~
(2)
Figure 10.12. Dependence of tl~e Size of a Space Landing Vehicle's (Variant `
A) Stability Zone on the Relative Magnitude of the Landing--
Gear Radius (a = 0�; 3 x 104 N; N= 4; f= 0.8; configu-
ration 2-2)
Key :
1. Vvert ~m/s)
2. Vhor ~m/s)
e = -io�
6 = 0�
~ 191
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~
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l~i)h Ol~I~ Lc: f~1L IrSI~. (~NI.Y
X
2,4 -
. X' ~~6
~ 1) H~~ Q~8 i ~ `
0 ~ ; ~
2 Y ~ ' i ~
~1~ % _p ' -
~ ! i I
_4 ' i ~ i
� ~ I '
I ~P !
(2) _08 ~ ~ .
1,6 ~
.~,o x , '
~3) M p~p ~ ~ - ~
i I I ~
(3,0 i ~ ,
~3) M~ { 1,6 ' ~
l 2,2 ~ ,
D ; i i ~
(4~ ~P, j-0,4 i
Pad ~-0,8 i ~ i ~
a~ ~'2 ~ ~ ; I
_ ~4) pad Q~ ~ I ~ ;
. , ~ ~
(4, . Pad1 ~ az ~ ' I -
0 : '
40,0 ' ~ ; ~
r ~
~5~ ~ T O,A/6 0,41 ,1 ,7 GO 1,91~ ;dd
rv^ 6
0,? 0,4 0,6 0.8 1,0 ;1 f,4 ;6 t,c
I 11 Ill � IY �
Y
YI PU
. Monanm BKnbveNUn cmddunu~upyroufr~x d6uzameAeri (7 )
Figure 10.13. Behavior of Dynamics Parameters During an Unstable Landing of
a Variant A Space Landing Vehicle (downward along the slope
with 9 = 20�)
Phases of the Landing Process:
I. Crushing of landing-gear leg no. 1
II. Rotation with respect to landing-gear leg no. 1
III. Crushing of landing-gear leg no. 2
IV. Rotation with respect to landing-gear leg no. 2
V. Operating period of'stabilizing engines
VI. Crushing of landing-gear leg no. 2 when the impulse from the
stabilizing engines is accounted for
192
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VII. Rotation with respect to landing-gear leg no. 2 when the impulse
from the stabilizing engines is accounted for
~ Key:
1. R and Y (m/s)
2. ~ (radians/s)
3. X and Y (m)
4. -al, and aZ (radians)
5. T (kJ)
6. t (s)
7. Instant of turning on stabilizing engines
. - L93
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i~'Uk OFi'IC [,11. I~SI: OtiI.Y
1,6 _
- f,2 ' -
(1) 0,8 - -
0~ ~ ~ 1 X I _ -
(1) ~~{-04 I i ' Y 1 I i
~ , i ~
0
(2) pa%~_O,B I 1 ' ~ ~ .
X, ~ JD ~ j ~ ' i I
(�3) M l 1,0 '
i
(3) Y.~ ,~o , ~ - . v ~ , .
M 2,6
?,2
( 0
~4) -0,2 ,
Pad -0,4 ~
~ ~
-0,6 ~ ~ ' 1
~a~ -a,, ~~4 ~ a, ,
pad ~6 ~ ~
~ ,O
i ; ' I
(4~ pad{ 0,3 a~----z 1
~
" 10,0~ ~ ~ - I ,
9,0 ~ ' ~ , ~
(5) ~Qm 40 ~ ! ' ~
z,a ~ 'T'
~ i , ;
~0 _0,7 ~74 G,6 18 ~0 ,2 f,4 f,6 f,Bt,c ~6)
! Q ~`r`_'- ~ .
MvneNm aKn~oveHUn
~ cma~unrc~upyrou{u~ d6utameneu (7 ) _
Fi.gure 10.14. Behavior of Dynamics Parameters During a Stable Landing of a -
Variant A Space Landing Vehicle (downward along the slope
with 6 = 20�)
Phases of the Landing Process:
I. Crushing of Ianding-gear leg no. 1
~ II. Rotation with respect to landing-gear leg no. 1
III. Crushing of landing-gear leg no. 2
IV. Rotation with respect to landing-gear leg no. 2
194
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~~~,E; c~i~ ~c ~ ni, ~~:;I? ~)NI,Y
x~y :
- 1. X and Y (m/s)
2. � (radians/s)
3. X and Y (m)
4. a , and a2 (radians)
5. T (kJ}
6. t (s)
7. Instant of turning on stabilizing engines
~ (3) (4)
Bapr~aNm A Bapad'Hirf B
6�0� B-5�
B-5� Ve-6N~c_...,_.
