structures, mechanisms, launch vehicle selection
TRANSCRIPT
Structures, Mechanisms, Launch Vehicle Selection
Aerospace and Ocean Engineering Department
Virginia Tech
Blacksburg, VA
Team Members:
Ann W. Bergquist, Jessica M. Jensen, Brian J. Santiestevan,
Andrew T. Vaughan, Christopher P. Vlastelica, Christopher D. Weaver
November 16, 2001
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Table of Contents
Table of Contents................................................................................................................ ii List of Figures .................................................................................................................... iv List of Tables ...................................................................................................................... v List of Abbreviations ......................................................................................................... vi List of Abbreviations ......................................................................................................... vi List of Symbols ................................................................................................................. vii
1 Introduction................................................................................................................. 1 1.1 Structures and Mechanisms .................................................................................... 1 1.1.1 Structures ............................................................................................................ 1 1.1.2 Mechanisms ........................................................................................................ 2 1.2 Launch Vehicle ....................................................................................................... 2 1.3 Summary and Overview ......................................................................................... 3
2 Subsystem Modeling................................................................................................... 5 2.1 Structures ................................................................................................................ 5 2.1.1 Modeling ............................................................................................................. 5 2.1.1.1 Statics.................................................................................................................. 6 2.1.1.2 Dynamics ............................................................................................................ 7 2.1.1.3 Mechanics of Materials....................................................................................... 7 2.1.1.4 Flexible-Body Dynamics .................................................................................. 11 2.1.2 Interactions........................................................................................................ 12 2.2 Mechanisms .......................................................................................................... 14 2.2.1 Modeling ........................................................................................................... 14 2.2.2 Interactions........................................................................................................ 16 2.3 Launch Vehicle Selection ..................................................................................... 18 2.3.1 Modeling ........................................................................................................... 18 2.3.2 Interactions........................................................................................................ 21 2.4 Summary............................................................................................................... 22
3 Subsystem Examples ................................................................................................ 23 3.1 Structures .............................................................................................................. 23 3.1.1 Materials ........................................................................................................... 23 3.1.1.1 Metals................................................................................................................ 23 3.1.1.2 Composites........................................................................................................ 24 3.1.1.3 Shape-Memory Alloys ...................................................................................... 27 3.1.2 Structures .......................................................................................................... 29 3.1.2.1 Solar arrays ....................................................................................................... 29 3.1.2.2 Tethers............................................................................................................... 30
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3.1.2.3 Bus .................................................................................................................... 33 3.2 Mechanisms .......................................................................................................... 35 3.2.1 Low-Cyclic Mechanisms .................................................................................. 36 3.2.1.1 Solar Array/Antenna Retention/Deployment Mechanisms .............................. 36 3.2.1.2 Payload/Launch Vehicle Separation Mechanism ............................................. 38 3.2.1.3 Payload Retention Device................................................................................. 38 3.2.2 High-Cyclic Mechanisms.................................................................................. 39 3.2.2.1 Antenna Pointing and Tracking ........................................................................ 39 3.2.2.2 Solar Array Drive Mechanisms ........................................................................ 40 3.2.2.3 Attitude Control Reaction Wheels .................................................................... 42 3.3 Launch Vehicle Examples .................................................................................... 43 3.3.1 Launch Vehicles................................................................................................ 43 3.3.1.1 Ariane-4 ............................................................................................................ 44 3.3.1.2 Titan IV............................................................................................................. 45 3.3.2 Payloads ............................................................................................................ 47 3.3.2.1 Shuttle Missions................................................................................................ 47 3.3.2.2 Titan IV Missions ............................................................................................. 48 3.3.3 Summary........................................................................................................... 48
4 Summary and Conclusions ....................................................................................... 50 4.1 Structures .............................................................................................................. 50 4.2 Mechanisms .......................................................................................................... 52 4.3 Launch Vehicles.................................................................................................... 52 4.4 Recommendations................................................................................................. 53
References......................................................................................................................... 56
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List of Figures
Figure 1: Diagram showing normal and shear loads on a rigid bar .................................... 9 Figure 2: Operating rate profile39...................................................................................... 15 Figure 3: Torque-speed curve39 ........................................................................................ 15 Figure 4: Stall torque plot39............................................................................................... 16 Figure 5: Inflatable framing structure and panel 38........................................................... 30 Figure 6: Trend of tether materials through time9 ........................................................... 31 Figure 7: Sectional drawings of multiline Hoytether™ 11 ................................................ 32 Figure 8: Illustration of sandwich panel and isogrid structures23 ..................................... 35 Figure 9: Solar array retention/deployment mechanism23 ................................................ 38 Figure 10: Retention-latch actuator22................................................................................ 39 Figure 11: Antenna pointing mechanism24 ....................................................................... 40 Figure 12: Solar array drive mechanism25 ........................................................................ 41 Figure 13: Alcatel solar array deployment mechanism26................................................. 41 Figure 14: Reaction wheels35 ........................................................................................... 42 Figure 15: Configurations of the Ariane-4 launch vehicle8 .............................................. 44 Figure 16: Schematic of Titan IV launch vehicle18 ......................................................... 46
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List of Tables
Table 1: Subsystem interaction matrix ............................................................................... 5 Table 2: Properties of commonly used spacecraft structure materials4 .............................. 8 Table 3: Launch vehicle data 12,19,33.................................................................................. 20 Table 4: Characteristics of typical Polymer-matrix composites ....................................... 26 Table 5: Prices of typical composite materials used in space structures2 ......................... 27 Table 6: Alloys exhibiting shape memory effects5 ........................................................... 28
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List of Abbreviations
ADCS Attitude determination and control system C&DH Command and data handling cg Center of gravity DOF Degrees of freedom FBD Free-body diagram FS Factor of safety HEO High-Earth orbit MOE Measure of effectiveness MOI Moment of inertia MS Margin of safety PMC Polymer-matrix composite RF Radio frequency
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List of Symbols
A Cross-sectional area a Acceleration BH Earth’s magnetic field F Force G Modulus of rigidity I Moment of Inertia L Original length m Mass M Moment P Normal (axial) force Pw Power R Resistance t Slew time T Torque capacity V Shear force v Velocity γ Shear strain δ Change in length ε Normal strain θ Slew angle Σ Summation σ Normal stress σc Conductivity τ Shear stress ΦOC Voltage (open circuit)
1
1 Introduction
Structures and mechanisms are integral parts of any spacecraft, and the launch
vehicle is required to place the spacecraft into orbit. The structures and mechanisms
subsystem serves as the physical backbone supporting all other subsystems. Although
other subsystems are not directly affected by the launch vehicle, vital attributes of the
spacecraft are constrained by the launch vehicle selection. This chapter briefly describes
these three subsystems and illustrates how each is vital to the overall design.
1.1 Structures and Mechanisms
1.1.1 Structures
The structure of a spacecraft is one of its most vital subsystems; it serves as the
housing for the mission payload and for all of the spacecraft’s control systems. Important
parts of almost every spacecraft structure include the spacecraft bus, which holds all
electronic and power components, a solar array structure, a propulsion module, and a
cover and support for the communications equipment. Each of these pieces of the
structure must bear the loads and vibrations imposed during launch and orbital
maneuvers. Additionally, the structure must be able to tolerate the space environment for
many years, depending upon the mission duration. Skin panels, trusses, pressure vessels,
brackets, and equipment boxes are all examples of typical aerospace structures.39
Material selection plays a vital role in the total cost, weight, and lifetime of the
spacecraft. Some important considerations while selecting a material are thermal
conductivity, strength, stiffness, ductility, and corrosion resistance. Each of these can be
2
maximized or minimized as a measure of effectiveness (MOE) to help select an optimal
structure for a given mission. For example, a designer may choose to use titanium for its
high strength and low coefficient of thermal expansion, but he or she would be sacrificing
cost economy and machinability.39
1.1.2 Mechanisms
Sarafin23 defines a mechanism as “an assembly that moves to function.”
Mechanisms are typically used on spacecraft for deployment or retraction of specific
instruments and are operated by a control system. Space mechanisms must be more
reliable than ordinary mechanisms, as mechanical repair in space is difficult and
impractical. Some mechanical design considerations include high launch vibrations, the
micro-gravity environment, and power restrictions.23
Aerospace mechanisms can be divided into two categories: high-cyclic
mechanisms and low-cyclic mechanisms. High-cyclic mechanisms are mechanisms
requiring frequent operation including antenna gimbals, boom extensions, and
momentum wheels. These mechanisms usually fail due to excessive component wear, so
they are designed to withstand many loading cycles. Low-cyclic mechanisms include
launch vehicle separation components, antenna and solar array deployment mechanisms,
and other devices only operated once. Mechanisms such as these are designed to
withstand a one-time maximum load. 39
1.2 Launch Vehicle
Much of the design of a spacecraft will be constrained by the size and weight
restrictions particular to the launch system. In general, there are five steps to selecting a
3
launch system for a particular mission. The first step involves defining the requirements
and constraints for the mission. At this point, issues such as mission timeline, funding
constraints, and spacecraft dimensions are addressed. The next step involves identifying
and analyzing acceptable configurations for the launch system. During this step the
reliability, performance, and lifting capacity are considered in addition to other factors
including acceleration imparted to the satellite and vehicle vibration. The third step is the
selection of the potential launch system. A potential launch system will be evaluated
using the following criteria: lifting capability, cost, performance margin available,
reliability, and schedule versus vehicle availability. Next, the environments created by
the launch system are determined, as well as the spacecraft design envelope. This step is
required to determine how the launch system may negatively affect the spacecraft. The
fifth, and final, step for selecting a launch system is to iterate the previous four steps in an
effort to meet constraints on performance, cost, risk, and schedule. 39
1.3 Summary and Overview
This report details many of the aspects of the design and selection of spacecraft
structures, mechanisms, and launch vehicle. It is divided into four chapters that are
arranged as follows. Chapter 1 introduces some of the structures, mechanisms, and
launch vehicles that are important to the overall project. Chapter 2 describes the
subsystem modeling, including how each subsystem was modeled, what equations,
charts, or graphs were used, and what other information is required for an accurate
model. This chapter also gives a thorough description of the interaction between the
various subsystems. Chapter 3 presents subsystem examples, giving details of what
options are already available for use, or that will become available in the near future.
