Reusable Launch Vehical - Group 1

Transcription

MULTI-DISCIPLINARY DESIGN PROJECT
AUTUMN 2012
RE-USABLE LAUNCH AND PAYLOAD
DELIVERY SYSTEM
-FINAL REPORTFACULTY OF ENGINEERING & PHYSICAL SCIENCES
GROUP ONE:
JAMES ROPER
CHARLES OFOSU
RICHARD FIELDS
SAMUEL VEREYCKEN
EMILY ANN CARTER
NORMAN TANG FAI NG
ACADEMIC SUPERVISORS:
DR. ANDREW RAE
DR. PHIL HANCOCK
DR. YU LIU
Abstract
The report has concluded that, using new technologies, an otherwise conventional rocket design is
the most cost effective solution to achieving reusability. This proposal is made on the basis of
multiple concept appraisals, from which the most promising were taken for further analysis. Such
appraisals are made on the basis of financial, technical, safety, environmental and developmental
timescale compatibility. Integrated reviews were performed in order to make informed and
collective decisions. The final product is predicted to take seven years to develop and will achieve
a cost/kg of $1000, whilst providing a benign launch environment for both crewed and un-crewed
payloads.
I
1.
EXECUTIVE SUMMARY
The proposed launch vehicle offers greater than 92.5% reusability (by dry mass, excluding
payload). The stated goal of $1000/kg of payload is met, outdoing competition from Falcon 9
Heavy ($2340 per kilogram). This will allow financial break-even to occur in year 15, and $2.8b
net profit realisation by year 28. Confidence in these figures has been obtained by sensitivity
analysis, which has shown an 80% fall in predicted reusability to be an acceptable loss.
Developmental costs have been minimised; in comparison to the Skylon, which also offers
$1000/kg with at a cost of $12b, this proposal will cost only $1.8b and carries significantly
reduced technical risks. In addition, development, design and deployment is achieved in 7 years.
PROCESS
Areas considered include launch philosophy and propulsion, payload design, structural design,
operations, recovery methods and the applicable markets. Benchmarking, in order to judge the
value of the concept, was performed in relation to existing and planned spacecraft. Initial concept
appraisals were integrated and refined to satisfy financial, technical, safety, environmental and
developmental timescale compatibility. Refined engineering analysis was then performed in order
to display feasibility and financial viability and satisfy the project aims.
CAPABILITIES
The system has been designed with a 30T payload to 330km LEO. GEO is achieved by means of
a space tug, (would be developed in parallel) for delivery of three 6-tonne payloads at a cost of
$1839/kg. It will be rated for both human and mechanical payloads, limited to 4G accelerations.
This will allow for cost-effective satellite structural design as well as crew comfort.
The 30T payload will be sufficient to deliver multiple payloads per flight, satisfying the current
and future satellite-launch market also provide the opportunity for space infrastructure to be
delivered and constructed, or alternatively to give a contained, multiple-month workspace for a
human crew utilising new inflatable habitat technology. Re-usability is key to minimising the
costs, in order to make these markets available to investors and promote new ways to take
advantage of space as a resource.
REUSABILITY
The tanks are re-usable, with inspection, for the 200 flight vehicle life. The engines are reusable
for 50 flights with refurbishment at 10 flight overhauls. Modular design of both of these has
allowed for increased efficiency, as parts can be easily interchanged due to commonality of
design. A single engine-stage is used throughout each flight, which also reduces costs. Parachutemethods are used to recover these items, as they are inexpensive and reliable. Re-usable heat
II
shielding has been achieved by means of the PICA-X material, allowing 13-flight re-usability of
the crewed, 6-person module and 25-flights for the engine recovery system.
ENVIRONMENTAL ASPECTS
The use of hydrogen fuel has a primary advantage over toxic solid-rocket fuels and green-house
gas producing hydrocarbons. This fuel is also secure in terms of future costs, due to its
widespread availability, unlike diminishing fossil fuel resources. Cost unity with fossil fuels will
be achieved by the year 2037.
OPERATIONS
The recovery and re-assembly of hardware achieves a four week turn-around, allowing 13
launches per year for each vehicle. It is recommended that the added complexity of immediate relaunch is not beneficial, as there will not be demand to launch with this regularity. However, this
could still be supported by using multiple vehicles. Controlled gliding re-entry of stages reduces
the range from which they must be recovered, in order to facilitate the 4-week turn around.
GROWTH
From design modularity and commonality, a rocket family may be easily and inexpensively
designed by lengthening the tanks, or removing stages for smaller payloads. 14, 30 and 70 tonne
payload capabilities are proposed. In addition, tourism has been proposed as a market and may be
viable once the system has fully matured, although at present it is too expensive for all but a few
individuals.
RECOMMENDATIONS
It is the opinion of this report that a conventional launch system design be adapted in order to
facilitate reusability, in order to give the lowest technical risks and significantly reduced
developmental costs. In order to aid reusability, new technologies including reusable heat
shielding, altitude compensating engines, parafoil recovery and a permanent space-tug for GEO
insertion are employed.
A low launch cost may drastically improve the support for space exploration, which by many is
currently considered to be too expensive and wasteful of resources. If costs can be further
reduced, it may also be a viable platform for orbital space tourism. It may also attract private
investors with interests in capitalising on space resources, such as zero gravity manufacturing and
space based power generation.
Further research is required in the areas of aerospike engine performance, thrust chamber life and
hydrogen embrittlement protection. In addition, the development of a more flexible and
operationally efficient human module is suggested for future consideration.
III
2.
NOMENCLATURE
CCDev
Commercial Crew Development
COTS
Commercial Orbital Transportation Service
CRS
Commercial Resupply Service
CST&EI
Commercial Space Transportation & Enabled Industries
CoCom
Coordinating Committee for Multilateral Export Controls
FAA
Federal Aviation Agency
GEO
Geostationary Orbit
GTO
Geostationary Transfer Orbit
GSO
Geosynchronous Orbit
GNC
Guidance Navigation and Control
H&S
Health and Safety
ISS
International Space Station
ITAR
International Traffic in Arms Regulations
LVM&SI
Launch Vehicle Manufacturing & Service Industry
LH2/Lox
Liquid Hydrogen and Liquid Oxygen
LNG
Liquid Natural Gas
LEO
Lower Earth Orbit
MEO
Medium Earth Orbit
NGSO
Non Geosynchronous Orbit
RTG
Radioisotope Thermoelectric Generator
R&D
Research and Design
RCM
Risk Control Measure
TDRS
Tracking and Data Relay Satellites
VSAT
Very Small Aperture Terminal
IV
3.
ACKNOWLEDGEMENTS
The Authors would like to thank Dr Philip Hancock, Dr Yu Liu and Dr Andrew Rae for your
continuous guidance throughout the project.
We would also like to thank the following people for their help;
Kevin Cheung (Pilot) for the information about aircraft landing
Roy Slocombe (Senior Engineer) for tunnel mining information
Benoit Geline (Sodern in France) for information on star trappers
Joseph Shoer for permission to use his graphic
Jay L Perry (NASA) whose information on Environmental control and Life support was of much
assistance.
Stuart Dalrymple (Project Manager C-Tech Innovation Ltd) for information on ohmic heaters
Marc M Cohen (Space Architect) for your insights into the study of water walls development.
Finally an extra special thanks to Dr Andrew Rae for making the journey from Scotland every
week.
V
Table of Contents
1.
Executive Summary ................................................................................................... II
2.
Nomenclature ........................................................................................................... IV
3.
Acknowledgements .................................................................................................... V
4.
Introduction ................................................................................................................ 1
5.
Initial Research ........................................................................................................... 3
5.1.
Launch Vehicle Marketing ................................................................................................... 3
5.1.1.
Orbital Launch Market ................................................................................................. 3
5.1.2.
Orbital Payload Market ................................................................................................ 4
5.1.3.
Launch Market in the Beginning of 2012 ..................................................................... 5
5.1.4.
Launch Market Forecasts.............................................................................................. 5
5.1.5.
Commercial Human Spaceflight Market ...................................................................... 7
5.1.6.
Economic Impact of Commercial Space .................................................................... 10
5.1.7.
Funding and Prizes ..................................................................................................... 12
5.1.8.
Launch Market Analysis ............................................................................................. 12
5.2.
Mechanical Payload............................................................................................................ 15
5.2.1.
Requirements .............................................................................................................. 15
5.2.2.
Mechanical Payload Selection .................................................................................... 15
5.2.3.
Attachment Methods................................................................................................... 16
5.2.4.
Multiple Satellite Considerations and Pico-Satellites................................................. 16
5.2.5.
Potential Infrastructure ............................................................................................... 17
5.3.
Human Modules ................................................................................................................. 18
5.3.1.
Introduction to Payload Structural Considerations ..................................................... 18
5.3.2.
Pressure Vessel Consideration.................................................................................... 19
5.3.3.
Seating and Layout ..................................................................................................... 19
5.3.4.
G Forces...................................................................................................................... 20
5.3.5.
Vibration and Damping .............................................................................................. 22
5.3.6.
Environmental control and Life Support Systems ...................................................... 22
5.3.7.
Air Revitalisation........................................................................................................ 23
5.3.8.
Regenerative Methods ................................................................................................ 25
5.3.9.
Carbon Dioxide Removal ........................................................................................... 26
5.3.10.
Water and Waste......................................................................................................... 26
5.3.11.
Trace Contaminant Control/Filtering ......................................................................... 27
5.3.12.
Pressure Management ................................................................................................. 27
VI
5.3.13.
Nutrition ..................................................................................................................... 28
5.3.14.
Fire Prevention ........................................................................................................... 29
5.3.15.
Rendez Vous/Docking ................................................................................................ 29
5.3.16.
Radiation..................................................................................................................... 30
5.3.17.
Future Proofing/Innovation ........................................................................................ 32
5.3.18.
Capsule Reusability .................................................................................................... 33
5.4.
Fuel Options ....................................................................................................................... 35
5.4.1.
Liquid Chemical Propellants ...................................................................................... 35
5.4.2.
Solid and Hybrid Propellants...................................................................................... 36
5.5.
Launch Philosophy ............................................................................................................. 37
5.5.1.
Space-Plane Configurations ....................................................................................... 37
5.5.2.
Enabling Technology: Air-Breathing Engines ........................................................... 38
5.5.3.
Vertical Launch Assessment For Lace ....................................................................... 39
5.5.4.
Proposal for Space Plane Configuration ..................................................................... 40
5.5.5.
Space Plane Conclusions ............................................................................................ 40
5.5.6.
Rocket Configurations ................................................................................................ 41
5.5.7.
Enabling Technology: Staging ................................................................................... 42
5.5.8.
Nozzle Configurations ................................................................................................ 43
5.6.
Launch Concepts ................................................................................................................ 46
5.6.1.
Exotic Propulsion ....................................................................................................... 46
5.6.2.
Electric Propulsion ..................................................................................................... 46
5.6.3.
Space Tethers.............................................................................................................. 47
5.6.4.
Rotovator .................................................................................................................... 48
5.6.5.
Space Elevator ............................................................................................................ 48
5.6.6.
Maglev Space Transportation ..................................................................................... 49
5.6.7.
Solar Sails ................................................................................................................... 50
5.6.8.
Technology Review .................................................................................................... 51
5.6.9.
Ramp Launch System Design..................................................................................... 53
5.6.10.
Mountain Tunnel Launch System Design .................................................................. 54
5.7.
Launch Vehicle Hardware .................................................................................................. 55
5.7.1.
Propellant Management .............................................................................................. 55
5.7.2.
Guidance, Navigation and Control ............................................................................. 56
5.7.3.
Communication and Data Handling ........................................................................... 57
5.7.4.
Electrical Power.......................................................................................................... 57
5.8.
Launch Abort Systems ....................................................................................................... 59
5.8.1.
LAS............................................................................................................................. 59
5.8.2.
SpaceX LAS ............................................................................................................... 59
VII
5.9.
Spacecraft Hardware .......................................................................................................... 60
5.9.1.
Attitude and Orbit Control .......................................................................................... 60
5.9.2.
Electrical Power.......................................................................................................... 60
5.10.
Heat Shielding ................................................................................................................ 62
5.10.1.
Deceleration ................................................................................................................ 62
5.10.2.
Heat Shield Design ..................................................................................................... 64
5.11.
Re-Entry Trajectory ........................................................................................................ 66
5.11.1.
Ballistic Entry Trajectory ........................................................................................... 66
5.11.2.
Glide Trajectory.......................................................................................................... 67
5.11.3.
Skip Re-Entry ............................................................................................................. 68
5.12.
Environmental Impact - Propulsion System ................................................................... 69
5.12.1.
Petroleum.................................................................................................................... 69
5.12.2.
Hybrid ......................................................................................................................... 70
5.12.3.
Solid............................................................................................................................ 70
5.12.4.
Cryogenic ................................................................................................................... 71
5.13.
Environmental Impact – Space Debris ........................................................................... 72
5.13.1.
Risk of Impact ............................................................................................................ 72
5.13.2.
Reducing Debris Pollution.......................................................................................... 73
5.14.
Conclusion ...................................................................................................................... 74
5.14.1.
Summary of Research Areas ...................................................................................... 74
5.14.2.
Down-Selections......................................................................................................... 74
6.
Initial Design Phase .................................................................................................. 77
6.1.
Vehicle Mass Estimations .................................................................................................. 77
6.1.1.
First Estimate of Total Delta-V Required................................................................... 77
6.1.2.
Estimation of Fuel Mass ............................................................................................. 77
6.1.3.
Initial Tank Sizing ...................................................................................................... 80
6.2.
Layout Options and Configuration ..................................................................................... 80
6.2.1.
Feeding Packet............................................................................................................ 81
6.2.2.
Carrying Packet .......................................................................................................... 81
6.2.3.
Feeding Packet - Donut Tank ..................................................................................... 82
6.3.
Operation ............................................................................................................................ 82
6.3.1.
Mission Planning ........................................................................................................ 82
6.3.2.
Ground Operations ..................................................................................................... 83
6.3.3.
Ground Operations Architecture ................................................................................ 84
6.3.4.
Vehicle Elements Retrieval ........................................................................................ 85
6.4.
Stage Recovery ................................................................................................................... 86
VIII
6.4.1.
1st Stage Recovery Options......................................................................................... 86
6.4.2.
Parachute .................................................................................................................... 86
6.4.3.
Parafoil ....................................................................................................................... 87
6.4.4.
Wings.......................................................................................................................... 88
6.4.5.
Powered Return .......................................................................................................... 89
6.4.6.
1st Stage Recovery Conclusion ................................................................................... 90
6.4.7.
2nd Stage Recovery Options ........................................................................................ 90
6.4.8.
3rd Stage Recovery Options ........................................................................................ 90
6.5.
Conclusion .......................................................................................................................... 92
6.5.1.
7.
Down-Selections......................................................................................................... 92