B->0�,~_.
~,NM
~ODa \ ~2~ ~
~ ,6 ~
\ b~
600 ~ y~
~ ~ /
. 6C~7 \ /
~ ~ %
yaa ` ~
~ 3
/
i
100 ~ \ ~ ~N~c ~2)
~ V~r/
/
2 J ~ N 1
Figure 10.15. Typical Deformations of Landing-gear Legs for the Variants
. A and C Space Landing Vehicles (landing upward along the
slope, a= 180�; Vhor = 1 m/s; configuration 1-2-1)
Key:
1, d (mm)
2. Vvert C~/s)
3. Variant A
4. Variant C .
f-
195
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BIBLIOGRAPh'Y
l. Avduyevskiy, V. S.; Anfimov, N. A.; Antonov, B. r1.; and others.
"Osnovy teorii poleta kosmicheskikh apparatov" [Foundations of the
Theory of Spacecraft Flight], edited by G. S. Narimanov and M. K.
Tikhonravov , Moscow, Mashinostroyeniye, 1972.
~ 2. Avduyevskiy, V. S.; Marov, M. Ya.; Rozhdestvenskiy, M. K.; and others.
"The Venera-9 and Venera-10 Automated Stations: The Operation of the
Descent Mocisle and the Measurement of Atmospheric Parameters,"
KOSMTCHESiC~�~ ISSLEDOVANIYA, Vol 14, No 5, pp 655-666, 1976.
3. Anon. "Avtomaticheskiye planetnyye stantsii" [Automated Planetary
Stations], edited by Yu. K. Khodarev , Moscow, Nauka, 1973.
4. Anon. "Algor~thms and Programs for the Solution of Engineering Opti- -
mization Problems on Digital Computers, Scientific-Engineering Report,"
Moscow, Gosfond for Algorithms and Programs, 1970.
5. Alekseyev, K. B.; Bebenin, G. G.; and Yaroshevskiy, V. A.
- "Manevrirovaniye kosmicheskikh apparatov" [Maneuvering of Spacecraft],
Moscow, Mashinostroyeniye, 1970.
� 6. Andreyevskiy, V. V. "Dinamika spuska. kosmicheskikh appar.atov na
zemlyu" [Dynamics of Spacecraft Descent to Earth], Moscow,
Mashinostroyeniye, 1970.
7. Arkhangel'skiy, B. A. "Plasticheskiye massy" [Plastics--A Reference],
Leningrad, Sudpromgiz, 1961.
~S. Bazhenovs V. I.; Kova1', A. D.; and Straut, E. K. "From ~~fodel to
Lunokhod," AVIATSIYA I I~OSMONAVTIKA, No 8, pp 42-46, 1966.
9. Bazhenov, V. "Lunokhod Ground School," VESTNIK APN PO S0~'ETSKOMU
SOYUZU, No 274, pp 1-S, 23 Nov 1970.
10. Bazhenov, V., and Sakharov, Yu. "Earth Rehearsal for Venera," News-
paper LENINSKOYE ZNAMYA, No 277, 4 p, 26 Nov 1975.
11. Bazhenov, V. I.; Goncharov, A. I.; and Osin, M. I. "Nek~toryye
problemy vozvrashcheniya cheloveka iz kosmosa i ikh razvitiye ot
trudov K. E. Tsiolkovskogo do nashykh dney" [Several Problems in
Returning Maii from Space and Their Development, Starting with the
Works of K. E. Tsiolkovskiy to the Present~, Works of the Tenth
Tsiolkovskiy Readings, Moscow, IiYeT [Institute of History of Natural
Sciences and Technology] of the USSR Academy of Sciences, pp 28-40,
1977.
~ 196
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~
FOR OFFICIAL USE ONLY
12. Barer, A. S.; Sorokina, Ye. I.; and Murakhovskiy, K. I. "On the Ques-
tion of Man`s Tolerance for Acceleration After a Prolonged Stay Under `
Conditions Simulating Weightlessness," KOSMICHESKAYA BIOLOGIYA I
MEDITSINA, Vol 6, No 3, pp 49--53, 1972.