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Examples of similar selection processes are also presented from previous spacecraft
designs. Chapter 4 summarizes the report, presenting the conclusions drawn by the group
and recommendations for future research.
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2 Subsystem Modeling
Structures, mechanisms, and launch vehicles all play integral roles in the success
of any space mission. Informed decisions may be made for the final design by accurately
modeling each of these systems. The subsystem modeling chapter presents many of the
charts, tables, and equations that are used to make calculations that are relevant to these
systems. This chapter also describes how each of these systems interact with other
systems onboard the spacecraft. These interactions are illustrated in Table 1 and are
described in subsequent sections.
Table 1: Subsystem interaction matrix
Astro
dynam
ics
Guidan
ce &
Nav
igat
ion
Propulsi
on
ADCS
Comm
unicatio
ns
Comm
and &
Dat
a Han
dling
Power
Therm
al
Environm
ent
Progra
m M
anag
emen
t
CostMiss
ion O
p & G
round S
yste
ms
Econom
ics, P
olitics
, Leg
al
Mechan
isms
Struct
ures
Launch
Veh
icle
Launch Vehicle 2 0 0 0 0 0 0 1 1 1 2 1 1 2 2 -
Structures 1 1 1 2 1 1 2 2 2 0 2 1 0 1 - -
Mechanisms 0 0 2 2 2 1 2 1 2 0 2 1 0 - - -
2.1 Structures
2.1.1 Modeling
Basic principles of engineering mechanics must be understood to begin modeling
a spacecraft structure. Some important concepts include principles of statics, dynamics,
mechanics of materials, and properties of flexible bodies. Some assumptions about the
shape and orientation of structural members are made, and environmental testing verifies
and reinforces calculations made using these assumptions. Structural modeling and
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verification is performed to predict how a spacecraft structure will react under thermal
and environmental loadings during launch and throughout its lifetime.23
Two types of displacements are of particular interest when modeling a spacecraft
structure. Translation is considered in all three directions, measured in units of length.
Rotation is considered about the x-, y-, and z-axes, measured in radians. A typical rigid
body is one that does not deform and has six degrees of freedom (DOFs), three
translations and three rotations. These DOFs together completely describe a rigid body.
A flexible structure has infinite DOFs. In the structural modeling process, the rigid body
assumption is often made in order to limit the number of DOFs of a particular part of the
spacecraft structure and provide for ease of modeling.23
2.1.1.1 Statics
The first step in modeling a structure is to examine its static equilibrium state.
The reaction forces, apparent after constructing a free-body diagram (FBD) for the
structure, must balance the applied forces in order to keep the body at rest or moving with
constant velocity. Forces and moments summed about their particular axes, as shown on
the FBD, should equal zero:17
ΣFx = 0, ΣFy = 0, ΣFz = 0, ΣMx = 0, ΣMy = 0, ΣMz = 0 (2-1) 17
For statically determinate structures, the reactions are determined using equilibrium
equations. For statically indeterminate structures, the reactions are dependent on how the
structure deforms. Reaction interfaces serve to constrain motion in certain directions.
For example, a pinned end constrains all translations but no rotations.23 The structure
design and materials chosen must be adequate to statically support the spacecraft
structure prior to external loading.
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2.1.1.2 Dynamics
If unbalanced external loads act on a structure, dynamic principles must be
considered in the structural model. External forces cause the body to accelerate, and
external and inertial forces balance each other so that the body remains in dynamic
equilibrium. Body acceleration occurs in the direction of the vector sum of the forces,
which no longer equals zero. Newton’s Second Law of motion governs dynamic
equilibrium:
ΣFexternal = ma (2-2)
where m is the mass of the structure and a is its acceleration
An example of a structure in dynamic equilibrium is a rocket during uniform
thrust.23 A spacecraft is in steady-state acceleration during launch. Certain launch
vehicles impose certain load factors on a spacecraft. A load factor is defined as the
product of the mass of the spacecraft times the acceleration on the spacecraft. The
structure of the spacecraft is designed to withstand the loads imposed by the selected
launch vehicle.39 Dynamic loading imposes random vibrations on the spacecraft
structure. The effects of these vibrations on the structure’s integrity are discussed in
section 2.1.1.4.
2.1.1.3 Mechanics of Materials
Mechanics of materials describes how structural members react to environmental
loads. Since the spacecraft structure is by function a load-bearing structure, the stresses
and deformations must be modeled for strength verification. The structure’s ability to
withstand normal and shearing stresses as well as in-plane and out-of-plane deformations
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should be adequate for the structure’s desired lifetime. Table 2 lists some properties of
commonly used materials that aid in determining the integrity of a structure. The
stiffness and strength of load bearing structures depend on cross-sectional area, length,
and material.4
Table 2: Properties of commonly used spacecraft structure materials4
Tension Compression Shear Tension Shear
Steel: Structural (ASTM-A 0.284 58 -- -- 36 21 29 11.2 6.5 23
High-strength low a 0.284 70 -- -- 50 30 29 11.2 6.5 22
High-strength low a 0.284 67 -- -- 46 -- 29 11.2 6.5 21
High-strength low a 0.284 60 -- -- 42 -- 29 11.2 6.5 24
Quenched and tem 0.284 110 -- -- 100 55 29 11.2 6.5 18
Cold-rolled stainle 0.286 125 -- -- 75 -- 28 10.8 9.6 12
Annealed stainless 0.286 95 -- -- 38 22 28 10.8 9.6 50
Aluminum: Alloy 1100-H14 0.098 16 -- 10 14 8 10.1 3.7 13.1 9
Alloy 2014-T6 0.101 66 -- 40 58 33 10.9 3.9 12.8 13
Alloy 2024-T4 0.101 68 -- 41 47 -- 10.6 -- 12.9 19
Alloy 6061-T6 0.098 38 -- 24 35 20 10.1 3.7 13.1 17
Alloy 7075-T6 0.101 83 -- 48 73 -- 10.4 4 13.1 11
Magnesium: Alloy AZ80 0.065 50 -- 23 36 -- 6.5 2.4 14 6
AlloyAZ31 0.064 37 -- 19 29 -- 6.5 2.4 14 12
Titanium Alloy 0.161 130 -- -- 120 -- 16.5 -- 5.3 10
Copper: Aluminum bronze 0.301 90 130 -- 40 -- 16 6.1 9 6
Manganese bronze 0.302 95 -- -- 48 -- 15 12 20
Cold-rolled yellow 0.306 74 -- 43 60 36 15 5.6 11.6 8
Annealed yellow b 0.306 46 -- 32 15 9 15 5.6 11.6 65
Cold-rolled red bra 0.316 86 -- 46 63 -- 17 6.4 10.4 3
Annealed red bras 0.316 39 -- 31 10 -- 17 6.4 10.4 48
Ductility, percent elongation in 2
in.
Yield strength, ksiModulus of
Elasticity, 106
psi
Modulus of
Rigidity, 106 psi
Coefficient of thermal
expansion, 10-
6/oFCategory Material
Specific weight,
lb/in-3
Ultimate strength, ksi
Stress is the most basic concept in the mechanics of materials. It is defined as the
load acting in a certain direction over a cross-sectional area. Normal stress, σ, is the load
acting normal to an object: 4 (see Figure 1)
σ = P / A (2–3) 4
where P is the load acting in either tension or compression parallel to the long axis of the
member and A is the cross-sectional area perpendicular to the specified axis.1
Shear stress, τ, is the out-of-plane load acting parallel to the same cross-sectional
area:
τ = V / A (2-4) 4
9
where V is the load acting perpendicular to the same axis as described in the normal
stress equation. 4
Figure 1: Diagram showing normal and shear loads on a rigid bar
Stress analysis considers uses the rigid body assumptions to statically model a
structure. Strain is another important concept of structures that takes into account the
deformations caused by applied loads on the structure. Avoiding loads that cause plastic
deformations is imperative to the success of a spacecraft structure. Normal strain, ε, is
the deformation per unit length of a structural member along the axis of loading: 4
ε = δ / L (2-5) 4
where δ is the change in length of the member and L is the original length of the member
being deformed. 4
Shear strain, γ, is the angular deformation of a structural member, which is found
using Hooke’s Law and knowing the shear stress: 4
γ = Gτ (2-6) 4
Normal Force (P) Cross-Sectional Area (A)
Area (A)
Cross-Sectional Area (A)
Area (A)
Shear Force (V)
10
where G is the modulus of rigidity of the material (see Table 1) and τ is the shear stress
associated with the direction of angular deformation. 4
Stress and strain modeling serves to verify the strength quality of a structure, but
developing an optimal structure for space flight is more complicated. Some degree of
failure must be accepted to design an adequate structure that is also light and cheap.
Structural reliability is never fully defined due to material flaws and environmental load
uncertainties. Making design criteria assumptions allows for approximations in structural
reliability. Such criteria include:39
1. Design allowable strength – the structure has 99% chance of withstanding
predicted stresses and loads, based on material heritage
2. Design limit load – the maximum load expected, equals the mean value load
(available from environmental data) plus three standard deviations
3. Factor of safety (FS) – factor applied to the design limit load to further
prevent structural failure
4. Design stress – stress caused by the design limit load multiplied by the FS,
must be lower than design allowable stress
5. Margin of safety (MS) – allowable strength (load or stress) divided by the
design strength (load or stress) minus 1; value should be positive and as close
to zero as possible
Application of these criteria during the modeling process allows for over-design and
acceptance of a small failure risk. 39
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2.1.1.4 Flexible-Body Dynamics
The last step in the modeling process is to consider multiple DOF systems, or
flexible structures. The rigid-body assumption is relaxed. Flexible structures fail only at
low vibration frequencies. Frequencies of concern depend on the structure’s size and
shape, and on the environmental forces. This section addresses failures associated with
the primary structure’s modes of vibration and the associated natural frequencies.