Finalised Design Phase ............................................................................................. 94
7.1.
Second Iteration of Mass Estimation .................................................................................. 94
7.1.1.
Engine Mass Prediction .............................................................................................. 94
7.1.2.
Final Re-Sizing ........................................................................................................... 95
7.1.3.
Engine Design for Re-Usability. ................................................................................ 96
7.2.
Financial Analysis for The Propulsion System .................................................................. 98
7.3.
Launch Trajectory ............................................................................................................ 100
7.4.
Fuel Tank Design ............................................................................................................. 100
7.4.1.
Material Choice ........................................................................................................ 101
7.4.2.
Thermal Protection System ...................................................................................... 102
7.4.3.
Structural Design and Analysis ................................................................................ 103
7.4.4.
Rigidity Sizing .......................................................................................................... 104
7.4.5.
Applied and Equivalent Axial Loads........................................................................ 105
7.4.6.
Tensile Strength Sizing............................................................................................. 106
7.4.7.
Sizing for Stability .................................................................................................... 106
7.4.8.
Mass Calculation ...................................................................................................... 107
7.4.9.
Common Bulkhead ................................................................................................... 108
7.5.
Layout Design .................................................................................................................. 109
7.6.
In-Orbit Trajectories and Propulsion ................................................................................ 110
7.6.1.
Orbit Definitions ....................................................................................................... 110
Polar Orbits................................................................................................................................... 110
Sun Synchronous Orbits ............................................................................................................... 110
Geosynchronous Orbits ................................................................................................................ 110
Molniya Orbits ............................................................................................................................. 111
7.6.2.
Orbital Elements ....................................................................................................... 112
7.6.3.
Orbital Mechanics .................................................................................................... 112
IX
7.7.
Powered Manoeuvres ....................................................................................................... 113
7.7.1.
Hohmann Transfer .................................................................................................... 113
7.7.2.
Plane Change Manoeuvres ....................................................................................... 114
7.7.3.
Combining Hohmann Transfer with Plane Change .................................................. 115
7.8.
GTO Delivery ................................................................................................................... 115
7.9.
Space Tug ......................................................................................................................... 116
7.10.
Re-entry Philosophy ..................................................................................................... 119
7.10.1.
Atmospheric Density ................................................................................................ 119
7.10.2.
Parachutes ................................................................................................................. 120
7.10.3.
Vehicle Shape Choice............................................................................................... 121
7.10.4.
Heat shield material analysis .................................................................................... 122
7.10.5.
Heat Shield Material Choice and Study ................................................................... 126
7.10.6.
Simulation Results and Conclusion .......................................................................... 127
7.10.7.
Heat Shield for The Engine ...................................................................................... 128
7.11.
Costs Associated with The Heat Shields ...................................................................... 128
7.12.
Infrastructure ................................................................................................................ 131
7.12.1.
Location Selection .................................................................................................... 131
7.12.2.
Construction ............................................................................................................. 131
7.12.3.
Infrastructure Requirements ..................................................................................... 132
7.12.4.
Storing and Transportation of Hydrogen .................................................................. 133
7.12.5.
Future Proofing ......................................................................................................... 134
7.12.6.
Conclusion ................................................................................................................ 135
7.13.
Financial Analysis ........................................................................................................ 135
7.13.1.
30 Tonnes Mechanical Payload ................................................................................ 136
7.13.2.
Human Modules ....................................................................................................... 136
7.13.3.
Sensitivity Analysis .................................................................................................. 137
7.13.4.
Conclusion ................................................................................................................ 138
8.
Risk Assessment ...................................................................................................... 139
9.
Technical Review .................................................................................................... 141
10.
Conclusion ........................................................................................................... 145
11.
Reference ............................................................................................................. 147
12.
Appendix .............................................................................................................. 163
12.1.
Marketing ..................................................................................................................... 163
12.1.1.
2011 Commercial Launch Events............................................................................. 163
12.1.2.
2011 Non-Commercial Launch Events .................................................................... 164
X
12.1.3.
Orbital Classification ................................................................................................ 164
12.1.4.
Payload Weight Classification.................................................................................. 165
12.1.5.
Payload Usage Classification ................................................................................... 165
12.1.6.
Economy Impact by Industries ................................................................................. 166
12.2.
Ramp Design ................................................................................................................ 167
12.3.
Existing tunnels ............................................................................................................ 170
12.4.
Launch Vehicle System Comparison ........................................................................... 171
12.5.
Antenna ........................................................................................................................ 172
12.6.
Infrastructure of Kennedy Space Center ...................................................................... 173
12.6.1.
The Location of the Infrastructure ............................................................................ 173
12.6.2.
Detail Description of The Facilities.......................................................................... 174
12.7.
Financial Analysis ........................................................................................................ 178
12.7.1.
Price of The Systems ................................................................................................ 178
12.7.2.
Timeline Costing ...................................................................................................... 183
12.7.3.
Balance Sheet ........................................................................................................... 184
12.7.4.
Income Statement ..................................................................................................... 160
12.7.5.
30 Tonnes Mechanical Payload ................................................................................ 160
12.7.6.
Human Modules ....................................................................................................... 161
12.7.7.
Sensitivity Analysis .................................................................................................. 162
12.8.
Risk Assessment ........................................................................................................... 166
12.8.1.
Quantify Risks .......................................................................................................... 166
12.8.2.
Application for NASA NAFCOM ............................................................................ 168
12.1.
12.1.1.
12.2.
Mass considerations/dimensions and calculations ....................................................... 169
Satellite dimensions and masses ............................................................................... 169
Human Payload ............................................................................................................ 170
12.2.1.
Thirty person human adaptation of shuttle cargo ..................................................... 170
12.2.2.
Capsule mass and calculations ................................................................................. 171
12.2.1.
Mass of human module and life-support .................................................................. 173
12.3.
Further Propellant Information. .................................................................................... 177
12.3.1.
Monopropellants ....................................................................................................... 177
12.3.2.
Liquid bi-propellants: oxidisers ................................................................................ 177
12.3.3.
Liquid bi-propellants: expansion on rp-1, lh2 and lng ............................................. 178
12.3.4.
Use of metals as fuels ............................................................................................... 179
12.3.5.
Solid and hybrid propellants ..................................................................................... 180
12.3.6.
Inert Propellants........................................................................................................ 181
12.4.
Rocket Sizing Spreadsheet ........................................................................................... 183
12.5.
MATLAB Program for Launch Trajectory. ................................................................. 185
XI
12.5.1.
Code for The Matlab Program. ................................................................................. 188
12.6.
Tank design .................................................................................................................. 195
12.7.
Re-entry Simulation Calculations ................................................................................. 196
12.8.
Project Management ..................................................................................................... 217
12.8.1.
Initial Gantt-chart ..................................................................................................... 217
12.8.2.
Final Gantt-chart ....................................................................................................... 218
12.8.3.
Deadline and Meeting Calendar ............................................................................... 219
XII
4.
INTRODUCTION
The purposes of this report are to fully and comprehensively analyse the current climate, market
and technological advancement with regards to space travel. There is particular focus on the
reusable launch and payload delivery aspect to provide a thoroughly considered proposal
programme, with the intention of advising potential investors of the potential capital reward from
investing, and the timeframe for that investment return. The potential for this programme is also
to work towards the on-going desire to make Space more and more accessible, reducing the
boundary that has kept Space out of the reach of the general public.
Since the first launch into Space over sixty years ago, the possible applications have boomed and
the world has progressively harnessed more and more of this extra-terrestrial resource. From
satellites coursing around the atmosphere, to probes darting out for planets and bodies that have
intrigued us for centuries, the open reaches of Space are used for multitudes of reasons, and where
there has been need, companies and organisations have facilitated. The Space industry has seen
many different projects and programmes, providing methods for accessing Space to those who
have the funding to provide. The technologies have evolved in this respect over half a century,
and this has provided different means to reach Space. The field of re-usable launch and payload
delivery systems has not been developed to a great extent thus far, and with growing interest in
Space and the opportunities it offers, the appeal of such a system may proffer a good incentive to
organisations interested in the potential rewards. In the light of the retirement of the NASA Space
Shuttle in 2011, the market for a reusable launch and payload delivery system has been ill-served
by the offering of current systems. This creates an opportunity for new developing programmes to
advance and claim a portion of the market.
Despite reusable launch and payload delivery not being an original field, the opportunity for new
investment is worthy of analysis, as it is a service field more than a product field, and this entails
an open-ended plan regarding the future of the market, and the contracts available on that market.
As a field, it fluctuates like any other, and as a market it eludes the attentive focus of general
consumers. It is generally a sector that has remained commercial only in what it offers companies
and the public. In other words, the field has thus far only aspired to provide launch methods for
commercial products and scientific research. However, how far does the horizon stretch? How
does the future of space travel look and how far off does is that image? The market exists, and has
grown significantly since the American/Russian bluster “Space-race”, but what requires analysis is the actual commercial opportunity that exists, and whether it is a viable investment that will
make a worthwhile return. The concerns of today (economy, environment, etc.) all contribute to
Author: Group
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the potential risks involved with any new project, and without due assessment they can cause
costly delay and even ruin a project, so these are duly examined as well.
The report starts by identifying and explaining the key concepts involved, and analyses the
current market, and the potential markets of the future. One of the key goals was to attain a
preferable level of reusability, as this determines to a great extent the success of the programme.
The report lays out the process that was followed for the research that was conducted. This
research turned up a number of technologies and concepts, which were then rated and reduced
down to a selection chosen for the capability and practicality. Through further processes of
analysis, the many different combinations were refined to the most advisable solution that could
be determined from the research done. This solution has been fully analysed for its market
competitiveness, its commercial suitability, and this report finishes with a technical review of the
solution, and a conclusion determined from this study. This report therefore provides a detailed
plan, backed up by research and calculations, for a reusable launch and payload delivery system
with a 10 year development period and the process by which this plan was reached.
Author: Group
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RE-USABLE LAUNCH AND PAYLOAD DELIVERY SYSTEM - AUTUMN 2012
5.
INITIAL RESEARCH
INITIAL RESEARCH
5.1.
LAUNCH VEHICLE MARKETING
5.1.1.
ORBITAL LAUNCH MARKET
Orbital launch companies categorise the delivery of the payload into two orbits, the commercial
geosynchronous orbit (GSO), and non-geosynchronous orbit (GSO) as defined in Appendix
12.1.3. The industry classifies the payloads by their weight, which is defined in Appendix 12.1.4.
The satellite manufacturer will design and set price according to these classifications.
Table 1 – 2011 Worldwide orbital launches (FEDERAL AVIATION ADMINISTRATION, 2012)
Commercial
Launches
United States
Russia
Europe
China
Japan
India
Iran
Multinational
Total
0
10
4
2
0
0
0
2
18
NonCommercial
Launches
18
21
3
17
3
3
1
0
66
Total
Launches
18
31
7
19
3
3
1
2
84
In 2011 there were 84 worldwide orbital launches. 79% of the launches were non-commercial and
21% were commercial, as shown in Table 1. This is a decrease from 2010 where 31% of launches
worldwide were commercial. One possible reason for the decrease is the deferral of 4 commercial
launch projects which were the Commercial Resupply Service (CRS) and the Commercial Orbital
Transportation Service (COTS) by NASA that was originally planned for 2011 (KREMER, Ken,
2012).
US $ Millions
1000
$ 880
800
$ 707
2
600
400
200
United States
2
$0
$ 140
$ 200
Russia
10
4
China
Multinational
0
Commercial Launch Revenues
Europe
Number of Commercial Launches
Figure 1 – International commercial launch market (FEDERAL AVIATION ADMINISTRATION, 2012)
The commercial launch industry produced approximately $1.92B in revenue in 2011. Europe had
the highest revenue of $880M with only 4 launches, which were conducted by ESA. The Russian
Author: Norman Tang Fai Ng
| Page 3 of 262
INITIAL RESEARCH
commercial launch revenue was $707M with 10 launches, followed by 2 launches each from
China and multinational co-operations, as shown in Figure 1.
The launch service corporation of Sea Launch AG is classified as multinational and produced
revenue of $200M. China had the least revenues in comparison with a total of $140M by its Long
March 3B vehicles. Figure 1 shows the market share of each nation in the commercial launch
market. Each launch contained different payloads that were delivered to selected orbits, as shown
in Appendix 12.1.1, Figure 82.
5.1.2.
ORBITAL PAYLOAD MARKET
Table 2 – 2011 Payloads launched worldwide (FEDERAL AVIATION ADMINISTRATION, 2012)
NonCommercial
Payloads
28
32
9
17
3
8
1
0
98
Commercial
Payloads
United States
Russia
Europe
China
Japan
India
Iran
Multinational
Total
0
21
8
4
0
0
0
2
35
Total
Payloads
28
53
17
21
3
8
1
2
133
There were 18 commercial launches in 2011, which carried 41 payloads into orbit containing both
commercial and non-commercial satellites, as shown in Table 2. Eight government payloads were
launched commercially, including three remote sensing, two communications, two sciences and
one development. The remaining 33 payloads were commercial communication satellites.
Appendix 12.1.5 illustrates and defines the payload classification.
10
Commercial
Communications
2
Government
Civil
34
46
Government
Military
Non-Profit
Figure 2 – Payload type delivery by non-commercial launches (FEDERAL AVIATION ADMINISTRATION,
2012)
The remaining 90 payloads were delivered by 66 non-commercial launches, as shown in
Appendix 12.1.1, Figure 82. The 90 payloads were for civil government, military and non-profit,
as shown in Figure 2. In Appendix 12.1.2, Figure 83 shows the detail breakdown of launches.
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There were 21 satellite payloads delivered to Geostationary Earth Orbit (GEO) and the rest were
sent to NGSO through commercial launches. However, Russia failed to deliver one payload to
GEO, as shown in Appendix 12.1.1, Figure 82. In addition, there were five non-commercial
launch failures.
5.1.3.
LAUNCH MARKET IN THE BEGINNING OF 2012
Micro
12
13
Small
5
Medium
19
17
Intermediate
Large
8
10
Heavy
Unknown
Payload Mass in First half of year 2012
Figure 3 – The delivered payload mass in first half of year 2012 (FEDERAL AVIATION ADMINISTRATION,
2012)
In the first half of 2012, 55 launches were carried out in which 18 of them were commercial
launches. These launches have generated total revenue of $1.90B, which is 98% of the whole sum
of 2011. Even though 13 of the payloads are of an unknown mass classification, the medium and
above classes still accounts for roughly 50% of the payload mass, as shown in Figure 3.
5.1.4.
LAUNCH MARKET FORECASTS
Figure 4 – The historical and forecasts of GSO and NGSO Launches (AST and COMSTAC, 2012) (FEDERAL
AVIATION ADMINISTRATION, 2012)
The Federal Aviation Administrations’ Office of Commercial Space Transportation (FAA/AST) and the Commercial Space Transportation Advisory Committee (COMSTAC) use the previous
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and present launch data and the information given by the industries to predict the future forecast,
as shown in Figure 4. The following organisations responded to participate in this forecast:

Arabsar

Arianespace

Boeing

Hispasat

Intelsat

Loral

Sea Launch

Sirius XM

SpaceX

Tenor

TerreScar

Figure 5 – The payload forecast of GSO and NGSO launch (FEDERAL AVIATION
ADMINISTRATION, 2012)
The global demand of launch in both GSO and NGSO is expected to continue its growth in the
future, as shown in Figure 4. For the GSO market, the 2012 forecast predicts an average of 16.3
launches and 21.2 payloads per year, an increase of 0.7 payloads per year from the 2011 forecast.
Moreover, the forecast predicts 43% higher mass class payloads will be launched between the
year of 2012 and 2021, as shown in Figure 5. The forecast also shows an increasing payload size
in the future.
In addition, there may be a further increase in satellite launches as dual-manifesting vehicle
technology is becoming more mature, for example SpaceX announced that their vehicles also
have the capability to implement such technology (SPACEX, 2008). Dual-manifesting technology
allows the launching two or more satellites in one vehicle. Furthermore, countries like China and
Russian are currently launching payload on a regular basis with increasing numbers of launches
every year.
The NGSO market is expecting12.8 launches per year between 2012 and 2021, as shown in
Figure 5. An average of 12 medium-to-heavy class vehicle launches per year is expected. The
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remaining launches will be small class vehicles. The industry will be focusing on the commercial
cargo and crew transportation services, which is expected to dominate 50% of the launches within
the next ten years (AST and COMSTAC, 2012). Currently, technical or financial issues are the
only causes for deferring the ISS resupply launches. However, SpaceX has recently carried out a
resupply mission to ISS in 2012 as the first commercial resupply launch (STAFF, Wire, 2012).
Commercial telecommunication is expected to increase its contribution to the NGSO market until
2017. This is because companies such as the Globalstar (ARIANESPACE, 2011), ORBCOMM
(BUSINESS WIRE NEW RELEASES, 2008) and Iridium (IRIDIUM, 2012) are planning to
replace their constellation satellite. Subsequently, the demand of commercial telecommunication
is expected to drop. The forecast also predicts that other areas of the NGSO market, such as
commercial remote sensing and science and engineering will remain stable (AST and
COMSTAC, 2012).
5.1.5.
COMMERCIAL HUMAN SPACEFLIGHT MARKET
SUBORBITAL FLIGHT
The suborbital reusable vehicle (SRV) is an emerging market in the spaceflight industry. SRVs
are designed to deliver high flight rates and relatively low costs for carrying humans or cargo to
space. The Tauri Group had carry out a forecast in this market using data based on high net worth
individuals, researchers, governmental intention and the market capability.
The study shows that around 8,000 high net worth individuals would be interested and have the
intention to purchase a suborbital flight and this group is expected to grow 2% annually.
Furthermore, 3,600 individuals are expected to fly within the 10 year forecast period. Other
people than the high net worth individuals are expected to create an additional 400 participants,
which increase the total number to 4,000. 335 seats are expected to be purchased in the first year
and sequentially increase to 400 seats by year 10. This total may increase to 11,000 in the case of
rapid growth. This could be caused by breakthrough in technology or increase in market demands.
However, interruption such as the impact of the global economy could alternate the market with a
reducing of 2,000 of seats over 10 years. Corporates, contests and promotion or even space
personal training could also affect the demand the market. Excalibur Almaz is a company
currently developing lunar transport system for 3 seats with $150M per seat. The company has
taken advantage of the technology and is planning to train with suborbital flights as early as 2015
(TAURI GROUP, 2012).
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Figure 6 – SRV reservation to date (TAURI GROUP, 2012)
Only 8 people in the history of orbital space flight have personally paid to do so. The tickets cost
from $20M to $35M. 6,000 people have flown in parabolic flight with Zero-G since year 2004; 40
MiG fighter jet flown annually with Edge of Space and 115 future tourists were trained with
Nation Aerospace Training Research (NASTAR) Centre (TAURI GROUP, 2012). These all show
the interest of the public for tour to space. Moreover, there are 925 individuals who have reserved
tickets for suborbital flight, as shown in Figure 6. The highest price is $200,000 with Virgin
Galactic and is expecting to launch in 2013 (DAVID, Leonard, 2012).
ORBITAL SPACE TOURISM
Is there a desire or willingness for space tourism at present? Research was done into the level of
feasibility for space tourism at current prices. Sub-orbital flights will soon open up the market to a
wider range of people interested in space tourism, but who perhaps may not be able to afford the
grand costs of Orbital flight. Space Adventures is continually seeking new ways to get private
citizens to space, and recently entered an agreement with Boeing to market seats on their new
spacecraft, the CST-100, which is expected to be operational in 2016/7. (SPACE
ADVENTURES, 2012)
Space Adventures have already managed to carry about seven people aboard the Russian Soyuz,
with one person travelling twice. However these trips could hardly be referred to as tourist trips,
as most if not all of them also conducted experiments in space taking advantage of microgravity.
The price for civilian cosmonauts to orbit has been $20M and there have been very few
participants; as said before most of which also conducted experiments on board the ISS and spend
on average 8 days in space. According to Space Adventures the Russian Soyuz is the only viable
way of carrying humans to orbit, and will remain that way for the next few years. With the
decommissioning of the space Shuttle the USA are now paying significantly more than previous
years to charter seats on the Soyuz to transport their astronauts to space, as shown in Figure 7.
Author: Norman Tang Fai Ng / Emily Ann Carter
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Figure 7 - Yearly Cost (per seat) for U.S. Crew Transportation Services aboard the Soyuz Vehicle for launches
through 2015
One can see that the price for travel is significantly rising which suggests the need for more
manned spacecraft is not a small one. This could be a potential avenue for investments if this new
reusable system can provide contracts to other space agencies and carry some of their cargo for a
price and even perhaps a more reasonable one.
As for tourism: to investigate this aspect a number of surveys have been carried out by various
organisations to answer this question. One such reliable study was conducted by Futron/ Zogby.
Futron is a well-established Corporation known for providing Innovative Decision Management
Solutions, performance and results. (FUTRON CORPORATION, 1999-2010) Zogby also is
known for consultancy and research provision to companies for the purpose of decision making.
(IBOPE INTELIGÊNCIA, 2011)
The survey was conducted on 450 USA participants and they were asked a number of questions to
establish the size of the market, the potential growth of the market and the customer
characteristics.
On review of this study it would appear that for orbital space tourism to take off the demand for it
would have to be increased to the thousands and the price would have to be significantly lowered
in order for more of the population to take part. At current prices according to the Futron/Zogby
survey (FUTRON CORPORATION, 2002) only 30% of the (450) respondents stated that they
would be willing to pay between $1M and $25M, whereas 70% of respondents were not willing to
pay based on the mentioned price range, as shown in Figure 8.
Author: Emily Ann Carter
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Figure 8 - Showing the respondent's willingness to pay for spaceflight within stated amounts (FUTRON
CORPORATION, 1999-2010)
This suggests the great need to cut costs and that at present it is infeasible or at best difficult to
achieve amortization and return on investments if this were the sole function of the launch system.
To qualify for this survey participants had to have a yearly income of or above $250,000 or a net
worth of $1M as these would be the only people who could actually afford a price this high in a
reasonable amount of time. With prices at $20M the passenger would have to be classified as an
Ultra Net-Worth Individual as was the case for Charles Simonyi and Guy Liberte to name a few
of the previous space travellers (who both have a net worth equal to or well above $1B as verified
by (FORBES LLC, 2012)).
That being said if a vehicle can be developed so that ticket prices can be reduced, tourism on its
own may not be able to amortise the cost of R&D (ABITZSCH, S and Eilingsfeld, F, 1992)
needed for this new vehicle and thus other potential markets should be able to be satisfied by this
re-usable system for the best return on investment to occur, thus explaining why this paper later
considers mechanical payloads among other markets.
5.1.6.
ECONOMIC IMPACT OF COMMERCIAL SPACE
Table 3 – Economic impact cause by the CST&EI (ADMINISTRATION, FEDERAL AVIATION, 2010)
Total Impact
Commercial
Launch
Economic
Activity
($000)
Earning
($000)
Jobs
1999
2002
2004
2006
2009
36
24
17
21
24
61,313,711
95,025,746
98,086,960
139,262,027
208,329,012
16,431,192
23,527,745
25,045,888
35,659,935
53,257,346
497,350
576,450
551,350
736,130
1,029,440
Author: Emily Ann Carter /Norman Tang Fai Ng
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The FAA released the most recent study on the economic impact cause by the commercial space
from United States. The Commercial Space Transportation and Enabled Industries (CST&EI)
have created a total of $208.3B in economic activity, more than one million employment positions
and earning exceeding $53B, as shown in Table 3.
Table 4 - Economic impact cause by the LVM&SI (ADMINISTRATION, FEDERAL AVIATION, 2010)
Total Impact
Economic
Activity ($000)
Earning ($000)
Jobs
1999
2002
2004
2006
2009
3,515,978
791,759
1,658,384
1,166,723
827,817
1,071,722
206,328
437,674
308,087
218,595
28,617
4,828
8,870
5,690
3,820
The Launch Vehicle Manufacturing and Service Industry (LVM&SI) are parts of the CST&EI,
which are indicated in the launch industries. This includes the USA industries related to the
manufacturing and processing of orbital and suborbital launch vehicles and its payload into space.
$828M in economic activity was recorded in 2009 by LVM&SI, creating about 4,000 jobs with
total earning exceeding $219M, as shown in Table 4.
Figure 9 – Total economic activity impacts on the U.S. economy (ADMINISTRATION, FEDERAL AVIATION,
2010)
The CST&ET has not been growing uniformly across all sectors, as shown in Figure 9. In recent
year, the economic impact from launch vehicle manufacturing and satellite manufacturing has
been declining. On the other hand, the ground equipment manufacturing has shown a positive
growth with VSAT service, satellite data services, mobile satellite telephony and satellite remote
sensing achieving modest growth. The satellite service sector showed a rapid growth, due to the
increase of demand for HD-TV and transponder leasing. The distribution industries have
continued to gain strength, however the LVM&SI has depreciated over the years. This could
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mean the industry is manufacturing parts for export to other countries as more countries and
companies start competing with the USA.
The study shows the economic impact on the USA. This created 1,029,440 jobs over the USA in
2009 and earnings of approximately $53B. This brings an economic activity of a total of $208B.
This increases the employment rate and the secondary manufacturing industries, which create
more opportunities for the public. The definition of CST&EI and a more refined breakdown is
shown in Appendix 12.1.6, Figure 84.
5.1.7.
FUNDING AND PRIZES
The main difficulty in a commercial space programme is obtaining the funding to sustain the
project development. In the past, space projects were primarily funded by governments and did
not impact the project in the same way commercial funding does.
Currently, there are four main types of funding available to this industry. Governments tend to be
the most reliable source for funding such projects, which requires a large amount of money to
maintain the development. The Skylon project is a good example of the main funding coming
from the government. Alternatively it may be possible to find a private investor such as Aabar,
which invested $110M in Virgin Galactic and boosted the stake up to 6% from the public
(MALAS, Nour, 2011). This showed that Virgin Galactic had the capability and gained the trust
of the consumers. This will allow contracts to be in place before launches. This might also win
prizes from the industry. The main prizes available are NASA’s Centennial Challenges and the X PRIZE Foundation. However, the topic of the prizes change all the time, thus the project cannot
be relying on this source of funding.
5.1.8.
LAUNCH MARKET ANALYSIS
Figure 10 – Trends in satellite mass class distribution (AST and COMSTAC, 2012)
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There is a slight increase in demand of space launches every year, as previously described in this
chapter. Although the USA postponed 4 projects till 2012, in 2011 it was still able to achieve
about a 3% increase in comparison to previous years. The data clearly shows evidence of the
payload mass classifications shifting to include larger masses over the years. In addition, the
forecast done by the FAA has also suggested an increase in demand of payload launchings and the
increase of payload mass, as shown in Figure 10. The main contribution is the replacement of the
commercial telecommunication satellites in the year of 2012 to 2017. There will be an increase in
demand of satellites below 2,500 kg by the year 2014. This is due to the increased usage of cubesatellite and micro-satellite technologies. The technology is relatively cheap in price and will be
broadly used after the cost per kilogram to orbit is further reduced.
The commercial human spaceflight industry has shown strong evidence that it can become one of
the biggest market shares apart from delivering payload. The suborbital space flight technology is
becoming more mature, as the first expected flight by Virgin Galactic would be taking place in
2013. The price per ticket is soon expected to lower as competition increases. However, the
orbital flight market is treated differently, as the price is still relatively high. Only people with
yearly income of or above $250,000 or a net worth of $1M could potentially afford the price
range. This market will be ignored temporarily by the industry and will not be the focus of the
industry in the near future.
The space industry could increase the local economic activity. The entire launch industry has been
constantly growing, as shown by CST&EI. However, the LVM&SI has shown a constant decrease
starting from 2004. The economic activity and the jobs shown had halved over the year. This
might be because of the efficiency of the LVM&SI has increased, requiring less labour force and
leading the relative industry to decrease. The information has shown how much economic activity
and employment can be generated in the country. There will be a similar effect in other countries,
even though this economic study was only carried out in USA.
Previous data and forecasting are showing clear evidence that there is a potential market space for
new launch providers to enter. Older launch systems such as the Ariane 5 and Proton M are not
designed to be launch as frequent as the newer providers such as Space X. The modularity and
simple design enables mass production and the launching of heavier masses with the similar
designs such as the Falcon Heavy. The biggest advantage of newer providers is the cost per
kilogram has been significantly lowered. Thus, to dominate the market share of the industry the
proposed launch vehicle will need to have the capability of mass production or high reusability to
decrease the turnover time of the vehicle. This will automatically lower launch costs and allow it
to be more competitive than other launch vehicles in the market. However, the existing customers
will not just simply shift to the new providers. This is because the customers might have already
placed orders or contracts on the existing providers and the uncertainty of capability from the new
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providers. The customers will need to gain confidence from the new provider. Especially when
delivering humans to the ISS as it requires greater reliability. Since the demise of the Space
Shuttle program there has been intense pressure to fill this operational gap, the proposed launch
vehicle should aim to fill this gap.
Finally, civil government, military or non-profit carried out about 80% of the launches in 2011. If
the launch vehicle can accommodate for this market it will further increase the market share
potential of the proposed vehicle.
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5.2.
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MECHANICAL PAYLOAD
Figure 11 Number of operational Satellites in Orbit (UNION OF CONCERNED SCIENTISTS, 2012)
There are many satellites currently in operation, as can be seen in Figure 11 Number of
operational Satellites in Orbit . It is a highly evolving market and new constellation ideas are
being developed such as Galileo as well as benefits being shown for smaller satellites such as
cube satellites. The launch vehicle should be able to carry a few satellites as it is a great source of
revenue.
The mass of the six person system was estimated between 8-10tonnes (humans and ECLSS)
which would leave a mass of ~20 tonnes for mechanical payload, subsystems/navigations in
addition to the structural mass of stages and fairings.
Payloads identified in the inception report included:




communications (2550kg-6900kg)
Earth observation(1150-6650kg)
Military (2850-18000)
Earth Sciences (1550-4850)
Most of which could potentially be accommodated apart from the types exceeding 5000
kilograms unless less crew and supplies were carried on-board.
5.2.1.
REQUIREMENTS
Satellites require clean room conditions which can be supplied via High-Efficiency
Particulate Air/HEPA filters which are also used on the human module for the internal cabin. In
addition methods are required to limit the vibrations of the rough ride to space and these are
discussed in later in the form of attachment methods.
5.2.2.
MECHANICAL PAYLOAD SELECTION
With a 4 metre diameter these are this section will provide a list of the payloads which may be
accommodated. Table 31 of Chapter12.1.1 a list of possible satellites to launch can be seen. With
a payload fairing of 6m most are accommodated and the majority may be capable of being
launched in multiple batches of for example four medium/small sized satellites evenly distributed.
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EnviSat has the largest height of 10m and a mass of 8.2t may not be accommodated as it would
take the payload mass over the target amount. All other satellites should be accommodated.
5.2.3.
ATTACHMENT METHODS
The mechanical interfaces are well defined such that they direct the pathway for the loads on the
spacecraft structures to so these systems usually employ a central thrust-load-bearing member
such as cone or cylinder which all other structures/stages and payloads are attached to strong
points or via platforms or trusses. (FORTESCUE, Peter and Stark, John, 1995). RUAG Space AB
provides reliable Payload adapter with several decades of experience and have launch 489
separation systems which were 100% successful. (RUAG SPACE AB, 2010)
Figure 12 -RUAG Payload Adapter System (RUAG SPACE AB, 2010)
5.2.4.
MULTIPLE SATELLITE CONSIDERATIONS AND PICO-SATELLITES
In recent years there has been increased interest in producing small satellites and even NASA
conduct the NASA CubeSat Launch initiative (CSLI) which provides the opportunity for small
satellite payloads to fly on planned rockets launches (KEETER, Bill and Lind, Rocky, 2012).
However the industries experience stringent budgets which limit the launch of small satellites as
there would be large costs associated with launching them and not enough return as compared to
larger satellites. With the aim of helping with these constraints many studies have been conducted
to reduce the cost of launching small satellites and reduce the vibration experienced during
launch. One such study developed a method known as Secondary payload adapter ESPA to isolate
the entire spacecraft structure from these launch vibrations currently used on Expendable Launch
vehicles but the benefits applied to this Re-usable system would also have beneficial impacts.
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Figure 13: Schematic of ESPA mounted to EELV upper stage at Standard Interface Plane with one primary
payload and six small satellites; all payloads have whole-spacecraft launch isolation (MALY, Joseph R et al.,
2000)
The ESPA is expected to reduce the launch dynamic environments for all payloads and thereby
reduce the cost of launching small satellites to less than 5% of the cost of a dedicated launch
vehicle (MALY, Joseph R et al., 2000).
Satellites such as CubeSats were designed to have standard units whereby the width and height
remain unchanged and the length may increase in specified units, this allows for a common
deployment method for the Satellites. The system is known as the Poly Pico satellite Orbital
Deployer (p-pod)
Figure 14: Poly Picosatellite Orbital Deployer (P-POD) and cross section (MUNAKATA, Riki et al., 2008)
5.2.5.
POTENTIAL INFRASTRUCTURE
INFLATABLE SPACE HABITAT
Bigelow Aerospace has developed a new as one of the potential avenues for space infrastructure
this type of system could be launched and set up as a possible intermittent step for even longer
missions. The BA330 has its own suite of Life support and is being proposed to carry 6 people
with a volume of 330 cubic metres. It has a diameter of 6.7m and a length of 13.7m with an
approximate mass of 20t (BIGELOW AEROSPACE 2012). From past inflatable structures
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namely Genesis I testing of the construction and the inflatable technology was successfully
launched in 2006. It had contracted diameter of 1.6m compared to the inflated diameter of 2.54m
this gives reason to believe that the BA330 will have an approximate contracted diameter of 4.22
metres assuming the same ratio is followed and thus could be accommodated in the proposed
fairing size.
The technology includes superior radiation technology better than the ISS and substantially
reduces the dangerous impact of secondary radiation. (BIGELOW AEROSPACE 2012) It has
improved debris shielding over traditional aluminium structures (such as the ISS) with its
innovative micrometeorite and Orbital Debris Shield. Kevlar and Vectran-like materials are the
building block behind the technology, which is present in multiple layers in the 15cm walls of the
Genesis I (THOMPSON, Mark, 2012)
Figure 15 Depiction of the BA330 compared to an ISS module (BIGELOW AEROSPACE, LLC, 2012).
5.3.
5.3.1.
HUMAN MODULES
INTRODUCTION TO PAYLOAD STRUCTURAL CONSIDERATIONS
It was determined that the aim of this re-usable spacecraft should be to replace the Space Shuttle
to potentially keep supplying the ISS as this would be a potential investment area, and largely
focussing on the construction of space infrastructure a largely growing sector of space activity.
To ensure that the tourism market was not forgotten in the investigation considerations were made
for 30 people as well as the common astronaut missions of 6 people. In addition to the more space
plane type notion it was felt that the use of a capsule should also be investigated.
In line with this thinking as a starting point the payload bay design would be a modification of
the dimensions of the cargo/payload area previously used in the Space Shuttle. The space shuttle
orbiter dimensions where 18.3m by 4m (diameter) after further investigations the decision was
made to slightly increase this to include larger diameter satellites such as the SpaceBus series with
a diameter of 5.5m? However there is a bit of concern about converting this previously
unpressurised cargo bay into a fully pressurised cabin to accommodate thirty people as this would
increase the mass significantly and the cost to produce could possibly be prohibitive.
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5.3.2.
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PRESSURE VESSEL CONSIDERATION
For a thirty person cylindrical (shuttle-like) spacecraft the wall thickness was estimated using the
typical pressure vessel equations an example calculation may be seen in the 161Here a minimal
wall thickness of 0.67metres was required and the vessel would weigh 8468kg
Manned Spacecraft include truss structures, semi-monocoque concepts in essence thin walled
vessels with supports and stiffeners, (stiffened pressure vessels).
The use of a capsule versus a lifted body (space plane) system was debated the benefits of using
each were discussed among the group and although the space plane system would be able to
provide a more gentle landing and create a great deal of media attention, the monocoque capsule
design would be more technically achievable as such designs tend to have less complexity when
compared to lifting bodies. They are traditionally designed to have Service Loads and Strength
requirements which imply a margin of safety > 1. The mass and volume calculations can be seen
in the Appendix and the capsule shape was estimated as a truncated cone for calculation purposes,
from this using a typical wall thickness of 10cm the mass would be
5.3.3.
SEATING AND LAYOUT
To quantify the size of the module needed for occupying the humans some basic dimensions were
collected and some calculations for people seated were done to determine what size the module
might need to be to fit the humans sufficiently. However as this was a conceptual project this was
not the same as calculating the habitable volume which would involve predicting how much of
the pressurised volume was anthropometrically useable. Once the spacecraft reached the design
phase it would have to be taken into account the optimal, performance or tolerable volumes as
well as number of crewmembers and mission duration in order to select the most optimal
scenario. Habitable volume is important as there is a minimum required amount of free or
accessible volume necessary for the crew to perform tasks without incurring physical,
physiological, or psychological impairment for the duration of the mission (SIMON, M et al.,
2012).
A tourism market study was conducted which inspired the investigation of the possibility of a 30
person craft, the overall feeling within the group was that although it may not be achievable in the
early stages of the mission; there should still be consideration of this as in future years a tourism
market should be incorporated.
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Figure 16 Common Dimensions Used for Seat Sizing and Layout (95th %ile Male Represented) in inches
(0.0254m) (GOHMERT, Dustin M, 2011)
The important Anthropometric data to determine seating specifics were the following: Seated
Height, Mid-shoulder Height, Bi-deltoid Breadth, Buttocks to Popliteal Length, Heel to Popliteal
Length, Hip Breadth (GOHMERT, Dustin M, 2011). These measurements must also take into
account the change in dimension when the Launch/Entry suit is activated as shown in Figure 16
Common Dimensions Used for Seat Sizing and Layout (95th %ile Male Represented) in inches
(0.0254m) , which will provide the interface between the body and the seat. Assuming the seats
are roughly 34 inches (0.864m) in width this was used to make a prediction of the diameter
requirements. Assuming passengers are seated in two layers, three on an upper level and three
below similar to the Space X dragon layout (REGAN, Rebecca for (NASA's John F. Kennedy
Space Center), 2012). A diameter of 4metres would be feasible for the capsule as three seats side
by side would amount to 2.58metres and provided a minimum amount of space is allowed
between seats this seems achievable.
For the height allocation, assuming each seat is placed in a recumbent orientation (12° to floor)
the height with the addition of a seat liners etc. can be estimated at approximately 0.71metres.
Therefore leaving adequate space for ingress/egress for instance 0.8metres between upper and
lower seats, it can then be inferred that a required space of roughly 2.31 metres just for the seat
logistics is needed. Thus validating a capsule sizing of roughly 3metres at highest point (due to
conical exterior shape)
5.3.4.
G FORCES
The difficulty with G forces is that the resultant between the applied and gravitational
acceleration has the effect of increasing the pressure of blood circulating from heart to brain. This
is effectively because the g component increases the equation P = ρgh. Author: Emily Ann Carter
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Where ρ = blood density, g = the applied gravitational acceleration, h = the height of the brain to heart column of blood
This then makes it difficult for the cardiovascular system to supply blood to the vital organs and
this is the reason why there is a limit of human tolerance. Blood flow to brain is diminished force
is directed in the +Gz (head to foot). On the other hand if directed in the –Gz direction (foot to
head) blood poles and astronauts may experience “red-outs” where vision becomes tinted red. Scientist have discovered that directing the acceleration in the chest through back direction
provides the best tolerance form humans providing the most normal distribution of blood flow.
The best way to orient the G forces is in the recumbent position.
The effective + Gz Tolerance therefore would be increased if the h component in the P = ρgh
equation is reduced i.e. reducing the aortic valve/eye column which is achieved by reclining seats.
The G tolerance improvements are fairly linear with reduction in effective column height for
o
example at 75 seat back angle, column height is reduced to one half and +Gz tolerance is almost
doubled. (MCDONALD, P. Vernon et al., 2007)
For a space plane consideration such as the Space Shuttle an angle of 6 degrees would be placed
between the seat and floor; NASA suggested it should be no greater than that (NASA, 2011).
For the capsule however astronauts showed a preference for an angle of 12º for the torso
inclination resulting from research carried out in early phase of capsule design research.
(MCDONALD, P. Vernon et al., 2007) So from this an angle of 12 degrees is suggested for
this system.
Figure 17: Direction of g Forces Experienced During Landing in a Soyuz Capsule (Left) and the Space Shuttle
(Right). (CLÉMENT, Gilles, 2011)
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The G forces that would be experienced by this conceptual design were calculated and were
shown to be within 4 G’s in G load simulation diagram Figure 104
This would be a fairly tolerable level and is comparable to currently existing launch systems. The
seats should also allow for a five/ six point attachment seat belt to help stabilise the body from the
high G loads and prevent excess movements or straining of neck areas especially.
5.3.5.
VIBRATION AND DAMPING
Figure 18 - Showing Wire rope insulator produced by Enidine (WR16 Series) (VON BENGTSON, Kristian,
2012)
Research has identified the use of a very simple system using some heavy duty wire (as shown in
Figure 18) which would be capable of supplying enough damping for the seat. Wire Rope
Isolators are widely used in industry and are favourable over systems such as hydraulic damping
cylinders due to their lower cost as well as high damping capacity.
These simple wound wire is capable of max compression of 500kg which is could be placed at
each corner of the seat to withstand the high compressive loads even in the event of Launch Abort
systems at 10G’s 5.3.6.
ENVIRONMENTAL CONTROL AND LIFE SUPPORT SYSTEMS
The protection of humans in space is a difficult and expensive issue. The requirements for human
survival were outlined in the Inception report (MDDP-REUSABLE SPACE SHUTTLE
GROUP1, 2012). The various methods for supplying these requirements will be discussed here.
Two different scenarios were proposed one a short thirty person tourist trip and then a six
astronaut mission for a longer period. It can be seen that for a thirty person spacecraft the
regenerative systems would not be sufficient to supply their needs as they have are designed to
supply only 3 to 6 people (JAMES, John T and Macatangay, Ariel, 2009).
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There is also not a need for them as it was only initially proposed that the tourist flights would be
somewhat of a novelty occasion and not a long deep space endeavour. Passengers would
experience microgravity and stellar views of earth and space, with flights lasting no more than
two days and would not leave the craft once in space due to the lack of feasible accommodations
for such a large number of people. After retrieving the results from the market research, although
there was a desire to go to space the price would be a large deterrent for most people so the
considerations progressing are largely for a six person craft.
On researching the methods for supplying the cabin with sufficient consumables it was found that
there are three proposed methods for achieving this. These would be to use in situ resources
brought on board, Physiochemical Life Support Systems or Bio-regenerative Life Support
Systems. However most of the passed programs such as the NASA shuttle or the Russian Soyuz
use stored consumables or are highly reliant on physiochemical processes such as on the ISS.
Physiochemical life support systems are what programs like the ISS currently use and use a range
of chemical reactions such as electrolysis for oxygen production in addition to mechanical
processes.
A biological life support system (BLSS) is usually considered as complementary to a
physiochemical life support system, able to support certain functions of a life support system, in
order to maintain or improve the habitability for the crew. (HORNECK, G et al., 2003) It
becomes more and more apparent that Bio-regenerative methods may become the future of life
support especially with the research done by the Ames Research Centre.
5.3.7.
AIR REVITALISATION
To extend the possible mission length regenerative systems would provide a way of supplying the
astronauts without the need to carry large stores of consumables but perhaps just small amounts
for contingency purposes and re-entry when a capsular system is used. At present there are
technologies used on the ISS which provide regenerative supplies and it would be possible to
modify these technologies and install them into the crew vehicle. However there is a particular
level of risk associated with these systems due to their complexity and in the future one should
not rely wholly on these physiochemical systems due to their large duty cycles, repair needs and
reliability. At present, however these systems are not used for missions less than three weeks and
will become very crucial if missions are to eventually lead to the far reaches of space like mars or
moon bases.
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Figure 19 Crew needs and by products (PERRY, Jay L and LeVan, M. Douglas, 2002).
The provision requirements for a six person craft on a typical eight day mission are as follows;
each person requires 0.84 kg of oxygen a day therefore for an eight day mission with six
astronauts this suggests a need for 40.32kg of oxygen for the duration. This mass of oxygen has a
volume of 30.5 m3 (see appendix for calculation) however as a liquid it would take up much less
volume requiring only 35.32 litres of liquid oxygen according to conversion tables (AIR
PRODUCTS AND CHEMICALS, INC, 2012), which is equivalent to a volume of 0.04 m3. Thus
explaining why most spacecraft use liquid oxygen storage tanks as they take up much less of the
craft’s already limited volume. To reduce the mass budget for these tanks, NASA developed
composite overwrapped pressure vessels which would store the liquid oxygen and nitrogen.
Oxygen tanks –Typical oxygen tanks used use an Overwrapped Pressure Vessels Program over
the use of all-metal designs, these are capable of storing high pressure gases between 20.7MPa
and up to 33.6MPa for gases such as oxygen, nitrogen, and helium. The agency used six 66 cm
such vessels in the Environmental Control and Life Support for nitrogen (FORTH, Scott et al.)
And thus a similar strategy will be used in this project as well as 6 for oxygen as oxygen has a
critical pressure of 5MPa and would need to be stored above this pressure.
Contingency oxygen supplies use multiple oxygen tanks so if one fails another one can be utilised
as well as it provides a more balanced storage capability so they may be distributed evenly around
the base of the capsule/trunk.
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Figure 20 - Typical Stress Pattern. (KEDDY, Christopher P)
Spherical pressure vessels offer better stress distribution patterns over some of the alternative
configurations and for liquid oxygen which requires due care when storing these seem to be the
better choice. A typical Stress pattern can be seen above in Figure 20.
Some of the common materials used as overwrap fibres are shown in Figure 21
Figure 21 - Comparison of Composite Overwrapped Pressure Vessel Fibre Properties
Extensive tests are carried out on the materials used especially Kevlar due to its durability. To
build a spherical pressure vessel, two titanium hemispheres had to be welded together to form the
liner which would also add a level of risk to the vessel and increase the potential for failures.
Failures in pressure vessels may lead to the loss of the spacecraft or potential fatalities. NASA
employed many fracture control test programs to ensure its safety and later lowered safety factors
as the materials properties became more understood.
5.3.8.
REGENERATIVE METHODS
The Oxygen generation Assembly (OGA) used on the ISS has the capacity to supply six
astronauts and has a production rate of 2.3- 9kg of oxygen per day during continuous operation
(NASA, 2008) and also has the option to operate on a cyclic basis producing less oxygen perhaps
as a form of redundancy. The system has a mass of 113 kg and is the size of a common