13. Bilik, Sh. M. "Pary treniya metall-plastmassa v mashinakh i
mekhanisma.kh" [Friction Between Metal and Plastic in Machines and
Mechanisms], Moscow, Mashinostroyeniye, 1966.
14. Blandov, P. T. "Amortizatsiya i upravleniye vsletno-posadochnyk:~
ustroystv samoletov" [Shock Absorption and Control of Aircraft Landing
Gears], MAI, 1962.
15. Bogdanov, V. A.; Gurfinkel', V. S.; and Panfilov, V. Ye. "Test Stand
for Modeling Reduced Gravity at Surface Conditions," Invention certifi-
cate no 215404, IZOBRETATEL' I RATSIONALIZATOR, No 12, p 12, 1970.
I.;. Borisov, M. "Na k.osmicheskoy verfi" [At the Space Shipyard], Moscow,
Mashinostroyeniye, 1976.
17. Kemurdzhian, A. L.; Gromov, V. V.; Cherkasov, I. I.; and others.
"Avtomaticheskiye stantsii dlya izucheniya poverkhnostnogo pokrova
luny" [Automated Stations for the Study of the Surface Layer of the
Moon], Moscow, Mashinostroyeniye, 1976. .
_ 18. Klimin, A. V. "At-mospheric Braking for Manned Spacecraft at the ~on-
clusion of Interplanetary Flight," KOSMICHESKIYE ISSLEDOVANIYA, Vol l.l,
- No 1, pp 31-37, 1973.
19. Krotikov, V. D., and Troitskiy, V. S. "Radio Emission and Nature of
the Moon," USPEKHT F?ZICHESKIKH NAUK, Vol 81, pp 51-56, 1963.
20. Lokh, U. Kh. T. "Dinamika i termodinamika spuska v atmosferG;ch planet"
[Dynamics and Thermodynamics of Descent Through Planetary Atmospheres),
' Moscow, Mir, 1966.
21. Anon. "Lunnyy grunt iz morya izobiliya" [Lunar Soil from the Sea of
Fertility], edited by A. P. Vinogradov, Moscow, Nauka, 1974.
22. Mak-Farland, R. K. "Hexagonal Honeycomb Structures Under the Action
of Supercritical Axia1 Loads," RAKETNAYA TEKFINIKA I KOSMONAVTIKA, No 6,
pp 125-131, 1963.
23. Mishin, V. P., and Osin, M. I. "Vvedeniye v mashinnoye proyek-
tiro~Taniye letatel'nykh apparatov" [Introduction to Machine Design of
_ Flight VehiclesJ, Moscow, Mashinostroyeniye, 1978.
24. Moroz, V. I. "Fizika planet" [Physics of Planets], Moscow, Nauka,
1967,
197
FOR OFFICIAL USE ONLY
-I
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25. Moroz, I. "Structure of the Martian Soil Based on Optical and
Infrared Observations," KOSMICHESKIYE ISSLEDOVANIYA, Vol 14, Issue 1,
pp 85-96, 1976.
� 26. Naumov, V. N.; Rozhdestvenskiy, Yu. L.; and Kharitonova, V. Ye. "Pro-
duction of Artificial Soil," Transactions of MVTU [N. E. Bauman Higher
Technical School in MoscowJ, No 231, Issue 1, Moscow, MVTU, pp 32-37,
1976.
J 27. Neyland, V. Ya., and Snigirec, Yu. I. "Thermal Regime of Flight Vehic- R_
les at En~ry Velocities into t:~e Earth's Atmosphere of Approximately
15 km/s," KOSMICHESKIYE ISSLEDOVANIYA, Vol 5, Issue 2, pp 205-211,
1967.
_ 28. Pellinets, V. S. "Ob otsenke pogreshnostey pri izmerenii parametrov
vibratsiy" [On Estimating Errors in the Measurement of Vibration Param- -
eters~, Leningrad, LDNTP [Leningrad House of Scientific and Technical
Propaganda], 1967.
29. Anon. "Pionery raketnoy tekhniki" [Pioneers of Rocket TechnologyJ,
szlected works, Moscow, Nauka, 1964.
30. Sveshnikov, A. A. "Osnovy teorii oshibok" [Foundation of the Theory of
Errors], Leningrad, LGU [A. A. Zhdanov State University in Leningrad], _
1972.