Another concern is modes associated with the coupled loads of components attached to
the primary structure.23
Modes of vibration are of concern because at the mode natural frequencies the
structure experiences its highest displacement amplitudes. These displacements impose
high stresses on the spacecraft structure. Each mode of vibration of a structure has an
associated mode shape and natural frequency. A mode shape refers to the deformed
shape of a vibrating structure.23
Basic structural dynamics equations can be used to determine modes, natural
frequencies, and displacement amplitudes. However, computer programs, such as
IDEAS, are typically used for structural modeling. IDEAS uses finite element analysis of
a structure to predict modes, areas of maximum deformation, and maximum stress values.
Finite element analysis is ideal for modeling structures with low natural frequencies in
the first, second, and third modes. Statistical energy analysis is used to model high
frequency structures, but these are not usually of concern due to their small displacement
amplitudes.
Historical vibro-acoustic data is used to develop test specifications for
determining structural modes. This data aids in simulating the expected environment to
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which the structure will be subjected.39 A force-time history for random-vibration
environments is unpredictable, however the input random vibrations may be
characterized and a frequency-domain spectrum may be predicted. This spectrum is used
to estimate peak point accelerations, loads, or stresses. 23 A structure can be tested using
the simulated environment to verify modes and natural frequencies found using IDEAS,
or other modeling techniques. 39 Testing is the final step in modeling spacecraft
structures.
2.1.2 Interactions
The structure serves as the housing for all components on the spacecraft, and
therefore interacts with most other subsystems. The different degrees of subsystem
interactions with the structure are illustrated in the first row of Table 1. These
interactions are based upon the two main impacts of the structure on the mission: it
serves as the housing, structural support, and protection for all other components of the
spacecraft, and it makes up 10-20% of the mass as well as the size and shape of the
spacecraft bus. 39
Several subsystems interact with the structure only because the structure holds
their components. These subsystems include guidance and navigation, communications,
and command and data handling (C&DH). Some other subsystems with weak
interactions include propulsion, mission operations, and astrodynamics. The structure
houses all components of the propulsion system. Propulsive efficiency and
astrodynamics are dependent on the mass of the structure. Possible repairs on the
structure will involve mission operations. 39
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The remaining subsystems interact strongly with the structure. The mechanisms
on a structure do not interact as strongly as launch vehicle selection, ADCS, power,
thermal, environment, and cost modeling. Mechanisms serve as assemblies that make the
functions of other components possible. They are attached to the structure, but they only
serve as an active interface between the structure and the components that they support.
For example, an antenna is attached to a structure through a gimbal mechanism. This
gimbal serves to point the antenna in a desired direction so that the antenna may
successfully complete some task laid forth by the communications subsystem. In this
case, the mechanism only serves as the interface between the communications subsystem
and the structure. 39
Launch vehicle selection is based on the size, shape, and mass of the spacecraft
and is therefore highly dependent on the structure. The structure must also interface with
the separation system inside the launch vehicle, and must bear the loads imposed during
launch and separation event shocks. The ADCS depends on the structural housing to
support its control actuators and attitude sensors. It also depends on the shape, size, and
mass of the structure for accuracy. Attitude control actuators, such as momentum wheels
and control moment gyros, use the mass, center of gravity (cg), and moments of inertia of
the spacecraft for pointing control. The power subsystem depends on the structure to
house its internal components such as wiring, batteries, and connectors, and its external
components such as solar cells. The exterior of a spacecraft structure is often covered
with solar cells, or may have a solar cell structure attached to it that is larger than the
bus.39
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Materials chosen for the structure of a spacecraft often reflect thermal and
environmental concerns of the mission. The material of a structure should provide a path
for heat to be channeled away from internal components that could overheat. These
materials should also have low thermal expansive properties due to potentially high
temperatures. The structure protects all internal components of the bus from harmful
environmental affects such as corrosion, orbital debris, and random vibrations. Structural
materials also have a large effect on the overall cost of the structure. Materials should be
as inexpensive as is possible for their desired characteristics as well as being easy to
machine to lower fabrication costs. 39
2.2 Mechanisms
2.2.1 Modeling
Aerospace mechanisms are divided into high- and low-cyclic applications. High-
cyclic applications, such as antenna pointing and tracking or attitude control reaction
wheels, require frequent or constant manipulation. Low-cyclic applications, such as
antenna deployment or solar array retention, restrain a payload on launch or retrieval, or
they propel the payload to the deployed or restored position. An important requirement
for all spacecraft mechanisms is the demand for precision pointing and a long operating
life. 39
Functional requirements for the mechanisms derive from mission requirements
and divide into torques or forces and operating rates. Figure 2, an operating rate profile,
establishes the payload deployment rate.
15
Figure 2: Operating rate profile39
From this profile, the maximum angular acceleration, α, is determined. With the payload
moment of inertia determined, MOI, the mechanism's operating torque can be calculated
as:
T = αMOI (2-7) 39
Generally, for rough torque sizing a 20% friction torque is added to the operating
torque.39 The constant-speed part (s2), in Figure 2, represents the mechanism operating
torque since there is no acceleration at this point. With the two operating points known,
s1 and s2, a torque-speed curve (Figure 3) can be generated.
Figure 3: Torque-speed curve39
16
This linear relationship establishes the stall torque and theoretical no-load speed for the
mechanism. Figure 4 allows for approximations of the mechanism parameters with a
known stall torque.
Figure 4: Stall torque plot39
In addition to these parameters, spacecraft mechanisms must withstand the launch and
vibration tests. The mechanisms must also operate in the space environment where the
thermal vacuum will influence the selection of materials, lubricants, and coatings. 39
2.2.2 Interactions
There are several elements of a space mission which are not directly affected by
the mechanism design – astrodynamics, guidance and navigation, program management,
and the political, legal, and economic issues. The mechanism design/selection will not
impact these systems.
Spacecraft mechanisms interact slightly with the C&DH, thermal, launch vehicle,
and mission operations segments. The C&DH system sends signals to high-cycle and
low-cycle mechanisms, so each mechanism must be designed to interface with the
C&DH system. This interface changes depending on the design and function of the
mechanism. The thermal system employs and interfaces with various mechanisms.
17
High-cycle mechanisms may be used in the various cooling loops to open and close
valves and to force cooling fluid flow. The launch vehicle interfaces with several
spacecraft mechanisms as well, since low-cycle mechanisms may be used to separate the
spacecraft from the launch vehicle. The mission operations segment may also be
impacted by the characteristics of the various mechanisms. Mechanisms designed to
require human control need a more extensive ground support system than would
autonomous mechanisms. 39
Other subsystems exhibit strong interaction with the spacecraft mechanisms:
propulsion, ADCS, power, communications, and cost. Many mechanisms are utilized in
the propulsion system; some can regulate fuel flow for conventional thrusters, others can
deploy or retract the tether for electromagnetic propulsion. The design of certain
mechanisms directly affects the effectiveness and reliability of the spacecraft propulsion
system. The ADCS system also interacts heavily with various spacecraft mechanisms. If
momentum wheels are used for attitude control, high-cycle mechanisms are required to
accelerate and decelerate the wheels. If electromagnetic or conventional thrusters are
used for attitude control, the mechanisms and ADCS systems interact similarly to the
mechanisms/propulsion interactions. The power subsystem requires several mechanisms,
both high-cycle and low-cycle, in order to provide electricity for the spacecraft. Solar
panels need a low-cycle mechanism for deployment, and high-cycle mechanisms to orient
the panels toward the sun. The communications system interact strongly with spacecraft
mechanisms in a similar manner to the power system; antennas must be deployed and
positioned to communicate with ground and/or space systems. Also, the cost of the
spacecraft relates directly to the selection/design of the various mechanisms. Using low-
18
cost, off-the-shelf mechanisms costs much less than researching, designing, and
fabricating new spacecraft mechanisms. 39
2.3 Launch Vehicle Selection
2.3.1 Modeling
Table 3 presents data useful in selecting a launch vehicle for any given space
mission. The first and second columns of the table present the family and model number
of the launch vehicle, such as Atlas II or Ariane 40. In general, each family contains a
base model and several different configurations, created by upgrading the vehicle stages
or adding strap-on boosters. The third column of the table indicates the location of the
primary launch site for the vehicle, in degrees of north latitude. Every launch vehicle
presented in Table 3 is launched from a site north of the equator. The fourth column
indicates the nation that builds the launch vehicle, and the fifth column indicates the date
that the vehicle was first launched.
The ‘cost range’ column lists the approximate cost of the launch vehicle, in
millions of US dollars, where such figures could be found. The ‘payload dimensions’
columns indicate the maximum length and diameter of the payload that will fit in the
launch vehicle’s faring. When two pairs of dimensions are given, the launch vehicle can
be fitted with an alternate faring, either to increase the allowable payload size at the cost
of vehicle performance, or to decrease payload and faring size to gain performance. The
maximum mass that can be lifted to low Earth orbit (LEO) or geostationary transfer orbit
(GEO) are presented in the next columns, where such data could be found. Available
data varies slightly, but a typical LEO is defined as a one hundred nautical mile (185 km)
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circular orbit at the approximate inclination of the launch site. The last column presents
the legacy of the launch vehicle, listing number of successful launches over the number
of attempted launches.