refrigerator now for a launch vehicle the volume and mass constraints would prohibit the use of a
system such as this one for missions less than three weeks.
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5.3.9.
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CARBON DIOXIDE REMOVAL
Carbon dioxide removal may be achieved by using Lithium hydroxide canisters as the Shuttle did
however this is an expendable process and there have been new developments in technology
which use regenerative methods. A major disadvantage to using Lithium hydroxide systems is
that to eliminate the Carbon dioxide that one human produces in a day there would be a need for
1.5kg of lithium hydroxide. Canisters used on the shuttle had a capacity of 3kg (volume:
0.006m3) (JAMES, John T and Macatangay, Ariel, 2009) which means that several canisters
would have to be taken on-board to satisfy mission requirements, suggesting the need for
improvements to be made.
One such technology is the use of solid amine solvents which is perhaps and improvement upon
some of the other proposed methods such as molecular sieves to provide better Co2 capacity and
less capacity for O2 and N2 in essence drier beds. The amine swing-bed system will use an
adsorbent material comprised of interleaved layers of beads coated with a proprietary amine
compound noted for its affinity for CO2 and water. (JAMES, John T and Macatangay, Ariel,
2009) It is also advantageous because it may also be used to regulate humidity levels in the cabin
and thus the use of heat exchangers is not required.
5.3.10.
WATER AND WASTE
Water although it may be produced as a by-product of many of the other systems such as the
amine swing bed it is traditionally produced via the use of fuel cells which are used for power
supplies for many of the crafts electrical systems. This system will still be employed as it is a
proven technology. The Oxygen Generation System if eventually used can include a Sabatier
process making the ECLSS a closed loop whereby the wasted CO2 and Hydrogen are used to
process water rather than venting overboard. Future applications will see the use of technologies
to recycle wasted water and purify it for re-use such as the technologies used on the ISS and new
technologies such as the AMES group water walls program. As of now wasted water will be
stored for miscellaneous uses if any (for example heat exchanger cooling) with the excess vented
overboard as was the case with the shuttle (NASA, 2002)
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5.3.11.
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TRACE CONTAMINANT CONTROL/FILTERING
Table 5 - Spacecraft Cabin air quality Parameters (PERRY, Jay L. and LeVan, Douglas, 2002)
The craft should be able to maintain these standards for various particles and gasses as discussed
in a NASA report. With these needs there would be a marginally high risk factor as if any of the
systems deemed to carry out the task fails the cabin could become a hazardous environment for
the occupants.
Table 6 - CO2 measuring capabilities on the ISS (JAMES, John T and Macatangay, Ariel, 2009)
All are used in the ISS for the space craft using two the MCA and the CDM should suffice and
provide enough monitoring of the waste gas levels etc.
5.3.12.
PRESSURE MANAGEMENT
Pressure suits: Contingency Hypobaric Astronaut Protective Suits was specifically developed for
the use by commercial space passengers
NASA uses the Advanced Crew Escape Space Suit System (ACES) which is worn during Launch
and Re-entry and provides a means of survival in the event of cabin depressurisation and
emergency egress. There are a number of benefits to using this system over newer alternatives; it
is a flight proven technology whereas other systems such as CHAPS may have only been tested in
a centrifuge or spaceflight but few by comparison to ACES. It has a range of capability including
inbuilt parachute systems and early full operational capability and flight readiness (NASA, 2012)
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Full pressure suit (launch & re-entry), gloves, boots, helmet
Provides 3.46 psi nominal operating pressure
Protection in low altitude bailout and ground egress scenarios
Ability to operate as open or closed-loop demand breathing system
Emergency breathing system
Liquid cooling system
Headset communication system
Search and Rescue identification and emergency communication hardware
High-altitude, automatic-inflation parachutes
Automatic-inflation life preserver
Survival drinking water packets
Figure 22 - NASA ACES -CAPABILITIES (NASA, 2012)
5.3.13.
NUTRITION
The methods used to prepare food on the shuttle used a forced-convention oven or additionally a
hot water supply to reheat foods for serving (PANDIT, Ram Bhuwan et al., 2007). However these
systems were relatively inefficient due to the indirect heat transfer which led to the development
of further technologies. Using the new technology for heating food on the spacecraft there can be
a reduction in mass, traditional systems weighing convection heating 3475kg. Originally the
‘Ohmic heater’ was developed by EA Technology in the UK and was subsequently further
explained by Sastry in 1994. (SUN, Da-Wen, 2005) The idea developed by a group at Ohio State
university for using Ohmic heating for space application aimed to produce a very light weight
system which could potentially be used on long missions to Mars where life support systems must
operate under even more restricted volume and mass requirements. The Ohmic system equivalent
mass was estimated to be 839 kg which is significantly less than the traditional system, there is a
level of risk in this evaluation since the authors did analysis on a mars mission which means that
for Lower orbits the equivalent mass may change however they predict mass savings over the
traditional systems even for these missions.
Figure 23 - (a) Pouch interior showing electrode; (b) exterior of pouch, showing one of the electrode tabs used to
contact the power source (PANDIT, Ram Bhuwan et al., 2007).
The proposed layout of the system included a pouch with electrodes inside and assumed a more
rectangular shape to make improvements on existing pouches in terms of more uniformed heating
as well as ability to stack easily (PANDIT, Ram Bhuwan et al., 2007).
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Figure 24 - Schematic of the food heater enclosure.
Typical units used in ground industry for other uses are in the price range of 35 thousand pounds
(C-TECH INNOVATION, 2012) however specific units could be assembled, which may cost
significantly less.
5.3.14.
FIRE PREVENTION
Fire extinguishers should be present in convenient locations around the craft and where possible
the use of non-flammable materials used. Also the atmospheric composition would have to be
sufficiently balanced to avoid having too much oxygen making fires more likely. For this reason
an atmosphere mixture with pressures equivalent to sea level should be used and Nitrogen should
be used for dilution. Oxygen levels should be kept greater than 20kPa to avoid hypoxia and less
than 50kPa to prevent toxicity and limit combustion. A cabin pressure of 101kPa is targeted with
21% O2
5.3.15.
RENDEZ VOUS/DOCKING
This is the mechanism by which docking to space habitat / station is achieved and typically a
common berthing mechanism is needed if docking to the ISS. There have been international
agreements which suggest the use of a design developed by ESA as shown in Figure 25 and the
use of similar but not identical systems may be used with this system.
Figure 25 -ESA-developed berthing and docking mechanism (INTENATIONAL DOCKING STANDARD, 2011)
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5.3.16.
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RADIATION
DOSAGE
Due to the fact that this mission is aiming at Low Earth Orbit (LEO) the need for substantial
radiation shielding was debated. For the basis of down-selection some of the plausible
technologies were discussed. The aim of a radiation shielding system would be to keep the risks
as low as reasonably achievable (ALARA principle) taking into account the potentially large mass
penalty this could have on the overall launch load. To more accurately assign a radiation shielding
technology the most likely mission scenario needed to be chosen and the level of shielding needed
was selected based on the severity of the risks. To help ascertain the severity of risks the radiation
doses of some common human activities were found for comparison. The effective dosage for
radiation associated with these common human tasks is shown in the table below.
Table 7: Showing the Radiation doses of some human activities and spaceflight
Procedure
Effective
Comparison to
Additional risk of developing
Dose
Background radiation
cancer (1in 5 risk in general)
10
3 years
Low
7
2 years
Low
12
4 years
Low
0.03
3 days
Negligible
1.5 years
Low
/mSv
Computed Tomography
(CT)- Colonography
Computed Tomography
(CT)-Chest
Coronary Computed
Tomography Angiography
(CTA)
Commercial Airplane Dosage
Space flight ( LEO )
(8-day mission orbiting the
Earth at 460 km)
5.59
From this table it can be seen that the highest sauce of radiation comes from a chest CT scan
(7mSV) which when compared to the risks of travelling in LEO in 8 days which received a
dosage of 5.59mSv. The shuttle would have provided some radiation shielding from the walls of
the craft. Typical LEO missions would provide between 10 and 30 cm of shielding and no more
due to high mass penalty.
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Table 8 – Equivalent limits recommended as career doses (keeping excess lifetime risk lower than 3%) (REITZ,
G.)
Table 8 shows the dose limits for humans of different ages based on a 10 year career, assuming
the dose is being kept below a 3% increased risk as suggested by the National Council on
Radiation Protection and Measurements (NCRP). Missions within LEO are somewhat protected
by the Earth’s magnetic fields and thus radiation exposer can be limited by setting up career limits
whereby a humans number of days in space are limited to some defined value which ensures that
there risks of developing cancer are below 3%. Most launch systems abide by this method
because the implementing of current shielding technologies requires too much mass to extend the
shielding any further.
RADIATION SHIELDING METHODS
The current technologies used to protect against SPE are sufficient however it is clear that major
optimisations are needed to extend these technologies for longer missions as well as to adequately
account for GCR as well. This is the challenge.
Currently to withstand a nominal atmospheric pressure the walls may only be a few mm thick
leading to very limited protection (not taking into account the added thickness of the structural
supports and heat shielding). With walls of thickness of just a few millimetres (0.67mm) the
radiation shielding provided is 0.01g/cm2 which according to Figure 26 the received dose would
be just above 1mSv/day. Typically however there are wall thicknesses of 10 cm including some
structural support in the form of trusses etc. A wall of 10cm of aluminium would thus provide
shielding of depth 27g/cm2 and exposure is then reduced to about 0.5mSv/day
Some of the common materials used for radiation shielding are: Aluminium, Water, Cryogenic
Liquids, and Polyethylene. The protection against SPE can be sufficiently achieved by using these
common technologies and with a minimum amount of shielding of 10g/cm2 less than 1 mSv per
day
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Figure 26 - Showing the comparison dose equivalent calculations against shielding thickness for several
materials from trapped protons at an orbit of 51.6°x390km (CUCINOTTA, F.A et al., 1999)
From Figure 8 it can be seen that liquid hydrogen remains the better option for shielding and
research is still being conducted to develop materials which are highly impregnated with
hydrogen. Active Shielding technologies such as magnetic shielding or plasma shielding were
also researched however these devices require significant power sources or are exceptionally
massive to achieve significant GCR risk reduction. (Durante and Cucinotta, 2011) Future
missions into deep space will increase cosmonaut exposure when compared to short duration LEO
missions and different strategies or multiple layered systems may need to be installed if these trips
become likely.
To conclude Aluminium shielding will still remain the most feasible option as the technology to
produce some of the better ones may not be quite ready or require new materials to be developed.
An aluminium shield of this nature 10cm thick as said before would give the occupants an
estimated dosage of 0.5mSv/day and would sufficiently low enough to be sufficiently handled so
as to avoid increased cancer risks.
5.3.17.
FUTURE PROOFING/INNOVATION
Water walls which serve as not only radiation shielding but as life support as they contain algae
among other substances to eliminate CO2 and contaminants as well as provide O2 for
atmospheric revitalisation current studies look promising. Some of the underlying technologies
used in this system are already in application in other areas for example the FO membrane is
currently being used in the military for purification of urine to drinkable water as can be seen in
Figure 27 - X Drink water purification bag which uses Ames developed FO membrane.
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.
Figure 27 - X Drink water purification bag which uses Ames developed FO membrane (COHEN, Marc M et al.,
2012)
The potential architecture that the group is proposing can be seen in figure 10 and the aim is to
produce this system so that it can be wrapped around in a cylindrical fashion around the craft
however this technology is still in its research and developmental phases but would be beneficial
in future applications.
Figure 28 - Proposed Water Wall Architecture (COHEN, Marc M et al., 2012)
5.3.18.
CAPSULE REUSABILITY
It is estimated that the capsule may be re-used approximately 5-10 times depending on the
conditions of launch and re-entry. This largely depends on the effectiveness of the heat shield and
the re-entry trajectory employed. Also of large importance would be the parachute systems both a
breaking parachute and a main one to provide a soft landing a large risk factor is necessary here
since a large number of backup chutes may not be feasible. As outlined elsewhere (Cross
reference Charles Report Section) the use of the Space X Launch Abort System which acts not
only as launch abort in the event of mission failure but can provide extra assistance to landing so
by firing a few of the Super Draco engines at a particular altitude to slow approach before
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deploying drogue parachutes. In this way if speeds don’t decline as planned more engines may be fired and perhaps a main parachute and soft landing may be achievable on land. However this
may be
As for sea landing the capsule would have to be thoroughly coated in anti-corrosion coatings to
reduce maintenance and minimise damage. Once such coating became popular in NAS the , IC
531 Zinc Silicate a non-toxic water-based material, which dries within 30 minutes to a ceramiclike, hard, durable finish (NASA, 1995).
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5.4.
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FUEL OPTIONS
Propellant systems that have been successfully developed and deployed include liquid, solid and
hybrid propellants. In addition, inert fuels have been used with power provided by
electromagnetic radiation or by thermal conduction, however analysis performed suggests that
neither or these are viable choices for a launch vehicle.
5.4.1.
LIQUID CHEMICAL PROPELLANTS
Specific impulse is the main consideration in determining a suitable propellant or combination of
propellants. The specific impulses given in the following section are averages of the vacuum and
sea level specific impulse, multiplied by 0.9 to account for deviation from ideal values, due to
incomplete combustion, reaction kinetics, viscous effects and divergence (Huzel and Huang
1992). Fuels are considered first, with oxidisers separately. Monopropellants are not considered as
they do not offer high enough specific impulses for a launch vehicle.
The highest specific impulse is obtained from Liquid Hydrogen fuel with metallic additives. Since
these additives can be toxic or expensive (examples include lithium and beryllium), pure
hydrogen is the more viable choice. This offers a specific impulse of 380s (Huzel and Huang
1992). The drawbacks of hydrogen are that it must be stored at a deeply cryogenic temperature
(44°K), it has a very low density (67.8 kg/m ), leading to a high volume and therefore a larger
and more expensive vehicle. Hydrogen also causes embrittlement of metals, by dissolving and
then re-combining within the material. This causes porosity and initiates cracks. All of these
issues can and have been tackled in the past, however more benign fuels could be an attractive
alternative to ease the design process for reusability.
RP-1 is a highly refined gasoline derivative. This fuel has good material compatibility and can be
stored at room temperature and pressure. The fuel is flammable, leading to handling safety issues,
but these are overcome routinely in everyday life. The specific impulse of RP-1 is 296s (Huzel
and Huang 1992). RP-1 has reasonably high density (approximately 800 kg/m ); however its low
specific impulse will result in increased mass and size. The drawback of RP-1, as a gasoline
derivative, is the ever increasing cost and scarcity of crude oil and the impact on the green-house
effect of carbon dioxide exhaust.
A possible intermediate between RP-1 and Hydrogen could be liquid natural gas, or LNG. This
has a mid-range density (450 kg/m (Kumar 2011)) and a higher impulse than RP-1 (305s (Huzel
and Huang 1992)).This fuel may be more sustainable, as natural gas can be produced biologically.
LNG must be stored cryogenically, but not as deeply as hydrogen. A typically storage temperature
would be similar to that of liquid oxygen (~90°K), introducing the potential for tank design
commonality and thereby reducing overall costs.
Author: James Roper
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All of the above have been considered using Liquid Oxygen as the oxidiser. This is because
oxygen is widely available and relatively benign. Many other oxidisers, such as Fluorine,
Chlorine/Fluorine/Oxygen compounds and strong acids (red fuming nitric acid) were considered,
however they are too toxic or dangerous to employ for this project. Many are also hypergolic,
meaning that a spill would ignite not only the fuel but possibly any other material that came in to
contact with the oxidiser.
There are also numerous other possibilities for fuels, but again, many of these are too toxic or
dangerous to contemplate seriously. A toxic fuel or oxidiser would drive higher costs due to the
increased risks and the specialised equipment and safety systems required to handle them. A fuller
list of fuels and accompanying analysis is given in Appendix 12.3, including inert fuels.
5.4.2.
SOLID AND HYBRID PROPELLANTS
Solid propellants are used for the solid rocket boosters of the Space Shuttle. The most energetic
combinations can offer specific impulses in the region of 270s, which is poor compared to a
typical liquid fuel (a 10% reduction). They are also used as long term storable propellants for
missiles, including the LGM-30G Minuteman III. A risk is identified with such propellants, as
both the oxidiser and fuel are pre-mixed, making accidental ignition a higher possibility.
Once ignited, the burn of a solid propellant cannot be controlled or shut down. Such propellants
are also dangerous to manufacture, often including unstable explosive substances such as nitroglycerine (Sutton and Biblarz 2001). This entails that they will be expensive, requiring highly
specialised facilities for manufacture. In addition, the exhaust products are often toxic and can
contain corrosive substances such as hydrochloric acid (Sutton and Biblarz 2001). The Titan IV
failures in 1993 and 2009 (USAF 2009) and the Delta II in 1997 (Evans 2000) were all caused by
failures of solid rocket boosters and resulted in clouds of toxic, corrosive debris.
Solid rockets have been shown to be re-usable (at a cost) in the Space Shuttle. The spent casings
are cleaned, pressure tested and crack detected prior to re-loading. Mechanical fasteners are used
to facilitate maintenance (NASA).