31. Sedov, L. T. "Metody podobiya i razmernosti v mekhanike" [Methods of
Similarity and Dimensional Analysis in MechanicsJ, Moscow, Nauka, 1972.
32. Strelkov, S. P. "Vvedeniye v teoriyu kolebaniy" [Introduction to the
Theory of Oscillations], Moscow, Nauka, 1964.
33. Teyfel', V. G. "Atmosfera planety yupiter" [Atmosphere of the Planet
Jupiter], Moscow, Nauka, 1969.
34. Timoshenko, S. P. "Kolebaniya v inzhenernom dele" [Oscillations in
Engineering], Moscow, Nauka, 1967.
,
35. Trofimenkov, Yu. G., and Vorobkov, L. N. "Polevyye metody
isslednvaniya stroitel'nykh svoystv gruntov" [Field Methods of Studying
Structural Properties of Soils], Moscow, Stroyizdat, 1974.
36. Anon. "Angular"Accelerometer, Insensitive to ~.inear Accelerations,"
U. S. Patent No 316498, kl, 73-516. [sic]
37. Khint, I. A. "Osnovy proizvodstva silikal'tsitnykh izdeliy' iFounda-
tions of Manufacturing Silicalcite Products], Moscow, Stroyizdat, 1962.
198 ~
FOR OFFICIAL USE ONLY
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- 38. Tsiolkovskiy, K_. E. "Za atmosferu" jBeyond the Atmosphere], VOKRUG
SVETA, No 1, pp 10-14, 1934.
_ 39. Tsytovich, N. A. "Mekhanika grunta" [Soil Mechanics], Moscow,
Gosstroyizdat, 1963.
~+0. Shkadov, L. M.; Bukhanova, R. S.; Illarionov, V. F.; and others.
"Mekhanika optimal'nogo prostrantsveanogo dvizheniya letatel'nykh -
apparatov v atmosfere" [Mechanics of Optimal Three-dimensional Motion -
of Flight Vehicl~s in the AtmospherP], Mosco~a, Mashinostroyeniye, 1972.
;I. Shternfel'd, A. A. "Vvedeniye v kosmonavtiky" [Introduction to Space
Flight;, "'~oscow, Nauka, 1974.
42. Shchegolov, A. V. "Konstruirovaniye prot~azhek" [Designing oF
Broaches], Moscow, Mashgiz, 1960.
43. E1'yasberg, P. `Ie. "Vvedeniye ~ teoriyu poleta iskusstvenykh sputnikov
zemli" [Introduction to the Theory of Flignt or Artificial Satellites
of the Earth], rloscow, Nauka, 1965.
44. Admire, J., and :~Iackey, A. "Dynamic Analysis of a Multilegged Lunar
Landing Vehicle To Determine Structural Laads During Toucndown," NASA
TN D-z582, 1965.
45. Bartolomew, C. S. "Risk Appraisal for Program Planning and Manage-
- ment," ANN 4LS OF R~LIABILITY AND MAINTAlNABILITY, Vol 6, 1967.
46. Bellman, D. R., P~tatranga, G. J. "D~sign and Operational Characteris-
tics of a Lunar Landing Research Vehicle (LLRV)," Edwards Flight
Research.Center, NASA TN D-;s023, 1965.
47. Black, R. J. '~Quadrupedal Landing Gear System for Spacecraft," J.
OF SPACECRAFT AND ROC., Vol 1, No 2, p 196, 1964. ~
48. Blanchard, Ulysse, J. "Model Investigation of Technique for Conduct-
ing Full-scale Landing-Impact Tests at Simnlated Lunar Gravity," NASA
TN D-2586, 19b5.
49. Blanchard, Ulysse J., "Lvaluation of a Full-scale Lunar-Gravity Simu-
lator by Comparison of Landing-Impact Tests of a Full-scale and a 1/6-
scale Model," NASA TN D-4474, 1968.
50. Blanchard, U. .J., and Stubbs, S. M. "Martian Gravity Simulator for
Fu~]--srale Viking Landing Tests." Space Simulation, NASA SP 298,
pp 629-646, 1972.
51. Congelon, W. M. "Investigation of Reflecting Heat Shield Materials
for Outer Planet Missions," AIAA Paper, N 74-702.