20
Table 3: Launch vehicle data 12,19,33
First Cost RangeFamily Model (deg.) Country Year (Approx, $M) Dia. 1 Len 1 Dia. 2 Len 2 LEO GTO Success/Total As Of Atlas I 28.5 USA 1990 77-88 3.3 7.75 4.19 9.74 5,820 2,375 - -
II 28.5 USA 1992 84-88 3.3 7.75 4.19 9.74 6,580 2,610II-Star 48B 28.5 USA 1992 100-104 3.3 3.89 4.19 5.88 4,439 -
IIA 28.5 USA 1993 90-104 3.3 7.75 4.19 9.74 7,280 2,745IIA-Star 48B 28.5 USA 1993 106-120 3.3 3.89 4.19 5.88 5,139 -
IIAS 28.5 USA 1993 110-142 3.3 7.75 4.19 9.74 8,640 3,379IIAS-Star 48B 28.5 USA 1993 126-158 3.3 3.89 4.19 5.88 6,499 -
IIIA 28.5 USA 1999 - - - - - - 3,400 - -IIIB 28.5 USA - - - - - - - 4,500 - -
V (30x, 40x) 28.5 USA 2001 - - - - - - 5,100 - -V (5xx) 28.5 USA - - - - - - 8,200 - -
Athena I 34.7 USA 1997 - 1.98 4.29 - - 794 - - -II 28.5 USA 1998 - 2.74 6.65 - - 1896 - - -
Conestoga 1229 38 USA - 13-14 1.63 1.94 1.63 4.23 363 - - -1379 38 USA - 15-16 1.63 1.94 1.63 4.23 771 - - -1620 38 USA 1995 18-19 1.63 1.94 1.63 4.23 1,179 - - -1669 38 USA - 15-16 1.63 1.94 1.63 4.23 1,361 - - -1679 38 USA - 21-22 1.63 1.94 1.63 4.23 1,497 - - -3632 38 USA - 21-22 1.63 1.94 1.63 4.23 2,141 - - -5672 38 USA - 27-28 1.63 1.94 1.63 4.23 - - - -
Delta II-7925 28.7 USA 1990 55-60 2.54 4.67 2.79 4.08 2,760 1,450II-7920 28.7 USA 1991 49-60 2.54 6.38 2.79 5.78 5,045 1,270
III 28.7 USA - - - - - - - 8,345IV (EELV) 28.7 USA - - - - - - 17700 -
Pegasus Pegasus 28 USA 1990 8-14 1.27 1.9 1.27 2.14 455 125XL-C Star 24C 23 USA - 14-14 1.27 1.39 1.27 1.53 544 -XL-C Star 27 23 USA - 14-14 1.27 1.14 1.27 1.38 544 -
Taurus Taurus 28.5 USA 1994 21-26 1.37 2.54 1.37 2.54 1,420 514 5/5 1995XL Star 37XFP 28.5 USA - 23-26 1.37 2.54 1.37 2.54 1,565 595 - -XL/S Star 37FM 28.5 USA - 23-26 1.37 2.36 1.37 2.36 1,980 736 - -
Titan IIG/Star 37 28.7 USA 1988 33-38 2.84 5.13 2.84 6.66 2,655 -IIG/Star 48B 28.7 USA 1988 33-38 2.84 5.13 2.84 6.66 2,655 -
IIS-SSPS 99 USA 1988 44-55 2.84 2.41 2.84 3.94 2,445 -IIS-PAM D2 99 USA 1988 44-55 2.84 5.58 2.84 7.11 2,885 -
IIS-4SRM-SSPS 99 USA 1988 44-55 2.84 2.41 2.84 3.94 3,342 -IIS-4SRM-PAM D2 99 USA 1988 44-55 2.84 5.58 2.84 7.11 3,665 -
IIS-10GEM 28.7 USA 1988 44-55 2.84 5.13 2.84 6.66 5,470 -III/TOS 28.6 USA 1989 165-245 3.65 10.58 3.65 10.58 14,515 11,000 - -
IV/Centaur 28.6 USA 1994 435-475 4.57 15.7 4.57 15.7 18,144 -IV/SRM/Centaur 28.6 USA - 435-475 4.57 15.7 4.57 15.7 18,144 -
IV/SRM/IUS 28.6 USA - 330-435 4.57 10.5 4.57 10.5 23,350 2,360Scout Scout 34.7 USA 1979 10-12 0.86 3.27 1.07 3.78 270 54 99/113 1995
Space Shuttle Space Shuttle 28.6 USA 1981 130-245 4.7 18.6 - - 27,100 5,900 103/104 2001Long March CZ-1D 41 China 1991 10 2.05 3.99 - - 720 200
CZ-2C 41 China 1975 20 2.2 3.144 3.35 7.125 3,200 1000CZ-2E 28 China 1990 40 4.2 11.95 - - 9200 3370
CZ-2E/HO 28 China 1995 - 4.2 11.95 - - 13600 4500CZ-3 28 China 1984 33 2.6 5.84 3 7.27 5000 1500
CZ-3A 28 China 1992 - 3.35 8.89 4 12 7200 2500CZ-4 - China 1998 - 2.9 4.91 3.35 8.48 4000 1100
Ariane 40 5.2 Europe 1990 71-82 4 8.6 4 9.6 4900 207042P 5.2 Europe 1990 73-85 4 8.6 4 9.6 6100 292044P 5.2 Europe 199x 77-88 4 8.6 4 9.6 6900 338042L 5.2 Europe 199x 100-112 4 8.6 4 9.6 7400 3450
44LP 5.2 Europe 1988 106-108 4 8.6 4 9.6 8300 417044L 5.2 Europe 1989 130-141 4 8.6 4 9.6 9600 4700
5 5.2 Europe 1999 118-130 5.4 5.25 4.57 11.75 18000 6800 2/3 1998SLV ASLV 13.9 India 1987 - 1 3 - - 150 - - -
PSLV 13.9 India 1991 - 3.2 8.3 - - 3000 - - -GSLV 13.9 India 1995 - - - - - 8000 - - -
Shavit Shavit 31 Israel 1988 22 - - - - 2/2 1995H-Vehicle H-1 30.2 Japan 1986 90 2.44 7.91 - - 3200 1100 7/7 1995
H-2 30.2 Japan 1993 157 4.07 12 - - 10,500 4,000 6/6 1998J-Vehicle 1 30.2 Japan 1996 - - - - - 1000 - - -M-Vehicle M-3S-II 31.2 Japan 1985 29-31 1.65 6.85 - - 780 517 - -
M-V 31.2 Japan 1995 38-40 2.5 8.8 - - 1950 1215 2/2 1998Kosmos C-1 62.8, 48.4 Russia 1964 - 2.4 5.72 - - 1350 - 393/415 1995Energia EUS 51.6 Russia 1987 120 5.5 37 - - 88000 - - -
EUS/RCS 51.6 Russia 1987 120 5.5 37 - - 88000 - - -Proton D-1 51.6 Russia 1967 35-70 3.3 5.1 3.68 7.45 20000 -
D-1-e 51.6 Russia 1967 35-70 3.3 5.1 3.68 7.45 20000 5500D-1-e Star 27 51.6 Russia 1970 59-82 3.3 3.86 3.68 6.21 20000 -
D-1-e Star 48B 51.6 Russia 1970 59-82 3.3 3.27 3.68 5.62 20000 -M-5 51.6 Russia 1995 59-82 3.68 7.5 - - - -K 51.6 Russia 1996 - - - - - 20900 4800M 51.6 Russia 1996 - - - - - 22500 5500
Tsyklon F-1-m 51.6 Russia 1967 12 2.13 14.14 - - - - 114/116 1995F-2 51.6 Russia 1977 12 2.7 10.13 - - 4000 - 108/110 1995
Vostok A-1 51.6 Russia 1959 14 2.7 6.8 - - 4730 - 88/89 1995Molniya A-2-e 51.6 Russia 1961 - 2.7 7.8 - - 1500* - 265/284 1995Soyuz A-2 51.6 Russia 1963 18 2.85 9 - - 7000 - 671/688 1995Zenit J-1 45.6 Russia 1985 77-82 3.9 11.155 3.9 13.65 13740 4300 23/27 1998
Launch Vehicles Launch
23/25 1998
Payload Dimensions (m) Payload Mass (kg) to:
1998
170/198 1998
18/18 1998
20/22 1998
Operations
216/232 1998
199849/54
93/99 1998
142/155
21
2.3.2 Interactions
As seen in Table 3, there are many types of launch vehicles from which to choose.
No matter which launch vehicle is chosen for any project, the selection will influence
other parts of the overall mission. The load factors produced by the launch vehicle are
one of the strongest influences on the payload.
The selection of a launch vehicle greatly influences the astrodynamics, mission
geometry, cost, and structures portion of a given mission. Not all launch vehicles can get
a satellite to High Earth Orbit (HEO) or launch from specific latitudes so different flight
paths and orbits will be used depending on the selection. Payload size, weight, and
vibrations produced vary with each launch vehicle. Larger and more massive payloads
usually require larger and more powerful launch vehicles, which will increase the cost of
the mission. If the selected launch vehicle produces more vibrations than the structures
can withstand, it will shake the payload apart before it arrives in orbit.
Launch vehicle selection weakly interacts with the following: mission analysis,
thermal, environment, mission operations and ground systems, politics, and mechanisms.
The launch vehicle imposes a hazardous environment to the spacecraft, but only in the
beginning of the mission. If a specific launch vehicle produces more heat than the
structures and mechanisms inside the payload bay can withstand, the mission will need to
be re-analyzed to determine if a new launch vehicle should be selected or if different
materials should be used on the payload. Most launch vehicles can be launched from
only one location, so if the selected vehicle can only be launched from Russia then the
entire ground system crew will need to be moved there.
