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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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
ii
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
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
v
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
vi
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
vii
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.
4
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.
5
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
6
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
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
11
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
12
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
13
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
14
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)
19
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 -
-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,
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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
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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
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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.
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Figure 9: Solar array retention/deployment mechanism23