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

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

iii

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

iv

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

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

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

8

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

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 -

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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