22
Launch vehicle selection does not interact with ADCS, communications, C&DH,
and power. These sub-systems are all independent of the selected launch vehicle since
they only apply to the payload that is inside it. Once the payload is in the launch vehicle
and on the way to orbit there is nothing that can be done about these sub-systems until the
satellite is released into space.
2.4 Summary
There are many equations used to describe the structures and mechanisms of the
spacecraft. There are also many different types of launch vehicles to choose from, each
with advantages and disadvantages. The structures, mechanisms and launch vehicle
interact in varying degrees with all of the systems of the spacecraft. The models and
interactions presented in this chapter provide tools by which a system may be designed to
fulfill mission requirements. This system must be assembled into the final product,
however, this requires research into the manufacturer, mass, and cost of each individual
part. The next chapter presents examples of structures, mechanisms, and launch vehicles
that could be used in a space mission.
23
3 Subsystem Examples
There are hundreds of examples of spacecraft structures and mechanisms
available today, as well as dozens of launch vehicle options. While it is impossible to
present an example of every structure, mechanism, and launch vehicle that exists, this
chapter attempts to give a sample of the hardware that is available. Some of these
examples are presented with data on cost, availability, and suppliers. The references
cited in this chapter may also be used for further information.
3.1 Structures
3.1.1 Materials
This section discusses three different categories of materials used to build
spacecraft structures including metals, composites, and metallic shape-memory alloys.
Both raw metals and metal alloys are used in structural applications because of their high
strength, toughness, and ease of machinability. Composites are used because of their
stiffness and relatively low mass. Shape-memory alloys are sometimes used for their
resilience.
3.1.1.1 Metals
Metals have been used in spacecraft structural applications since Sputnik, launched
on October 4, 1957. The bus of this spacecraft consisted of an aluminum sphere. More
recent spacecraft utilized stronger, lighter materials such as titanium, which was used for
24
the structure of the Mars Odyssey launched April 7, 2001. For the most part, all
spacecraft’s bus structures are metallic.
Upon completion of all structural design, materials are chosen and bought for
fabrication. A general rule when choosing metals is to first find out if the metals are
flight quality. The NASA-MSFC Materials and Processes homepage20 provides a listing
of all NASA approved metallic flight materials. Once it is certain that a metal is NASA
approved, a supplier of that metal is located. The same materials website listed above, or
the Thomas Register website31 are used to search for metal suppliers. Both the MSFC
website and the Thomas Register website list thousands of materials to choose from and
corresponding suppliers. Costs vary with quantity ordered. Most materials have a lead-
time of no longer than a week.
3.1.1.2 Composites
Composite materials are created from two or more different materials to obtain
properties that neither material exhibits alone.23 Composites are utilized in spacecraft
structures primarily for their low mass and high stiffness. Aluminum can be 30% - 80%
heavier than composites. Composites also have a very low coefficient of thermal
expansion and can be fabricated for high stiffness and high strength.
Workmanship plays a large role in strength of composites. Additional testing
may be a necessary to validate the properties of the composite. Analysis of composite
structures requires accounting for property changes in specific directions due to the non-
isotropic properties of the material. Composites tend to be brittle and can be hard to
fasten.39
25
Composite materials typically cost more than other materials. However, the cost
of the entire system may be cheaper with composite products because of their higher
stiffness-to-mass and strength-to-density ratios compared to traditionally used metals.
These benefits coupled with resistance to fatigue, high thermal conductivity, and low
thermal distortion make composites a worthy choice for spacecraft materials.40 More
information about composites, including properties and analysis, can be found in the
Fiber Composite Analysis and Design handbook.10
The type of binding matrix and reinforcement most commonly classifies
composites. The four main classes of composites include polymer matrix, carbon matrix,
metal matrix, and ceramic matrix. Polymer matrix composites (PMC's), including carbon
(graphite) epoxy, are commonly used in spacecraft solar arrays and antenna covers.
Carbon matrix composites are beneficial in high temperature environments and are
especially useful in launch vehicles. Metal matrix composites are ideal for use in
gimbals, hinges, and brackets. 23, 40
Matrices hold the reinforcing fibers together and transfer shear to provide a
desired tensile strength or Young's modulus. Typical matrices include polymers
(epoxies), metals, carbons, and ceramics. The reinforcement provides extra strength and
stiffness and is either continuous or discontinuous. Aramid (Kevlar or Spectra), graphite,
and glass are typical reinforcers. Polymer-matrix composites are the most commonly
used and consist of a reinforcing phase and carbon fibers in an epoxy (polymer) matrix.
Graphite reinforcement with epoxy matrix is the most widely used PMC. Table 4 lists
advantages, disadvantages, and applications of typical PMC materials. 23
26
Table 4: Characteristics of typical Polymer-matrix composites
Material Advantages Disadvantages Typical Applications Aramid/Epoxy (eg. Kevlar fibers w/ epoxy matrix)
-Impact resistant -Lower density than graphite/epoxy -High strength-to mass ratio
-Absorbs water -Outgasses -Low strength -Negative coefficient of thermal expansion
-Solar array structures -RF antenna covers
Carbon/Epoxy -Very high strength-to mass
ratio -High modulus-to-mass ratio -Low coefficient of thermal expansion -Flight heritage
-Outgasses -Absorbs water
-Truss members -Face sheets for sandwich panels -Optical benches -Monocoque cylinders
Graphite/Epoxy -Very high modulus-to-
mass ratio -High strength-to-mass ratio -Low coefficient of thermal expansion -High thermal conductivity
-Low compressive strength -Ruptures at low strain -Absorbs water -Outgasses
-Truss members -Antenna booms -Face sheets for sandwich panels -Optical benches -Monocoque cylinders
Glass/Epoxy -Low electrical conductivity
-Well-established manufacturing process
-Higher density than graphite/epoxy -Lower strength and modulus than graphite/epoxy
-Printed circuit boards -RF antenna covers
Programs such as the Mars Global Surveyor, the Hubble Space Telescope, and
Clementine incorporated composite materials into their design. Clementine used a
composite isogrid as its solar array substrate and carbon (graphite) epoxy for its skin
structure.22 Several programs use Kevlar/fabric epoxy as face-sheets for solar array
substrates and graphite as solar array stiffeners. Composite Optics, Inc. produces solar
arrays, fuel tanks, and honeycomb structures made from composite materials.7 Applied
Aerospace Structures Corporation manufactures solar array structures, bus structures, and
electronics housings.3 Table 5 lists prices of various composite materials manufactured
by Aerospace Composite Products.2 Other manufacturers of aerospace composite
27
structures and suppliers of raw composite materials can be found using the Thomas
Register.31
Table 5: Prices of typical composite materials used in space structures2
Composite Material
Dimensions Price Notes
Carbon fiber laminate 0.030” thick 4” × 36”
$30.00 -Used to reinforce areas with high loads -High stiffness
Nomex honeycomb (Aramid fiber)
0.25” thick 12” × 12”
$16.00 -High stiffness-to-mass
Graphite plate (epoxy matrix)
0.08” thick 8” × 12”
$30.00 - $40.00 -High strength -Lightweight
Kevlar mat
25 oz. 4” × 35.5”
$12.50 -High impact resistance -High toughness
Aero mat ---- $10.00 / yard -Honeycomb foam mat
-Adds thickness -Flexible
Kevlar ribbon
0.125” wide 30” long
$2.00 -Used to stiffen structures
Carbon fiber ribbon
0.125” wide 30” long
$3.00 -Reinforcer
3.1.1.3 Shape-Memory Alloys
Shape memory alloys (SMA) are a group of materials that possess the ability to
“remember” their original shape, and return to it upon a temperature change. These
materials include nickel-titanium alloys and copper-base alloys. Table 6 is a list of
metallic compounds that demonstrate SMA characteristics.5
28
The fundamental property of these alloys is the ability to return to their un-deformed
shape upon heating, following a plastic deformation. There are two types of SMAs: those
that exhibit one-way shape memory and those that exhibit two-way shape memory. One-
way SMAs exhibit shape memory only upon heating. Two-way SMAs exhibit shape
memory upon heating and cooling. 5
Table 6: Alloys exhibiting shape memory effects5
Alloy Composition Transformation temperature range,
oC
Transformation temperature range,
oF Ag-Cd 44/49 at.% Cd -190 to -50 -310 to -60 Au-Cd 46.5/50 at.% Cd 30 to 100 85 to 212
Cu-Al-Ni 14/14.5 wt.% Al 3/4.5 wt.% Ni
-140 to 100 -220 to 212
Cu-Sn approx. 15 at.% Sn -120 to 30 -185 to 85 Cu-Zn 38.5/41.5 wt.% Zn -180 to -10 -290 to 15 In-Ti 18/23 at.% Ti 60 to 100 140 to 212 Ni-Al 36/38 at.% Al -180 to 100 -290 to 212 Ni-Ti 49/51 at.% Ni -50 to 110 -60 to 230 Fe-Pt approx. 25 at.% Pt approx.-130 approx.-200
Mn-Cu 5/35 at.% Cu -250 to 180 -420 to 355 Fe-Mn-Si 32 wt.% Mn, 6 wt.% Si -200 to 150 -330 to 300
Shape memory alloys have been used on recent space missions and are being
considered for future space applications. Specifically, SMAs are advancing solar array
technology. When used as hinges, or other parts of solar array structures, SMAs’ low
mass improves power-to-weight ratios. They also provide shock-free deployment, which
improves the dynamics of any spacecraft. Deployment devices made from SMAs are
cheaper, lighter, simpler, and more reliable than conventional technology.6
29
Some SMA materials can be found on the NASA-MSFC Materials and Processes
homepage20. Raychem Corporation manufactures the nickel-titanium alloy16, which has
been used in space applications.