199
FOR OFFICIAL USE ONLY
APPROVED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000300040031-0

APPROVED FOR RELEASE: 2047102/08: CIA-RDP82-00850R000300040031-0
FOR OFFICIAL USE ONLY
52. Douglass, D. R., znd Austin, C. R. "Mission Risk Appraisal," ANNALS
OF RELIABILITY AND ~~AINTAINASILITY, Vol 6, 1967.
53. Garden, H. D.; Herr, R. W.; and Brooks, G. ~d. "Technique for the Simu-
lation of Lunar and Planetary Gravitational Fields Including Pilot
Model Studies," NASA TN D-2415, 1964.
54. Herr, R. W., and Wayne, L. H. "Dynamic Model Investigation of Touch-
down Stability of Lunar Landing Vehicles," NASA TN D-4215, 1967.
55. Hylderman, R.; Mueller, W.; and 1tilantus, M. "Landing Dynamics of the
Lunar ~xcursion Mndule," J. OF SPACECRAFT AND ROCKETS, Vol 3, No 10,
pp 1484-14fi9, 1966.
56. "Langley'~ Lunar Landings. Lnnar Landing Research Facility (LLRF),"
:~IISSILE� Aiv~ ROCKETS, Vo1 17, :'ao 2, p 8, 1965; FLIGHT INTERNATIONAL,
Vol 88, No 2940, p].ll, 1965.
57. Leonard, H. W.; Wal.ton, W. C.; and Herr, R. W. "Studies of Touchdown
Stability for Lunar Landing Vehicles," J. SPACECRAFT AND ROC., Vol 1,
No 5, p 552, 1964.
58. Mayo E. E., and others. "Hypersonic Reentry Vehicle," U. S. Patent
3301507.
59. McGehee, John R., and Stubbs, S. M. "Experimental Validation of a
Landing Dynamics Computer Program for Legged Spacecraft Landers," NASA
_ TN D-703, 1973.
60. Miller, R. E.; Hansen, S. D.; Kawaguchi, A. S.; Redhed, D. D.; and
Southall, U. W. "Cost Effectiveness of Integrated Analysis (Design
Systems/IPAD/)," AIAA Paper N 74-960.
61. Moss, I. N.; Anderson, E. S.; and Bold, C. W. "Viscous Shock Layer
Solutions for Jovian Entry," AIAA Paper N 75-671.
62. Nicolet, W. E.; Morse, H. L.; Vo~vodich, N. S. "Outer Planet Probe
Entry Thermal Protection, Part 1: Aerothermodynamic Environment,"
AIAA Paper N 74-700.
63. Nicolet, W. E.; and Mizines, S. A. "Outer Planet Probe Entry Thermal
Protection, Part 2: Heat Shielding Requirements," AIAA Paper N 74-701.
64. 0'Bryan, T. C.; and Hewes, D. W. "Operational Features of the Langley
Lunar Landing Research Facility," NASA TN D-3828, 1967.
65. Rau, T. R.; and Decker, J. P. "Optimal Design Integration System for .
Synthesis of Aerospace Vehicles," AIAA Paper N 74-72.
200
FOR OFFICIAL USE ONLY
APPROVED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000300040031-0

APPR~VED FOR RELEASE: 2007/02/08: CIA-RDP82-00850R000300040031-0
FOR OFFICIAL USE ONLY
66. Schuring, D. "Scale Model Testing of Land Vehicles in a Simulated
Low-gravity Field," SAE-660Z48, 1966.
67. Smith, A. M. "Risk Assessment in Complex Unattended Aerospace Systems,"
ANNALS OF RELIABILITY AND MAINTAINABILITY, Vol 6, 1967.
- 68. Stubbs, S. M. "Investigation of Technique for Conducting Landing
Impact Tests at Simulated Planetary Gravity," NASA TN D-6459, 1961.
69. Stubbs, S. M. "Experi.mental Investigation of the Landing Dynamics of
Three-legged Spacecraft Models," NASA TN D-7664, 1974.
7U. Troitsky, V. S. "A Four-meter Thick Porous Ma.terial Layer Covers the
Moon," SPACE WORLD, Vol D-6 42, pp 44-45, June, 1967.
71. Warner, R. W.; Sorenson, R. M.; and Kaskey, A. J. "An Inves.tigation of
a Deforming Energy Absorption System for Space Vehicle Landings," NASA
TN D-3061, 1966.
COPYRIGHT: Izdatel'stvo Mashinostroyeniye, Moscow, 1978
691+8
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END
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