Hybrid propellants typically include a solid fuel and a liquid oxidiser. Small scale hybrid engines
are used on the Virgin Galactic SpaceShipOne and SpaceShipTwo (MALIK, Tariq, 2004). Unlike
pure solid rockets, these engines are controllable and can be shut-down on command (by shutting
off the propellant feed).
The solid fuel is often a hydrocarbon, resulting in benign combustion products including carbon
dioxide and water. In addition, loading and handling of the solid fuel is very much safer than for
liquid fuels. This makes hybrid propellants an attractive option.
Author: James Roper
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A problem faced in larger hybrid motors is combustion instability and pressure variations, as
intimate contact of fuel and oxidiser is difficult to achieve (Sutton and Biblarz 2001). As a result,
this technology is not mature enough yet to be deployed on large rockets, but focussed
development could overcome this in a few years.
The specific impulse of a hybrid system is comparable to that of a liquid hydrocarbon system, as
they are chemically similar. Since they use solid hydrocarbons, there is reason to believe that
waste plastic could be re-cycled in to rocket fuel, perhaps lowering the fuel cost and combating
the environmental damage caused by landfill of plastic waste.
Hybrid and Solid fuels have a significant drawback in terms of the dry mass of the motor. This is
typically 10-30% of the ‘wet’ (or fuel-and-oxidiser loaded) mass (Sutton and Biblarz 2001). This
is because the motor casing must be capable of withstanding the high temperatures and pressures
of combustion. As a result, they give much poorer delta-V than comparable liquid-fuelled
engines.
5.5.
5.5.1.
LAUNCH PHILOSOPHY
SPACE-PLANE CONFIGURATIONS
The Skylon design attracts the most credibility as a
potential single-stage to orbit space-plane. This has
been verified through independent review by
engineers at the European Space Agency.
Space Planes typically result in high dry mass
fractions, which are a driver for higher costs. This
is due to the need to heat-shield the entire
Figure 29 CGI of Skylon Space Plane
vehicle and also the addition of aerodynamic surfaces and structures.
However, the fully heat-shielded vehicle can be used to ferry payloads both to and from orbit. The
space-plane configuration provides a controlled soft landing capability. This would be an
advantageous attribute if there were a need to always return the payload (i.e. manned vehicles) or
if there were high customer demand to return other objects from orbit, such as defunct satellites,
debris, etc.
There is also an increase in safety in the event of an engine failure, as the vehicle should be able
to make a controlled horizontal landing. This should however be given a more realistic overview,
in that failures may not be confined to the engine. In this case a more conventional launch-abort
system would still be required (but may be difficult to integrate).
Author: James Roper
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The above performance advantages are realised at the expense of both high development costs
and high vehicle costs. The dry mass fraction for the proposed Skylon space plane and a
competing reusable rocket are considered, as these will be a basic indicator of their costs.
For the Skylon, a dry unladen mass of 53T and a propellant mass of 277T are quoted (Bond
2010), giving a dry mass fraction of 19.13%. By comparison, the Falcon 9 rocket will deliver a
similar payload mass. This rocket has a dry mass of 14.85T and a propellant mass of 411T
(SpaceX 2012), giving it a fraction of only 3.613%. This is approximately 5.3 times less than the
Skylon, suggesting that Falcon 9 would have significantly lower costs for both development and
manufacture.
In actual fact, the development cost of Skylon is projected to be $12b (OLSON, Parmy, 2012).
Comparatively, the cost of Space X Falcon development is estimated (private companies need not
disclose this data) as between $1.66b and $4b (NASA 2011). That is, the space plane costs are
between 3 and 7 times greater, which would be reflected in revenue required (from launch fees) to
break even. In the long run, the operational costs of recovering spent stages are likely to be less
costly than Skylon’s development.
There are technical issues associated with the design of vehicle for horizontal flight, which will
lead to increased developmental costs. The vehicle would have to be optimised for both lowspeed flight (take-off and landing) and near-hypersonic Mach 5 flight. In addition, the vehicle will
be significantly lighter on re-entry than on take-off, adding further complication. This entails that
the vehicle cannot be optimised as a typical aircraft would be, i.e. for cruise, thus optimisation
would be a difficult task requiring a numerical model (ESA 2011).
In terms of reusability, the space-plane has obvious operational benefits. The entire vehicle is
returned under controlled flight to a single position. Assuming that all of the components can be
immediately reused, the vehicle need only be refuelled before it can fly again. In comparison with
a staged rocket, which must have all its components recovered and re-assembled, there are
obvious savings in operational cost. In reality, however, the stresses placed on the vehicle during
launch and re-entry will cause damage to the structure and the engines. Access to these areas is
restricted by the aeroshell, making maintenance more difficult. In general, the vehicle has very
low modularity, which will increase its overhaul costs and reduces the potential for design
growth.
5.5.2.
ENABLING TECHNOLOGY: AIR-BREATHING ENGINES
The development of a single stage to orbit space plane is dependent on the development of airbreathing engines, as only they offer high enough specific impulses. This is achieved by obtaining
oxidiser from the surrounding air for as long as possible during the flight.
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The maximum dry mass fraction that can be achieved for non-air breathing engines is
approximately 8% (giving a payload fraction of 1%, for a LH2/Lox SSTO rocket), which is less
than half that proposed for Skylon (19.13%). This makes air-breathing engines essential to space
plane development. The increased development and research required for such engines will put
them at a disadvantage in terms of development cost versus an improved rocket engine, but this
cost-sacrifice is made in order to obtain the operational benefits of an SSTO space plane.
The REL SABRE engine is one possible candidate for an air-breathing propulsion system. This
engine has a maximum thrust to weight ratio of 14, significantly higher than other air-breathing
engines. This is a combined cycle engine, incorporating elements of a turbojet, rocket and ramjet
with a pre-cooler system (liquid air collection engine, or LACE). The major element of this
technology, the pre-cooler, has been successfully ground tested in the past year (Reaction Engines
2012). It is therefore reasonable to assume that the technology could be developed to maturity in a
ten year period.
Competing Scramjet-propulsion would have an excessively high mass, as the T/W is significantly
lower (LACE T/W is five times greater on average) (Varvill and Bond 2003). Scram vehicle
proposals claim to offset this penalty by using the engines as the vehicle structure. However, this
does not eliminate the need to carry large fuel tanks. Thus, this argument seems flawed, giving
Scramjet engines little credulity as a competitor for a launch system engine.
The reliability of LACE is questionable. If used on a horizontally launched vehicle, like an
aircraft, the risks that an aircraft’s propulsion system is subjected to must be accounted for. The
introduction of ice, rain, birds or general debris in to the fragile cooler matrix and the compressor
would be extremely detrimental. Although the engine may be able to return to non-air-breathing
operation after such an incident, it would result in a mission abort. By comparison, an engine that
is non-air-breathing from the beginning is unaffected by these issues. The rocket-type vehicle
spends less time exposed to meteorological risks, as it quickly escapes the atmosphere.
5.5.3.
VERTICAL LAUNCH ASSESSMENT FOR LACE
Another possible use for the air-breathing engine technology employed in the Skylon design is as
a stage for a vertical-launch vehicle. An analysis has been performed, based upon the thrust to
weight ratio of the Sklyon Sabre engines (Varvill and Bond 2003). In reality, new engines would
be developed that were optimised to the vertical launch requirement.
The analysis indicates that nine Skylon-type SABRE engines would be required to launch a 30T
payload vertically, giving a payload fraction of 8.4%. The air-breathing stage would be designed
to provide 1650m/s delta-V, plus the delta-V due to drag (200m/s) and gravity (650m/s). Two
additional rocket stages, with their own engines, are included to perform the non-air-breathing
flight phases.
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The purpose in this was to obtain the benefits from the engine technology, without the need to
develop a space plane. A very high payload fraction is achieved, but the large number of engines
required will drive up costs, nullifying any reductions from not developing a horizontal launcher.
The air breathing engines at present are not suited to vertical launch vehicles, as their high unit
price restricts their application.
In addition, the engines have a combined weight of 92T. This would make their recovery as part
of a jettisoned stage prohibitively difficult.
5.5.4.
PROPOSAL FOR SPACE PLANE CONFIGURATION
The Skylon proposal is adapted and modified to develop a proposal for a space plane. Some
potential faults are identified and reconfigured in order to improve the potential performance,
whilst employing the general concepts of air breathing vehicles.
In the currently proposed Skylon configuration, the mounting of the engines on the wing tips is a
possible failure point. High stress will be placed on the wings by doing so, increasing their
required mass. This will also result in higher drag than could be achieved if the engines were
more integrated with the fuselage. Each engine has its own pre-cooler and compressor, and feeds
directly in to the engine. The engines employ expansion deflection nozzles for altitude
compensation.
A suggested re-design would separate the cooler from the engine. To improve the mass-efficiency
of the pre-cooler, air would be taken in through a duct in the cryogenic hydrogen tank. This would
then feed the oxidiser tank, to fill it up during the flight. This simplifies the design as the engine
no longer needs to switch between oxidiser feeds, and the pressure input from the compressor is
decreased. The air intake would be at or near the nose of the vehicle.
Expansion-deflection nozzles are replaced by a single aerospike nozzle, which is much easier to
design. Ideally, this would be annular and would be incorporated in the tail of the vehicle,
dramatically reducing the frontal aspect of the engine.
A final point would be to improve the structural efficiency by creating a lifting body that could
also function as a wave-rider when approaching hypersonic speeds. This integrated lifting body
would have less sudden geometry changes and thus would reduce high localised heat flux during
hypersonic flight (both on re-entry and ascent).
5.5.5.
SPACE PLANE CONCLUSIONS
Space planes are highly attractive in terms of their performance, which allows objects to be
ferried to and from orbit, with a controlled soft landing, without specialist design of the payload
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itself. As a result, the payload itself is highly modular. They also have dramatically reduced turnaround requirements, increasing their operational efficiency.
This is achieved at the expense of very high developmental and maintenance costs as both the
aeroshell and the engines are new and complex technologies. High speed flight and multiple point
optimisation would be a sizeable technical challenge. In addition, if there were not a need to
return objects from orbit, expenses made in order to re-enter the empty space-plane would have
been excessive and could be out competed by more conventional stage-recovery.
From the point of view of future proofing, it should be noted that the space-plane offers little
flexibility. A modular rocket-type design can be quickly extended to accommodate a larger
payload. For the space plane, each item on the vehicle would have to be scaled.
In closing, the space plane offers great advantages, but requires high developmental cost and
associated risk due to technological immaturity. A more conventional, vertical launcher can be
quickly and easily adapted to reusability, making its development much cheaper. The downfall of
staged vertical launchers, namely operations regarding their recovery, will be the focus of
investigation of this concept – if properly achieved; this will out compete the space plane concept.
5.5.6.
ROCKET CONFIGURATIONS
Historically, only rockets have been able to deliver large payloads to space. Knowledge
accumulated during over sixty years of spaceflight has established an excellent experience-base,
making development a much simpler task. This, in turn, will reduce the developmental costs.
In addition, rockets have no need to generate lift; they can assume the lowest-drag configurations.
These shapes are much easier to construct than lifting bodies and generate less structural stresses,
thereby increasing the vehicle’s possible life and simplifying development.
All orbit-reaching rockets have been of a staged configuration. This is necessary in order to
deliver high payload fraction, thus reducing the vehicle cost. Staging also results in easier to
optimise design, as the envelope of flight of each stage is smaller and thus the requirements are
more uniform.
For a fully re-usable vehicle, the jettison of stages is a disadvantage, as they have to be recovered,
incurring operational expenditures on every flight. In addition, they must be re-assembled in order
to launch again. This will lead to a poor vehicle turn-around time unless measures are taken to
integrate recovery with vehicle design or launch philosophy.
However, it should be noted that return from sub-orbital flight (i.e. the point the stage is
jettisoned) is easier than return from orbit, which would require full heat shielding and complex
design. Alternatively, if the ejected stage is simply an empty tank, it may be more cost effective to
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not recover these parts at all. In this case, focus would be given to recovering only those items
that make up large fractions of the launch cost.
From the perspective of development finance, rockets are a more economical option than
competitors employing more complex launch vehicles or philosophies. The Space X Falcon
vehicle development has a previously-stated upper-estimate of $4b (NASA 2011).
Other vehicle development costs are more broken down. The re-development of the US Space
Shuttle main External Tank (ET) for super-lightweight design was delivered at a cost of only
$43m (LENGYEL, David, 2009). Assuming that common tank design was used, this would
include all stages of a given vehicle. Engine development, including re-usability, is under
investigation jointly by ESA and RKK Energia. This is predicted to total approximately $530m
over seven years (Parmalee 2002).
Taking these figures and recognising that integration, subsystem design would make up the delta
between tank-and-engine development and full vehicle development, it can be seen that
development of a rocket system is relatively low-cost.
The more detailed configuration of the vehicle depends on staging, which is discussed in the
following section.
5.5.7.
ENABLING TECHNOLOGY: STAGING
Staging is most effective when each stage carries the same mass fraction – i.e., the ratio of the fuel
it burns to the mass it carries is identical (Fortescue, Swinerd and Stark 2011). From the rocket
equation, this implies that each stage should provide the same total delta-V (assuming similar
specific impulses), including penalties due to drag and gravity. Staging is most commonly
achieved by firing pyrotechnics, which destroy fastenings securing one stage to another.
Staging configurations may be either stacked (one on top of the other), in parallel or a
combination of both. The Saturn V, Space Shuttle and the Soyuz rockets are examples of each
configuration, respectively. There are benefits and drawbacks to each.
Stacked stages mean that each stage must have its own rocket engine. This adds mass to the stage,
thereby increasing the vehicle size. It also means the number of engines will be increased. Due to
the fact that engines are expensive, this can drive high launch costs. There are also risks inherent
in this design, most notably that either the premature firing of an engine, or the failure to ignite
whilst in flight, can be disastrous. However, the stacked-staging approach does offer the most
aerodynamic configuration, as there are no sudden changes across the geometry of the crosssection.
Parallel stages were used in early rockets to avoid the problems inherent to igniting an engine
during flight. Because the stages are in parallel, the same engine or cluster of engines can be used
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throughout the flight, saving mass (if this is the case, the engine must be altitude-compensating to
avoid poor performance in changing pressure conditions). In parallel staging, there is a greater
risk that a jettisoned stage may impact the adjacent stages, but this has been successfully
demonstrated many times. One drawback of parallel stages is that the frontal area of the vehicle
can be larger, thus resulting in higher drag.
In either design, there are of course risks that the pyrotechnics will fail. In this case, a stacked
rocket would be unable to fire its engine and continue the flight. The parallel staged vehicle would
be able to fire its engine, giving it the possibility of continuing the flight to some safer altitude
(assuming the failed stage is still safely attached). However, pyrotechnics are extremely reliable,
making this a small risk.
Staging benefits reduce with increasing staging events. More events also lead to increased risk of
a failure. The optimum number of stages, to balance increased complexity and risk against
payload ratio augmentation, is either two or three (i.e. one or two staging events).
Staging leads to inherent modularity of the vehicle. Modularity is an attractive feature that
improves re-usability and cuts costs related to hardware spares, along with overhaul time. Indeed,
the rocket could be fabricated in such a way that each of its stages can be broken down in to
elements that can be readily interchanged with any other stage, meaning that they would have
common hardware spares. This would be achieved by using common tank diameters.
5.5.8.
NOZZLE CONFIGURATIONS
The purpose of the nozzle is to obtain maximum velocity from the exhausted gas, by supersonic
expansion. This is most efficient when the exit pressure matches the atmospheric pressure.
Bell nozzles, as typically used for current-generation launchers, are designed for expansion at
only one pressure (and hence one altitude). This results in a decrease in performance of 1-5%,
along with a drop in thrust of up to 15% (Sutton and Biblarz 2001). This drawback is corrected by
carrying multiple engines, which are suited to flight at each staging event. However, engines are
heavy and expensive items. The higher the number of engines carried, the lower the total payload
fraction and the higher the incurred cost and complexity of recovery.
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Figure 30 Nozzle configurations and relative sizes (Huzel and Huang 1992)
There are numerous nozzle designs that allow for pressure compensation at multiple pressure
conditions. Among these, several designs include enhancements to the present day bell
configuration. Multiple ‘steps’ are used, to re-configure the nozzle to the optimum area ratio for a
given flight condition. These steps can either be built in to the nozzle (dual bell) or mechanically
added or subtracted during flight (droppable inserts or extensible nozzles).
Mechanical nozzles suggest a reduced reliability, as a failure of the deployment mechanism is a
real possibility. If this were to occur, the engine would provide insufficient thrust to reach the
desired altitude. Thus, these are not considered as potential candidates. A performance loss is