3.1.2 Structures
3.1.2.1 Solar arrays
Lightweight solar arrays are desired for spacecraft so that the power-to-mass ratio
is as high as possible. Fixed panel and deployable panel are the two types of structural
possibilities for solar arrays. Fixed panel arrays can be used when a minimum amount of
power is needed. Deployable arrays are more desirable for high power requirements
since the arrays can be adjusted to absorb the maximum amount of solar radiation. Only
the solar array substrate is discussed in this paper. Discussions on solar cells can be
found in the power subsystem report.
A light honeycomb sandwich material is typically used in flat and rigid solar
arrays. Deployable arrays must be constructed to avoid low natural frequencies and to
maximize stiffness. Stored deployable solar arrays must be ground tested to ensure that
the loads imposed by launch will not destroy the array. Many substrates in the past were
constructed of aluminum, titanium, steel or metallic alloys. Composite materials are now
more common in array structures because of their low mass and high stiffness. 23, 38
Deployable arrays can be made flexible for storability. Roll-out arrays are
typically made of flexible sheet metal composed of stainless steel or beryllium copper.
Composite materials such as Kapton can also be used for roll-out arrays. Only
lightweight and limited extension panels can be used for roll-out systems. Motor, gears,
30
and a roller add mass to this type of system. An experiment performed by L'Garde, Inc
measured a roll-out array's specific power of about 100 W/kg. 23, 38
Inflatable arrays are another type of flexible, deployable array and are beneficial
because of their storability during launch and their low mass. These arrays are fully
extended when a gas is blown into the main blanket structure once the spacecraft reaches
orbit. Inflatable arrays can become permanently rigid and produce more power than roll-
out or rigid arrays. A sketch of an inflatable solar array is shown in Figure5. 38
Figure 5: Inflatable framing structure and panel 38
Fixed arrays are solar panels mounted directly on the spacecraft body. These
arrays are simple and reliable but provide relatively low amounts of power with respect to
the available surface area. Power collection depends on the incident angle of the sun so a
fixed array structure requires more surface area than a deployable, sun-tracking array.23
3.1.2.2 Tethers
Tether material performance is based on the strength-to-mass ratio and
conductivity of the material. Metals are typically used when conductivity is important for
31
electrodynamic propulsion or power generation. However, metals have a relatively low
strength-to-mass ratio of about 20 km. Composites have proved to have higher strength-
to-mass ratios and have been used in tether experiments since 1960. Figure 6 shows the
trend of tether materials through time.9
Figure 6: Trend of tether materials through time9
A tether can generate power as it is dragged through the Earth’s magnetic field.
Faraday’s law gives the amount of voltage that the tether can generate between its two
ends, Φoc:
Φoc = (v × BH)·L (3-1)15
where v is the orbital velocity, BH is the magnitude of the Earth’s magnetic field, and L is
the length of the tether. In a circular orbit, v is perpendicular to the Earth’s magnetic
field. For a spacecraft in a 400 km altitude circular orbit the velocity of the spacecraft is
approximately 7.7 km/s and BH is given as 2.6 × 10-5 T 15. So for a 20 km tether, the
voltage drop over the two ends of the tether equals 4004 Volts. The conductivity, σc,
32
length, L, and cross-sectional area, A, of the material gives the overall resistance, R, of
the tether by the following equation:
R = L / σc Α (3−2)15
Assuming the 20 km tether is constructed of aluminum with a cross-sectional area of
4 mm2 and conductivity of 3.5 × 107 (Ωm)-1, the overall resistance of the tether will be
143 Ω. Power is given by:
Pw = Φ2oc / R (3-3)
which makes the overall power generated by this tether equal to112 KW. This is for the
ideal case.
Material selection and configuration of the tether is also important for defining the
lifetime of tethers. Orbital debris poses a threat to any deployed tether since any collision
will probably slice the tether. Multiple parallel tethers or interconnected parallel tethers
increase the lifetime of the tether system. The Hoytether™ concept (Figure 7) developed
by Tethers Unlimited11 is a failsafe multiline tether design for long duration missions.
Figure 7: Sectional drawings of multiline Hoytether™ 11
33
3.1.2.3 Bus
There are typically four types of spacecraft main structures: trusses and frames,
skin-frame structures, monocoque cylinders, and cylindrical structures. Each of these has
its own set of design considerations, forms of construction, and materials. 23 This section
discusses in short detail each of these structures. Illustrations of each type of structure
can be found on pages 523 - 526 of Sarafin. 23
A truss structure can only withstand axial loads applied to its joints. Only trusses
whose members form triangles are considered structurally stable. A frame is a truss
whose members form polygonal shapes other than triangles, and can therefore carry shear
through its members as well as axial loading. A frame, however, is less efficient than a
truss. The weight efficiency of trusses and frames is highest for rectangular or triangular
bus cross-sections, and decreases as the cross-section becomes round. Trusses and
frames are usually machined out of one piece of metal, rather than pieced together from
individual members. Typical forms of construction include members made of sheet
metal and formed into structural shapes, truss sides machined from plate stock material
and fit together, and separately machined open-section members. Typical materials used
in truss and frame structures include aluminum, titanium alloys, and graphite/epoxy
composite. 23
A skin-frame structure consists of a framework made of stringers and lateral
frames covered with skin panels. Any bus shape is possible with this type of structure.
The skin in these structures is usually designed to buckle so that diagonal tension carries
shear. For stability, these structures must be closed on each end, include diagonal
members, or include frames radially internal to the structure. Typical forms of skin-
34
frame construction include machining frame members from sheet metal and using sheet
metal, sandwich construction, or isogrid for skin. Sandwich panels and isogrid are
discussed in more detail later in this section. Typical skin-frame materials include
aluminum, magnesium, and titanium alloys. 23
A monocoque cylinder is simply a cylindrical shell with no stiffeners or frames,
and is therefore limited in strength by buckling stress. Monocoque cylinders are only
effective under uniform axial loading over its cross-section. These structures cannot
support concentrated loads. Typical forms of construction include sheet metal or isogrid
rolled into a cylinder and sandwich segments either fabricated with curvature or pieced
together to form a cylinder. Sandwich or isogrid structures result in low mass. Typical
monocoque cylinder materials include aluminum and magnesium alloys or
graphite/epoxy composite. 23
Each of the cylindrical structures includes members that stabilize the skin and
help it carry loads, and include skin-stringer, stiffened-skin, and semi-monocoque
configurations. The stringers in a skin-stringer structure are attached to the skin and
designed to carry most of the axial load and bending after the skin buckles. Stiffeners in
a stiffened-skin structure are machined as part of the skin and are intended to increase its
buckling load. A semi-monocoque structure has no axial stiffeners, but intermediate ring
frames that increase the skin’s buckling load. Cylindrical structures are typically
constructed of aluminum alloys; the stringers, stiffeners, and skin are machined from
sheet metal. 23
Sandwich panels and isogrid (Figure 8) can be used for several of these types of
structures when high buckling strength relative to weight is desired. A sandwich panel is
35
formed of two thin face sheets, which carry axial loading and bending moments,
separated by a honeycomb core, which carries out-of-plane shear loading. A honeycomb
sandwich panel provides lightweight structures with high bending strength and stiffness.
Isogrid can be solid or open. It consists of a pattern of equilateral triangles machined
from plate metal, usually aluminum. Sandwich panels are lighter than isogrid panels, but
isogrid panels are stronger than sandwich panels. Components are attached to isogrid
panels through threaded inserts or studs at points where the isogrid ribs meet. Sandwich
panels require local potting material within the core to distribute loads induced by
fasteners. 23
Figure 8: Illustration of sandwich panel and isogrid structures23
3.2 Mechanisms
This section presents examples of low-cyclic and high-cyclic mechanisms. Some
examples of low-cyclic mechanisms include appendage deployment/retention
mechanisms and payload/launch vehicle separation mechanisms. High-cyclic
36
mechanisms include antenna pointing/tracking mechanisms, solar array drive
mechanisms, and reaction wheels.
3.2.1 Low-Cyclic Mechanisms
Low-cyclic mechanisms serve several purposes on a spacecraft, such as
restraining payload component during launch and deploying payload components after
launch. Specific examples include solar array and antenna retention mechanisms, solar
array and antenna deployment mechanisms, spacecraft/launch vehicle separation
mechanisms, and mechanisms for securing the spacecraft in the launch vehicle. Many of
the options for deploying and retaining solar arrays can be adapted for use with antennas;
consequently, solar array and antenna deployment/retention is discussed in the same sub-
section.
3.2.1.1 Solar Array/Antenna Retention/Deployment Mechanisms
Solar arrays and antennas can also be deployed using inflatable beams. These
beams contain an aluminum laminate layer and a layer of multilayer insulation. Upon
inflation, the aluminum laminate yields, and the resulting cold work causes the beam to
remain rigid after the beam is depressurized. These systems are reliable as long as the
pressurizing system used for inflation is reliable.
Light Flexible Solar Array Hinges (LFSAH) use a shape memory alloy to allow
controlled, shockless deployment of solar arrays or any other spacecraft appendage.
Electrical current is applied to the hinges, and the heat from the electrical resistance
causes the hinges to deploy the appendage. Although deployment can be controlled by
the amount of heat applied, the appendage cannot be retracted one deployed. Advantages
37
include low-shock controlled deployment, few parts, low mass, high reliability, and ease
of production and assembly. Although they have not been flight tested yet, LFSAH have
been tested in the weightless environment on STS-93, and is planned for use on the New
Millennium Earth Observer 1 (EO-1) and the Deep Space 3 (DS3) space vehicle34.
Alcatel26 provides information regarding several of their solar array deployment
mechanisms. Torsion springs, which use torque to deploy the solar arrays, mass around
0.210 kg per spring, provide from 1 to 6 N-m of torque. One the arrays are deployed, the
springs provide about 2000 N-m/rad of torsional stiffness to keep the solar array
deployed . Elastic hinges typically mass 0.25 kg per hinge, and provide about 0.2 N-m of
torque. Once the arrays are deployed, the hinges provide approximately 800 N-m/rad of
torsional stiffness. Alcatel alludes to other systems such as ADELE and mechanisms
using shape memory alloys, but does not provide any relevant information regarding
these systems. Attempts to contact this company have been unsuccessful.
The mechanism shown in Figure 9 can retain solar arrays during launch, release
them for deployment, and can recapture and stow them if necessary. A motor turns a ball
screw, which rotates the bell-crank linkage. This pushes the spring-loaded jaws outward
which causes them to open and release the two solar arrays. Limit switches measure the
position to indicate the end of travel22.
38
Figure 9: Solar array retention/deployment mechanism23
3.2.1.2 Payload/Launch Vehicle Separation Mechanism
One option for separating the spacecraft from the launch vehicle uses springs
between the spacecraft and the launch vehicle. Before launch, the springs are
compressed, and the spacecraft is fastened to the launch vehicle using pyrotechnic bolts.
Once in orbit, the pyrotechnic bolts are fired, thus allowing the springs to push the
spacecraft away from the launch vehicle. Springs and pyrotechnic bolts are relatively
low-cost, and do not require much lead time to obtain. Power requirements are minimal,
since power is only necessary to fire the pyrotechnic bolts22.
3.2.1.3 Payload Retention Device
A retention-latch actuator is an example of a mechanism used to hold payloads in
a cargo bay during launch. This device, shown in Figure 10, closes down over the
payload trunnions (interface shafts). Redundant alternating current brake motors, a
differential speed-reducing gearbox, and double four-bar linkage can prevent opening
under loads of 150,000 pounds22.
39
Figure 10: Retention-latch actuator22
3.2.2 High-Cyclic Mechanisms
High-cyclic mechanisms require frequent or constant articulation. Some typical
high-cyclic mechanisms used on spacecraft include antenna pointing and tracking, solar
array drives, and attitude control reaction wheels.
3.2.2.1 Antenna Pointing and Tracking
Astrium, the new company formed by the merger of Matra Marconi Space and the
space divisions of Daimler-Chrysler Aerospace provides an antenna-pointing mechanism.
The design, shown in Figure 11, is compact and flat. This mechanism has a mass of
5.6 kg and can support an antenna with a mass up to 22 kg. It has a rotation range of
+ 10° about two axes and a pointing accuracy of + 0.01°. The antenna pointing
mechanism has an operating temperature of + 90°C. Fourteen months are necessary for
the production and delivery of this mechanism.24
40
Figure 11: Antenna pointing mechanism24
Alcatel Space Industries also produces an antenna pointing and tracking
mechanism. This mechanism offers the ability to deploy the antenna in addition to
providing pointing and tracking. It has a low mass and is flight proven with over seventy
mechanisms sold worldwide and fifty mechanisms still in orbit. This mechanisms typical
lifetime is 100,000 cycles. It has a pointing angle of + 15° and a pointing accuracy of
+ 0.005°. It has an operating temperature of + 80°C. Twelve months are necessary for
the production and delivery of this mechanism. 24
3.2.2.2 Solar Array Drive Mechanisms
Astrium produces a solar array drive mechanism, shown in Figure 12. It has a mass
of 4.2 kg with 2 kg of redundant electronics. It requires 5 W of power during operation.
This solar array drive mechanism offers a reference position accuracy of + 0.6°. It has a
typical lifetime of greater than fifteen years. Fourteen months are necessary for the
reproduction and delivery of this mechanism.25
41
Figure 12: Solar array drive mechanism25
Alcatel provides a solar array drive mechanism, shown in Figure 13, compatible
with a large range of flexible or rigid solar arrays. It has a low mass of 3 kg and offers
full rotation capability. This mechanism can be easily customized to fit the needs of a
particular design. It has a step size of 0.12°. Twelve months are necessary for the
reproduction and delivery of this mechanism. Alcatel is also developing a new solar
array drive mechanism. This mechanism will be more powerful offering a step size of
0.01°. It will also be 2 kg more massive than the existing product.26
Figure 13: Alcatel solar array deployment mechanism26
42
3.2.2.3 Attitude Control Reaction Wheels
ITHACO, founded in 1962, produces many different attitude reaction wheels used
for three-axis reaction control. To date, ITHACO has produced over 150 reaction or
momentum wheels. Two examples of these reaction wheels are illustrated in Figure 14.
The maximum operating speed ranges from 1200 rpm to 6000 rpm for these wheels. The
momentum capacities range from 4 N-m-sec to 50 N-m-sec at their corresponding
maximum operating speeds. The mass ranges from 2.55 kg to 10.6 kg for the reaction
wheels. The additional mass of the motor drivers ranges from 0.91 kg to 3.3 kg. The
diameter for these wheels range from 20.5 cm to 39.3 cm.35
Figure 14: Reaction wheels35
Slew times can be determined for the lowest and highest capacity reactions wheels
produced by ITHACO. Slew times are calculated for slew angles of 30° and 90° using
the following equation:
T
It
θ4= (3-4)60
where t is the slew time (seconds), θ is the slew angle (rad), I is the moment of inertia of
the body (kg-m2), and T is the torque capacity (N-m-s). A body with a moment of inertia
of 500 kg-m2 can maneuver 30° in 4.4 minutes with the 4 N-m-s reaction wheel, but the
43
same body with a 50 N-m-s reaction wheel can rotate 30° in only 20.9 seconds. The
same body slews 90° in 62.8 seconds with the 50 N-m-s reaction wheel, or in 13.1
minutes with the 4 N-m-s reaction wheel.
The ITHACO reaction wheels have been used on such missions as the U.S. Air Force
Space Test Experiment Platform, the NASA Total Ozone Mapping Spectrometer - Earth
Probe, the APL Near Earth Asteroid Rendezvous satellite, the US Air Force Miniature
Sensor Technology Initiative satellite, and the NASA Mars Surveyor satellite.35
3.3 Launch Vehicle Examples
3.3.1 Launch Vehicles
There are dozens of vehicles currently available to lift a satellite payload into
orbit. Each of these launch vehicles is slightly different in its payload capacity, cost,
record, launch location, and many other factors. Examples of available launch vehicles
have already been presented in Table 3, which shows advertised performance and
reliability data for each. In this section, detailed descriptions are provided for two of the
more common vehicles in use today, the Ariane-4 and Titan IV. The Ariane 4 rocket is
the base model in a series of medium-lift launch vehicles produced by the European
Space Agency. It is capable of lifting between 4,900 kg and 9,600 kg to LEO, depending
on configuration. The Titan IV is a heavy-lift vehicle produced in the United States,
capable of lifting over 21,000 kg into LEO.
44
3.3.1.1 Ariane-4
Development for the Ariane rockets was authorized by the ESA in July of 1973,
after the unsuccessful Europa project was cancelled in April of the same year. The
Ariane-1, 2, and 3 rockets were first launched in 1979, 1986, and 1984, respectively, but
are no longer in production. The first Ariane-4 vehicle was first launched on June 15,
1988. Since then, it has been launched 106 times, with 64 consecutive successful
launches up to September of 200119.
The Ariane-4 (Figure 15) base launch vehicle is 58.4 meters tall, including a
payload faring that is 4 meters in diameter and 11.2 meters tall. Alternative fairings can
be mated to the vehicle for additional payload volume or increased vehicle performance.
The diameter of each of the fairings is the same, and the alternative heights are 8.6, 9.6,
2.8, and 3.8 meters. The mass of the base vehicle is 240,000 kg, but alternate
configurations can mass as much as 470,000 kg, based on the number of strap-on
boosters. 19
Figure 15: Configurations of the Ariane-4 launch vehicle8
45
The Ariane-4 is launched from Kourou, French Guiana, at 5.2° north latitude,
52.8° west longitude. Its primary missions are polar orbits and 7° geostationary transfer
orbits. 19
Both solid and liquid strap-on boosters can be used on the Ariane-4 vehicle. The
solid boosters have a nominal burn time of 34 seconds, and are jettisoned after burn-out
by four strong springs. The liquid boosters have a nominal burn time of 140 seconds, and
are jettisoned after burn-out using explosive bolts and small rockets. The first stage of
the Ariane-4 burns nominally for 205 seconds, the second stage for 126 seconds, and the
third stage for 725 seconds. 19
Launch processing for an Ariane-4 takes approximately nine weeks, beginning
with the transport of the vehicle hardware to the launch facility. The launch countdown
takes approximately 38 hours, spread over 3 days. If the countdown is halted more than
five minutes before launch, 24 to 48 hours are required to reset the vehicle. If the
countdown is halted less than 5 minutes before launch, the next attempt must be
postponed at least 10 days. 19
The cost of a dedicated launch in an Ariane-4 vehicle ranges from $45 to
$110 million, depending on configuration and number of boosters used. 19
3.3.1.2 Titan IV
The Titan IV is the most powerful launch vehicle produced by the United States.
Development of the Titan IV was initiated in 1985 when the US Air Force contracted
with Martin Marietta for 10 vehicles. The design is based on earlier Titan vehicles,
46
specifically the Titan 34D, and its first launch was executed successfully with no upper
stage on June 14, 1989. Since then, 30 Titan IV vehicles have been launched with a
success rate of over 95%.33
The Titan IV (Figure 16) is 62.2 meters tall, including the large payload faring
that is 5.1 meters in diameter and 26.2 meters in height. Smaller farings are available
with heights of 15.2, 17.1, 20.1, and 23.2 meters. The mass of the vehicle is
860,000 kg.19
Figure 16: Schematic of Titan IV launch vehicle18
Two primary launch sites are used to launch the Titan IV, Cape Canaveral Air
Force Station in Florida, and Vandenberg Air Force Base (AFB) in California.
Vandenberg AFB is located at 34.7° north latitude, and Cape Canaveral is located at
47
28.5° north latitude. The primary missions of the Titan IV are 28° geostationary transfer
orbits, 28° circular low earth orbits, and polar orbits. The Titan IV is capable of
launching 8,620 kg into a 28° geostationary transfer orbit, 21,640 kg into a circular 185
km LEO at 28°, and 18,600 kg into a circular 185 km polar orbit. 19
The Titan IV is launched with three stages and two Solid Rocket Motor Upgrades
(SRMU). Each of the two SRMUs produces an average thrust of 7.5 million N, with a
nominal burn time of 137.8 seconds. Stage one burns nominally for 164 seconds, stage
two burns nominally for 223 seconds, and stage three burns nominally for 104 seconds to
600 seconds, depending on configuration. 19
The cost of a dedicated launch in a Titan IV vehicle is approximately
$248 million, using the Martin Marietta Centaur upper stage. 19
3.3.2 Payloads
Launch vehicles are used to place spacecraft into orbit. This section presents
several examples of satellite payloads that were launched into space on United States
launch vehicles, with approximate mass and dimensions.
3.3.2.1 Shuttle Missions
The United States Space Shuttle is an important launch vehicle capable of lifting
heavy payloads and performing many complex deployment tasks. The Space Shuttle is
often used to launch many smaller payloads in a single mission, frequently those satellites
that require the personal attention of the astronauts. An example of a Space Shuttle
deployed mission is the TSS (Tethered Satellite System).
48
The TSS (Tethered Satellite System) was a sphere on the end of a 20 km long
tether. The sphere measured 1.6 meters in diameter and had a mass of 518 kg. The TSS
was built by the Italian Space Agency and was filled with scientific instruments. The
tether was 2.5 mm in diameter and ran between the shuttle and the sphere. The TSS was
launched on STS-46 (Space Transportation System), aboard the Space Shuttle Atlantis,
on July 31, 1992.30
3.3.2.2 Titan IV Missions
The Titan IV is a heavy lift US expendable launch vehicle used to launch many
large payloads for the US government. An example of a large spacecraft launched on a
Titan IV vehicle is the Cassini-Huygens probe to Saturn. The Cassini spacecraft is one of
the largest and heaviest spacecraft every built, massing about 5,600 kg at launch. The
spacecraft was attached to a Lockheed Martin-built Centaur upper stage, and was over
6.8 m tall and 4 meters in diameter14. Even though this is one of the largest spacecraft
every launched, it fit easily within the mass and size restraints imposed by the Titan IV
vehicle. Cassini was launched on October 15, 1997 from Cape Canaveral.
3.3.3 Summary
This chapter presents common examples of spacecraft structures, spacecraft
mechanisms, and expendable launch vehicles. Examples are given for typical spacecraft
structures, such as solar arrays, tethers, and spacecraft bus, and their comprising
materials. Mechanisms such as antenna pointing devices, solar array deployment drives,
and reaction wheels are also presented. Two examples are given for the many
expendable launch vehicles available, as well as examples of typical sizes and shapes for
49
the payloads. The following chapter summarizes the entire report, giving conclusions
drawn about launch vehicles, spacecraft structures, and spacecraft mechanisms, and how
they interact with the overall system design.
50
4 Summary and Conclusions
This chapter summarizes the information presented in this report and draws
conclusions regarding the structures, mechanisms, and launch vehicles associated with a
spacecraft. Conclusions involving modeling, analysis and procurement of various
components are stated. This chapter also recommends topics needing further research
before application.
4.1 Structures
Spacecraft structures include the spacecraft bus, solar arrays, propulsion modules,
and equipment support and covers. Each part of the structure must bear launch loads and
vibrations and tolerate the space environment for the mission lifetime. All subsystems
interact with the structure in some way. This section summarizes the most important
concepts and considerations involved in designing and building a spacecraft’s structural
components.
Structural modeling and verification is performed to predict how a spacecraft
structure will react under thermal and environmental loadings during launch and
throughout its lifetime. Some important modeling concepts include principles of statics,
dynamics, mechanics of materials, and properties of flexible bodies. Design criteria
assumptions are made to approximate structural reliability, which is never fully defined
due to material flaws and environmental load uncertainties. Environmental testing
verifies and reinforces calculations made during the modeling process.
51
Translation and rotation of a structure is modeled along and about all three axes.
External forces, including reaction forces and applied forces, must balance to keep the
body in static equilibrium. External forces must balance internal forces to keep the body
in dynamic equilibrium. Since the spacecraft structure is by function a load-bearing
structure, its stresses and strains are modeled to verify structural strength and quality.
Avoiding loads that cause plastic deformations is imperative to the success of a
spacecraft structure.
Dynamic loading imposes random vibrations on the spacecraft structure. Flexible
structures fail only at low vibration frequencies. Frequencies of concern depend on the
structure’s size and shape, and on the environmental forces. Modes of vibration are of
concern because at the mode natural frequencies the structure experiences its highest
displacement amplitudes, which impose high stresses on the spacecraft structure.
Material selection plays a vital role in the total cost, weight, and lifetime of the
spacecraft. Materials chosen for the structure of a spacecraft often reflect thermal and
environmental concerns of the mission. Both raw metals and metal alloys are used in
structural applications because of their high strength, toughness, and ease of
machinability. Composites are used because of their stiffness and relatively low mass.
Shape-memory alloys are sometimes used for their resilience. Sandwich panels and
isogrid can be used for several of these types of structures when high buckling strength
relative to weight is desired. Sandwich panels are lighter than isogrid panels, but isogrid
panels are stronger than sandwich panels.
52
4.2 Mechanisms
An important requirement for all spacecraft mechanisms is the demand for
precision pointing and a long operating life. In addition, they must withstand the launch
and vibration tests. The mechanisms must also operate in the space environment where
the thermal vacuum will influence the selection of materials, lubricants, and coatings.
High-cyclic mechanisms require frequent or constant manipulation. Some
examples of these are antenna pointing and tracking mechanisms, solar array drives, and
attitude control reaction wheels. Functional requirements for the mechanisms derive
from mission requirements and divide into torques, forces, and operating rates.
A spacecraft relies on low-cycle mechanisms for restraining and deploying
various appendages and for separating the spacecraft from the launch vehicle. These
mechanisms must be reliable and able to withstand the loads and vibrations imposed
during launch. Many different low-cycle mechanisms exist, and each design is usually
best suited to a specific application.
4.3 Launch Vehicles
Selecting the proper launch vehicle for a space mission is critical to the successful
completion of mission goals. Fortunately, launch vehicle selection is a straightforward
task that simply requires evaluating available vehicle options versus payload system
requirements. In particular, the selected vehicle must accommodate the physical
dimensions of the payload and meet minimum requirements for lift capacity, while
effecting the best balance of cost, reliability, launch location, and availability. Table 3
presents this data for twenty-one families of expendable launch vehicles from around the
world, plus the Unites States’ Space Shuttle. Many of these vehicle families have over
53
one hundred successful launches to their credit, and have proven themselves as reliable
ways to get to space.
Two US launch vehicle families, Atlas and Delta, have become standards for
launches within the United States. Together, these vehicles have logged over 300
successful launches of payloads as large as 6,500 kg, with an overall success of about
90%. The cost to launch on one of these vehicles can range from $50 to $150 million,
depending on configuration. Heavy lift capacity for the United States is provided by the
Titan family of expendable launch vehicles, or by the Space Shuttle. The Space Shuttle
is capable of lifting over 25,000 kg into low Earth orbit (LEO).
If political issues do not limit launch location to the United States, several other
vehicles are available with similar legacy to the Atlas and Titan vehicles. The European
Space Agency’s Ariane-4 vehicles have over 93 successful launches to its credit, with a
success rate of about 94%. Ariane-4 is capable of lifting 9,600 kg into LEO. Russian
launch vehicles such as the Kosmos, Proton, Tsyklon, and Soyuz have long histories of
success, with over 393 successful launches for Kosmos, 216 for Proton, 114 for Tsyklon,
and 671 for Soyuz. The Soyuz vehicle boasts a success rate of 97.5%, with a lift capacity
of about 7,000 kg into LEO.
4.4 Recommendations
The following recommendations are made regarding further research and
potential difficulties which could impact mission success.
Further research into specific types of structures, such as solar arrays and
propulsion modules, and their needs is required. The primary difficulty in structure
design is keeping mass low while maintaining strength, stiffness, adequate volume.
54
Another difficulty is keeping thermal paths open to ensure adequate thermal control.
Finally, the structure must be designed to accommodate all other subsystems’ needs,
including volume needs, thermal needs, and conductivity needs. When considering
materials or structural configurations, their heritage in space must be examined to learn
from others’ mistakes and to learn what best fulfills mission requirements.
The internet was primarily used in researching spacecraft mechanisms. Important
information, such as mass, size, pointing accuracy, operating temperature, and provider is
available, but limited, on the Internet. Additional research on the following topics is
needed:
• Cost of spacecraft mechanisms
• Materials of mechanisms and their outgassing properties
• Recent technological advances
• Providers of spacecraft mechanisms
• Operating conditions
• Necessary power
• Environmental limitations
• Launch and vibrational limitations
Accurate cost information is difficult to obtain, mainly because companies are
hard to contact and unwilling to provide pricing information.
Launch vehicle technology is constantly changing, as new vehicles are developed
and older vehicles are upgraded and improved. Successful and failed launches in the
news alter vehicles’ records almost daily. For these reasons, it is difficult to maintain a
complete and accurate database of currently available vehicles and their abilities. It is
55
recommended that system designers use Table 3 as a guide in selecting an appropriate
launch vehicle family, and then pursue detailed information and vehicle specifications
from the manufacturer. Payload faring size, boost capacity, cost, and launch location are
subject to change. Therefore, ongoing research must be applied to the expansion and
maintenance of Table 3, so that it may continue to be a useful tool for launch vehicle
selection.
56
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