associated with a dual-bell design, due to separation at the bump and shorter than ideal nozzles at
each stage. Thus, adapting a bell nozzle to provide altitude compensation is non-optimal.
Nozzles using aerodynamic boundaries are able to alter their ‘geometry’ automatically. As a result they can operate efficiently at any given pressure, rather than a finite number of design conditions.
These include aerospike nozzles and expansion-deflection nozzles.
The expansion-deflection nozzle consists of an ordinary bell, with a plug suspended near the
throat. This diverts the flow along the nozzle walls and it expands against the pressure inside the
re-circulating region. The plug needs to be capable of sitting in the most hostile part of the engine
(i.e. near the throat) and would likely need to be cooled to prevent it from melting. This is a
relatively complex design.
The alternative and preferred configuration is the aerospike nozzle. This nozzle is fully pressure
compensating, thus offering the highest performance. There is reason to believe that it could be a
more lightweight solution, as it is physically smaller than a bell nozzle. The limitation of the
aerospike nozzle is that multiple nozzles cannot operate efficiently in parallel (this would disturb
the aerodynamic boundaries and thus diminish pressure compensation). Thus, a single nozzle
would need to be used. The risk associated with a single engine can be offset by using multiple
combustion chambers and sectioning-off flow-paths on the nozzle for each chamber.
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Figure 31 Pressure compensation flow pattern of aerospike nozzle (Sutton and Biblarz 2001)
Cost-wise, testing of aerospike nozzles has occurred in the past (the XRS-2200 is an example), so
the technology is not completely unknown and would not be experimental. Flight experience with
such nozzles is yet to be gained. The majority of the engine machinery is identical to that of
conventional engines, so there is little concern in terms of technical ability.
An additional benefit of using multiple chambers for the aero spike is that differential thrust could
be used for thrust vector control. This would be a great improvement on the heavy mechanisms
employed to gimbal bell-nozzle engines, offering higher reliability and improved vehicle
integration.
Operationally, the recovery of a single engine stage, perhaps with the payload, is simpler than
recovering multiple engines. It is also more cost effective to overhaul a single large engine in
comparison to many smaller engines – not least, because it will only require one test firing, rather
than many. Hardware commonality within the engine, for example the thrust chambers or turbopumps, will also improve operational costs, because the same machines and processes are used
multiple times.
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5.6.
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LAUNCH CONCEPTS
5.6.1.
EXOTIC PROPULSION
Selection of rocket propulsion systems must take into account the factors such as the interface
between the propulsion systems and the vehicle’s overall systems. In order to allow all subsystems to operate efficiently with each order, careful analysis and systems engineering was
required to identify compromises that would this possible, (SUTTON, George and Biblarz, Oscar,
2010). Examples of such compromises include elements of propulsion systems and its ground
support, costs, risks and reliability. Once the mission requirements were defined, different
methods of propulsion systems were analysed to match the propulsion requirements.
5.6.2.
ELECTRIC PROPULSION
The best available exhaust velocity of 4.5 km/s for chemical rockets must be increased in order to
achieve more ambitious space missions e.g. manned missions to Mars, (TURNER, Martin, 2000).
Electric propulsion is a concept that uses stored electrical energy to generate thrust. There are
several approaches this can be done. However, most electric propulsion concepts are not designed
for high thrust as their accelerations are too low to overcome the gravitational force of Earth
launches. Instead they are used for exploration and interplanetary missions where they function
best due to high Isp and long operating times.
ION THRUSTERS
This is an example of an electrostatic thruster which uses static electric fields to accelerate
directly propellant ions to very high velocities. Even though, there are different ways to accelerate
ions, all these methods make use of the high charge to mass ratio of ions in order achieve very
high velocities.
One of the main advantages of an ion thruster is that it can achieve high levels of specific impulse
thereby causing a reduction in the reaction mass needed compared to conventional chemical
rockets. However, this means there is an increase in the amount of power needed.
Figure 1 shows a schematic of an ion thruster where gas particles are bombarded with electrons in
a chamber leading to the gas becoming ionised. Plasma is discharged from the collisions and
flows two screens; one positively charged and the other negatively charged, this causes a potential
difference, V between the two screens which is maintained by a high voltage power supply. The
ions exit the system based on the formula below;
𝑣 =
Author: Charles Ofosu
2(−𝑞𝑉)
𝑚
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where q is the charge of the ion, m is the mass of the ion and ve is the exit velocity. Outside the
system, an electron gun fires electrons into the ion exhaust stream to maintain a net positive
charge, (TAYLOR, Travis S., 2009). Providing the thruster has a continuous mass flow rate, 𝑚̇ of
ions through it, the thrust generated will be given by (TAYLOR, Travis S., 2009);
𝐹
= 𝑚̇𝑣 = 𝑚̇
2𝑞𝑣
𝑚
Figure 32 Schematic of the Deep Space 1 ion thruster. (Image courtesy of NASA)
HALL THRUSTERS
Hall thrusters are another type of electrostatic thruster and make use of an electrostatic field to
accelerate xenon ions to high exhaust velocities with no grids used, (TAYLOR, Travis S., 2009).
Electrons are trapped by a strong radial magnetic field and swirl about the axis of the thruster at
the exit of an engine. Xenon gas is used the propellant gas which is fed in through a positively
charged electrode and is accelerated by the swirling electrons because of the potential difference.
On the European lunar mission, SMART-1, this technology is used as well as a number of GEO
satellites because of the high Isp (~1500 s) which uses only a few kilowatts of power (TAYLOR,
Travis S., 2009). .
The main drawback of hall thrusters apart from being expensive is that ions can impinge on
spacecraft parts and cause contamination problems in the high vacuum conditions of space
because of the wide spread angle of the ions in the plume of a Hall thruster.
5.6.3.
SPACE TETHERS
This type of propulsion uses long incredibly strong strings attached to the Earth and going all the
way up into space to change orbits of spacecraft, (COSMO, M.L. and Lorenzini, E. C., 1997).
This technology will reduce the cost of space transportation considerably in the future. Recent
advancements in carbon nanotube technology which have a theoretical strength of 60 GPa is
bringing this technology closer, (CORNWELL, Charles F. and Welch, Charles R., 2011).
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ELECTRODYNAMICS TETHER
An electrodynamic tether (EDT) is a very simplistic and low-budget idea that is based on the
principles of electromagnetism. The EDT system is consists of two masses interlinked by an
electrically conductive cable thus a direct current is produced through the tether when it moves
through the Earth’s magnetic field and this field subsequently exerts a force on the current causing the system to slow down, (CHRISTENSEN, Bill, 2001).
Addition of a source of power such as a solar panel illustrated in Figure 2 to the EDT circuit
causes the induced current to be overcome, thus reversing its direction. This produces thrust in the
tether which can be used to accelerate spacecraft, (CHRISTENSEN, Bill, 2001).
Figure 33 Electrodynamic Tether System. Image courtesy of Tethers Unlimited, Inc.
5.6.4.
ROTOVATOR
Rotovators are momentum exchange tethers rotating at high speed so their tips reach speeds of up
to ~3 km/s, (BOLONKIN, Alexander, 2006). The mechanism involves a spacecraft latching onto
the tether at one end and being accelerated by its rotation in one orbit.
The spacecraft can then separate from the tether after reaching its altitude by an exchange of
tether’s momentum and angular momentum with must be re-charged. Two or three rotovators can
be connected and used to transport goods from the Moon to the Earth by exchange of momentum
making this technology exciting, (BOLONKIN, Alexander, 2006). However, this cannot be built
yet as no current materials can withstand the tip speed generated by the high spin rates required by
a rotovator.
5.6.5.
SPACE ELEVATOR
Space elevators are a type of a rotovator but are powered by the spin of a planet. The tether will
be attached to the ground on Earth and at the other end in GEO, a counterweight is attached to
keep the tether in tension. This is because gravity would be strongest on Earth and the centripetal
force would be strongest toward the counterweight, mechanical payload can then be attached to
the tether to be sent into orbit.
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Space elevators have the potential to reduce space transportation marginal costs down to $220/kg.
The estimate for construction of the space elevator is around $6-12B (GOUDARZI, Sara, 2005).
The Edwards proposal aims to launch 2,000,000 kg per year into orbit and estimates all operating
and maintenance costs to be $20B. Comparing this proposal to the Skylon project, which has a
research and development costs of $15B and has only a 15,000 kg cargo capacity to 300 km
(HEMPSELL, Mark and Longstaff, Roger, 2009). However this would be mechanical payload
only.
Figure 3 shows an illustration of a Japanese company, Obayashi Corp.’s concept of a space elevator with plans for it to be operational by 2050, (MACK, Eric, 2012).
Figure 34 Obayashi Corp.'s space elevator concept (Image courtesy: Daily Yomiuri)
5.6.6.
MAGLEV SPACE TRANSPORTATION
Maglev Space Transportation is based on the same technology used on high speed trains. This
concept is based on magnetic levitation using superconducting electromagnets to levitate and
catapult a launch vehicle and payload at an inclination into orbit, (Logsdon 1998). Lack of friction
with the track would make the launch vehicle capable of accelerating to orbital velocities of ~ 9
km/s (5.6 m/s).
Current passenger train using maglev technology can achieve speeds of 373 mph however, for
launching spacecraft; the vehicle would have to a velocity of 600 mph. A recent online article
(NUSCA, Andrew, 2010) claims maglev technology would be capable of this velocity within the
decade. The peak acceleration that would be reached is 3g which is roughly the same as the Space
Shuttle’s which makes it suitable to carry human passengers. Other factors required to make this technology idea a reality would be a track length of 1609 km
in length which is currently achievable for cross-country transportation, (ZYGA, Lisa, 2012). A
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vacuum tube would also be essential with vents to allow compressed air in front of the spacecraft
to escape the tube. Any flight vehicle travelling at hypersonic velocity at sea-level would
experience high aerodynamic drag and sonic shockwaves so the vacuum is needed to avoided this.
A vacuum equivalent of an altitude of 75 km would be sufficient and additionally, the exit of the
vacuum tube must be elevated to about 20 km, (ZYGA, Lisa, 2012). Otherwise acceleration force
at the exit would be about 100g due to aerodynamic drag at the exit; otherwise it would be similar
to escaping the vacuum tube travelling at a speed of about 20,000 mph (ZYGA, Lisa, 2012) and
hitting a brick wall.
The required technology in order to make maglev space transportation a reality exists and is well
understood however the engineering scale is simply too large. The main company behind this
initiative, StarTram, plans to build a passenger vehicle capable using maglev technology to reach
orbit. With a time frame of 20 years, the construction budget without taking into account inflation
is $60B. However, an estimate cost/kg of $50 would make it really cost effective enough to justify
the initial investment capital, (Powell 2010). Comparison with Space Shuttle which had a
development cost of $170B and the International Space Station which has cost $150B to date
shows that this method of transportation has huge potential and should not be ignored.
StarTram has plans to make its Maglev space transportation reusable after every launch without
extensive maintenance. Their unmanned Generation 1 vehicle is estimated to cost $19B and their
passenger version, Generation 2 is estimated to cost $67B to develop (Powell 2010).
Figure 35 Maglev train to space. (Image courtesy of StarTram)
5.6.7.
SOLAR SAILS
The principle of a Solar Sail is based on the fact that photons from the Sun hit a particular area of
some kind of thin film and propel this particular structure along with the rest of its components by
imparting a small force, (BOLONKIN, Alexander, 2006). The force imparted on the sail structure
is given by,
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𝐹 = 𝜂𝑃𝐴 𝑐𝑜𝑠 𝛼
where η is the sail coefficient (~1.8 with film wrinkles), P is SRP constant at one astronomical unit (AU) from the sun, A is the surface area of the solar sail and A is the sun angle between the
surface normal and the sun line.
Figure 36 free body diagram of a solar sail.
The idea missions to incorporate solar sail as a propulsion method are those involving long
distances with carrying propellants. Solar sails offer a significant mass reduction compared to
conventional rocket propellants which leads to an increase in payload capacity. Solar sails are
ideal for high ΔV missions because of their good build-up of acceleration.
However, very large sails would be required in order to launch a significant amount of payload
which is required for the proposed design based on market research.
Figure 37 Typical payload carrying solar sail (Image courtesy: NASA)
5.6.8.
TECHNOLOGY REVIEW
Here is a table showing the rating system used in order to down-select.
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Table 9 Progress review of exotic propulsion, values range from 1 as the worst outcome to 5 being the best.
Technology
Cost
Environmental
Technical
Impact
Feasibility
Reliability
Health
Time
&
scale
Safety
Maglev space
2
2
4
2
3
2
1
3
3
2
1
1
2
4
3
4
4
2
2
3
3
2
2
2
4
5
2
3
5
3
transportation
Space
elevator
Electric
propulsion
Space
Tethers
Solar sails
Since space elevators/tethers would be stationary, it would be vulnerable to objects in space such
as space debris, asteroids and satellites and also be a possible target for terrorists therefor this
technology was abandoned. Electric propulsion is suitable for deep space explorations and
interplanetary missions however it did not meet the mission requirements to launch about 30T to
LEO due to low thrust. Electric propulsion will be pursued for GTO delivery applications.
However, the maglev space transportation will be further pursued in the subsequent chapter.
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5.6.9.
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RAMP LAUNCH SYSTEM DESIGN
Figure 38 – Ramp launch system from When Worlds
Collide (BALMER, Edwin, 1951)
Figure 39 – Lunar Launch Ramp Design (BESCHIZZA, Rob, 2009)
The concept of using a ramp assisted launch can be dated back as early as 1951 to a movie, as
shown in Figure 38. In this film the ramp was built against of the mountain and the spaceship is
launched from the sea level. If the mountain is replaced by a steel structure, as shown in Figure
39, it could allow the proposed concept to be built in any location. The structure has a high level
of reusability and maintenance is a relatively simple process. If the ramp is equipped with a
propulsion system like the catapult launch in an aircraft carrier or induced propulsion it could
contribute to the launch ΔV required and provide a controlled acceleration. Thus reducing the
amount of fuel required to be carried.
The proposed design might however be difficult to construct, as it needs to achieve a terminating
altitude of 6 km to 10 km high to be beneficial, as shown in Appendix 12.2, Figure 85. This will
increase the cost of construction and the difficultly of maintenance which requires large amount
of resource for construction and will not be environmentally friendly. An initial analysis has
concluded that 363 Mt and substantially more steel than the largest steel structure in world, the
Bird Nest as shown in Appendix 12.2, Figure 86. Thus, it has been decided that this proposal
cannot be built with current or near term construction materials and methods. If a linear ramp
design is chosen it could reduce the material needed as the ramp curvature creates centrifugal
force, which create addition loading.
An alternative method is a design similar to the movie, as shown in Figure 87. The highest
mountain on Earth is the Mount Everest with 8.84 km and the highest cliff near equator is at the
northern Pakistan with 4.6 km. This will decrease the amount of structural steel required.
However, the environmental impact of this proposal will be huge.
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5.6.10.
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MOUNTAIN TUNNEL LAUNCH SYSTEM DESIGN
Figure 40 – The Maglifter Booster Concept (MANKINS, John, 1994)
NASA also has a concept that takes advantage of natural structures. It involves using the
electromagnetic catapult system through a mountain. It has been estimated to have a development
cost of $2B. The proposal is to launch at the mountaintop located above 4.3 km with a speed of
970km/h and acceleration of 3 g’s, as shown in Figure 40 (Mankins 1994).
The project can be constructed with current tunnel technology. The tunnel boring machine (TBM)
has been constructing tunnel with construction starting from the peak surface with 30° downward
or starting at the ground level with 42° upward. This information is from the email conversation
with Herrenknecht International Ltd employee, one of the leading tunneling companies. It also
depends on the location of the mountain, as the price varies with the type of material that must be
drilled through and its length, as suggested in Appendix 12.3. The advantage of this method is
similar to those proposed in Chapter 5.6.9. In addition, the tunnel could be near-vacuum sealed
resulting in an increased exit velocity due to nominal drag losses. However, these advantages do
not outweigh the price of the structure in comparison to a conventional launch facility.
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5.7.
INITIAL RESEARCH
LAUNCH VEHICLE HARDWARE
5.7.1.
PROPELLANT MANAGEMENT
Propellant management consists of controlling 3 propellant aspects, the pressure, the temperature
and the flow.
Pressurising the propellant is essential for achieving a high specific enthalpy in the combustion
chamber. This is then converted to kinetic energy, providing thrust, using the nozzles. This high
pressure can be achieved using two well established methods, a pressure-fed delivery system or a
pump-fed delivery system.
In a pressure fed delivery system a separate tank is filled with high pressure inert gas. This tank is
attached to the propellant tank via a high pressure regulator. This gas squeezes the propellant out
of its tank and into the combustion chamber at the same pressure as the gas. It is typically used for
in-orbit propulsion but has legacy uses in ballistic missiles. A pump-fed delivery system relies on
pumps to pressurise and move the propellant to the combustion chamber. The pump typically
burns part of the fuel to produce the mechanical energy required. This is the most widely used
pumping method for launch vehicles. The advantages and disadvantages of each system are
shown in Figure 41
Pros
Pressure-Fed Delivery System
Cons
Simple, no mechanical moving Achieves lower pressure, large
parts, cheap
pressure
vessel
required,
heavy
Pump-Fed Delivery System
High pressure achievable
Complex, many moving parts,
expensive
Figure 41: Advantages and disadvantages of different fuel delivery systems.
Flow management is required to ensure propellant is fed to the engine at the right time, pressure
and direction. A typical flow control scheme uses the following components:

Fill/Drain Valves – Used to fill/drain a tank prior to launch.

Pyrotechnic Valves – For liquid propellant tanks to initiate the flow into the system.

Pyrotechnic Isolation valve – Used to shut off an engine.

Pressure Relief Valves – Vent propellant in case the system is over pressured.

Check Valves – Control the flow direction and prevent back-flow.

Redundancy – Typically there is a duplicate set of fuel lines that allow bypassing of faulty
valves.
Author: Richard Fields
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In conclusion it would be best for the proposed launch vehicle to use a legacy pump-fed system.
This will help improve the vehicle reliability, decrease costs and decrease development time. An
example of one type of pump-fed system is shown in Figure 42: Example of turbo-pump used in
Centaur LH2/Lox upper stage (R.Wertz and Wiley J. Larson 2008)..
Figure 42: Example of turbo-pump used in Centaur LH2/Lox upper stage (R.Wertz and Wiley J. Larson 2008).
5.7.2.
GUIDANCE, NAVIGATION AND CONTROL
Guidance, navigation and control, GNC, systems are essential for ensuring the safest and optimal
trajectory is achieved during launch. For launch vehicles, GNC typically uses onboard sensors
and computers which are complimented by ground tracking to ensure the system is giving the
correct feedback. Traditionally groups of spinning mass gyroscopes are used for sensing. These
have the benefits of being simple but may be susceptible to strong vibrations. Ring laser and
fibre-optic gyroscopes improve accuracy and reliability so these should be pursued.
In recent launch vehicles, a GPS receiver is becoming the standard sensing method. Commercial
application accuracy is approximately 3 m at sea-level and 15 m to 100 m (R.Wertz and Wiley J.
Larson 2008) in LEO due to shortened triangulation capabilities. There are restrictions on
commercial GPS receivers that limit its sensing capability to 18km altitude and 515 m/s (ACA,
1993). The launch vehicle will vastly exceed both of these so a military grade GPS receiver must
be used instead, which requires a CoCom/ITAR export certificate and is very costly. An
alternative may be to reverse engineer the receiver to remove the CoCom restrictions which will
increase R&D costs but vastly reduced production costs and susceptibility to government
restrictions.
In conclusion, all of these GNC systems will be combined using a Kalman filter to provide
stabilised feedback to the trajectory control system. A COTS solution should be sought if a low
quantity of vehicles will be produced. If a large quantity of vehicles is to be produced it would be
preferable for a custom solution to be designed and produced to save expensive purchasing costs.
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COMMUNICATION AND DATA HANDLING
The launch vehicle must constantly communicate with the ground control station. This is to
ensure correct telemetry data and course changes are being applied during the flight. If there is an
issue with this process it could lead to the launch vehicle posing a threat to structures or people
downrange.
Traditional communication methods involved an array of dish antennas that would require
mechanical slewing to maintain pointing. Modern launch vehicles require a larger throughput of
digital data with the rise of advanced tracking techniques and live HD video feeds. Use of phased
array antennas that transmit through communication satellites rather than directly to ground
stations may provide high data rate capabilities.
A high level study on launch vehicle communications concluded that the inclusion of a TDRS-Ka
band antenna provides the minimum size and weight for the necessary communication throughput
of a modern launch vehicle (WELCH, Bryan and Greenfield, Israel, 2005).
5.7.4.
ELECTRICAL POWER
To calculate the amount of power necessary depends on the hardware systems implemented. The
hardware systems that require electrical power during launch are described below:

Avionics – This system includes the GNC and communication. These are relatively low
powered systems, in the order of 10’s of watts. They require sustained power during the whole launch process with a voltage tolerance of 10%.

Pyrotechnic and flight termination – This system involves the separation of flight stages
using pyrotechnics and the controlled termination of the flight to ensure range safety. It
requires very large currents but only over a very short duration.

Thrust vector control – This system involves the movement of nozzles to control the
vehicles trajectory. Traditionally gears and hydraulic actuators are used but more recently
fully electric systems have been favored due to their simplicity and weight savings.
An example of a traditional launch vehicle is the Ariane 5 Heavy launch vehicle which is capable
of launching 21,000kg to LEO. The power requirement for this vehicle is 4 kWh over a period of
6 hours with a peak consumption of 1100W (BARDE, H et al., 2002).
In the case of a launch vehicle there are some power methods that are impractical such as solar
panels due to the high velocities and vibrations. A radioisotope thermoelectric generator (RTG)
for launch vehicle power is also discouraged due to the risk of contamination in case of launch
termination or failure.
The only viable alternatives are fuel cells and batteries coupled with capacitors for high current
delivery. Fuel cells have the benefit of being more compact and lightweight but would require
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fuel to be carried onboard and are currently very susceptible to damage from long duration
vibrations (PAUL GEORGE 2008). Batteries are the traditional method and can be designed to
resist high acceleration and vibration. Typical battery chemistries and their specific energy are
shown in Table 10.
Table 10: Performance of battery technologies for space use (Fortescue, Swinerd and Start 2011).
Careful consideration must be taken when designing the electrical subsystem of the launch
vehicle. Some of the power transmission lines will be very long so to limit Ohmic losses a higher
transmission voltage, around 60V, is preferred. This leads to an issue that is present in all launch
vehicle and satellite design.
At low pressures but not quite a vacuum electrical arcing may occur in high voltage components.
This can lead to failure of the equipment and potentially failure of the launch. This is known as
the Paschen breakdown. To mitigate this, the electrical components should be housed with
venting holes large enough to alleviate internal stress of pressure buildup but small enough to
ensure that during operation, a high voltage electrical system is not exposed to a pressure where
Paschen breakdown may occur (LUX, Jim, 2004).
In conclusion for the proposed launch vehicle it would be best to use a COTS Li-Ion battery
solution. There are many providers so it is believed a competitive price could be achieved. This is
preferred to in house development because research into advanced Li-Ion technology is not a
business priority.
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5.8.
5.8.1.
INITIAL RESEARCH
LAUNCH ABORT SYSTEMS
LAS
Traditionally, Launch Abort System (LAS) has appeared as towers containing solid propellants
attached on top of the crew capsule, ready to pull the crewed capsule away from a falling rocket at
the launch pad or during early descent.
The Apollo and Mercury programmes used LAS however the Gemini programme used ejection
seats which provided little chance of survival. The Soyuz T-10-1’s LAS has only been required in
a real emergency once; the crew of four landed safely four miles away during this event.
5.8.2.
SPACE X LAS
Space X plans to incorporate Super Draco hypergolic rocket engines on later versions of its
Dragon spacecraft for crew transport to LEO, re-entry, descent and potential landing control.
Recent announcement of the success of the Grasshopper programme testing shows this is feasible
and not far in the future, the Grasshopper rocket achieved flight to a height of 40m in 29s then
landed safely back onto the launch pad (THISDELL, Dan, 2013). However, the Super Draco
engine’s primary purpose is to act as Space X’s launch abort system on the Dragon capsule.
The system is funded by NASA’s Commercial Crew Development (CCDev) 2 and is currently in development. During launch abort, eight Super Draco engines are expected to fire for 5 seconds at
full thrust providing 530 kN axial thrust and the engine has a transient from ignition to full thrust
of 100 m/s.
The benefits of this engine is include it can be put through a series of throttling ranges which
allows redundancy and also has additional dual redundancy in all axes. The engine can also be
restarted multiple times.
When compared to the traditional tower LAS, this system does not require jettison after the 1 st
stage of ascent and failure in this key sequence usually ends the mission because the flight profile
is not designed for carrying the tower along into orbit. In terms of reusability, this system can be
incorporated into the capsule. The capsule would return with the engines on board if not required
in an emergency so can be used on future launches. In the longer term and with improved certified
technology, the capsule would be capable of landing back on Earth by its own propulsion means
or on the Moon or Mars. All these factors make it an ideal proposal as the launch abort system of
the design.
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5.9.
INITIAL RESEARCH
SPACECRAFT HARDWARE
5.9.1.
ATTITUDE AND ORBIT CONTROL
The AOCS is dependent on 3 things, using sensors to collect information about the vehicles
attitude, interpreting this information through a control system and applying attitude adjustments
via
actuators.
AOCS ATTITUDE SENSING




“Looking out the Window” - These sensors use external objects to gauge the vehicles
attitude. They typically “look” at the Earth, the sun, or the stars. The Earth and the sun
provide orientation about only 2-axis of motion, the stars can provide orientation in 3axis.
Gyroscopes - As the vehicle rotates around an axis perpendicular to the gyroscopes spin
vector; a torque will be produced causing the gyro to move. Measuring this movement
can provide the new orientation. This system has the benefit of being simple but requires
knowledge of the initial orientation as a reference and will need calibration. There are two
modern forms of gyroscope called the ring-laser and fiber-optic gyroscopes. These offer
similar/better accuracy with greater reliability than a spinning mass.
Magnetometer - This measures the Earth’s magnetic field in 3 dimensions, a compass measures it in 2. The measurements are compared to a map of the Earth’s magnetic field. They are only useful in LEO due to the magnetic field reducing in strength further away
from Earth.
GPS - Sample principle as ground based GPS. As orbital height is increased this method
becomes less accurate.
AOCS CONTROL SYSTEM
The AOCS control system is primarily a computer that can receive signals from sensors and
guidance inputs either from ground control or onboard personnel. It is then required to process
the sensor data to get an accurate estimate of the vehicles attitude. Using guidance inputs it will
apply adjustments via actuators such as momentum wheels, thrusters or main engines.
5.9.2.
ELECTRICAL POWER
Providing electrical power for a spacecraft requires analysis of the spacecraft mission profile. One
of the aims of this project is to design a spacecraft capable of carrying and supporting humans
while in orbit. Typical mission profiles of such spacecraft range from a few days in orbit to allow
transition to a destination such as the ISS to a fully self-supporting lab that may be in orbit for
weeks or potentially months.
An example of a vehicle that was adaptable to different mission scenarios was the Space Shuttle.
The maximum mission duration without docking with the Mir/ISS was approximately 2 weeks.
The maximum peak power consumption of the Space Shuttle was 35 kW and continuous power
Author: Richards Fields
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consumption was 21 kW. This vehicle has 6 passengers which is the same as the target
specification.
The three primary methods for providing in-orbit power are PV-cells, fuel cells and a RTG. A
RTG will not be pursued due to the high costs, low health and safety and possible damage to the
environment in case of failure in launch or re-entry. Fuel cells were the primary source of
electricity for the Space Shuttle. To provide the power listed above required 3 fuel cells, each
weighing 115 kg and having a total volume of 0.42m3 (Dumoulin 1988). This is quite small and
light but it must be remembered that for fuel cells, the fuel must be carried too. This takes up an
additional 0.6m3 and weights 140kg including tank weight.
PV-cells can provide sustainable power during illumination and charge batteries for use during an
eclipse. Typical space grade solar cells have an efficiency of 16% but future PV-cell technologies
may increase the efficiency to 30%.
Author: Richards Fields
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5.10.
INITIAL RESEARCH
HEAT SHIELDING
5.10.1.
DECELERATION
One of the most important features of the re-entry procedure is the deceleration of the vehicle.
There are a number of contributing factors that decide the amount of deceleration the vehicle is
capable of. Without considering these factors seriously, the vehicle may be under-prepared for
deceleration, and this has a critical effect on the success of the vehicle’s re-entry phase.
The vehicle itself plays a constant part in the drag force the vehicle experiences during descent.
The surface area of the base of the vehicle is maximised within the constraints the overall rocket
applies, and this provides the best re-entry surface area that can be gained. This surface, and
incidentally the entire re-entry vehicle shape, is designed to provide the best possible drag
coefficient (based on the angle of re-entry), either maximised or at an appropriate compromise
with horizontal drag to allow the vehicle to achieve stable flight characteristics later in the descent
profile whilst achieving a sufficient deceleration rate earlier in the profile. The drag on the vehicle
(or other bodies attached to the vehicle) is the only way a falling vehicle can decelerate (unless it
is equipped with rocket engines, which can then be used to provide thrust). The contributing
factors all assist in maximising the drag on the vehicle.
The forces the vehicle comes under during descent can be summarised into two main parts, i.e.
gravity and drag force. The gravity obviously has its own variables. Whereas it can usually be
approximated that acceleration due to gravity is about 9.81m/s2, this number applies only at the
surface of the Earth. Gravity obviously varies altitude, because it varies with radial distance from
the centre of mass of the object, in this case the Earth. The actual relationship between them is as
follows:
Equation 1 Relationship between gravity and radius
𝑔
𝑟
=
𝑔
𝑟
Using this relationship, a value for gravitational acceleration at varying altitudes can be
calculated.
The other factor in the descent is drag force. Drag force, as the name may suggest, works
adversely to the direction of travel. Drag force takes into account atmospheric density, and this
again is a variable based on altitude. The calculation for drag is:
Equation 2 Drag force
𝐹
Author: Samuel Vereycken
1
= 𝜌𝑉 𝐶 𝐴
2
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These two equations can be used to find an equation of motion based on these dynamic variables.
For the purposes of this report, the equations have been implemented into a simulation created in
Microsoft Excel. However, due to the nature of the problem, an issue arises in that the variables
have a circular dependency. In other words, some of the variables rely on other variables that
either directly or indirectly relies on the primary variables in question. This would normally
prevent such equations from being calculated in a continuous system. This problem is averted
with the use of time steps. By using small time steps, the error can be kept minimal, and the
equations can now be used by referring in series to initial values, providing final values that
decide the initial values for the next iteration. Within the simulations created for the descent
profile, the time steps have been set at 0.1 seconds, except for the final step from which the time
step is calculated from the descent velocity, acceleration and distance from the ground. This
produces a total value for descent time.
A major consideration during descent is the heating the vehicle undergoes. This is not majorly
caused by the friction between the vehicle and the Earth’s atmosphere, but by the compression of the air ahead of the vehicle at supersonic speeds, causing the air to turn to plasma. This not only
causes lots of heating, but also opens up another problem in the form of plasma degradation
which is inevitable and prevents complete re-usability of any surface that comes into contact with
this air.
Atmospheric density is the defining feature of the amount of friction and plasma degradation
encountered. Therefore, it can be assumed that the vehicle encounters little heating in the upper
atmosphere, but as the vehicle reaches the heavier atmosphere the amount of heating increases
rapidly. This means creating a stable heat-resilient surface or layer is a priority of the vehicle
design, without which the vehicle is rendered incapable of surviving re-entry and will burn up
somewhere in the lower atmosphere.
Heating is a dangerous aspect to descent, and failure of heat shielding is far from impossible. A
previous case of failure due to just this problem is the Space Shuttle Columbia incident. This
wasn’t due to design failure, but the take-off of the vehicle suffering some complications that
weren’t fully comprehended until the Shuttle returned. Part of the foam from the fuel tank detached during take-off and struck the heat shield on the left wing of the Shuttle. Consultation
between NASA and external advisers did not throw up major concerns, and the issue was
dismissed as non-critical. However, the damage meant the heat shield was incapable of protecting
the vehicle. Temperatures in the left wing rose, and sensors in the wing stopped giving feedback.
Temperatures continued to rise to critical levels, and the result was catastrophic failure and
destruction of the entire shuttle on the morning of the 1st of February 2003 (CHEN, Yng-Ru,
2003). The vehicle exploded and the entire crew were lost. This tragedy illustrates the very real
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importance of the heat shield, and how it can fail if every aspect of the flight is not carefully
monitored and considered.
Figure 43 The Space Shuttle Columbia at take-off with evidence of the
problem that caused the disaster (CHEN, Yng-Ru, 2003)
There are two main variables that need to be considered to develop an appropriate heat shield
capable of resisting the amount of heating the leading surface experiences. The first is the
maximum temperature the shield may reach. This would determine whether the shield starts
charring on the way down. The danger this would pose is that if the shield is burning away, the
effective heat resistance that the shield has is greatly reduced. As the unburnt depth of the material
reduces, the minimum requirements for depth are reached and passed and the material is no longer
capable of absorbing and radiating energy at a rate greater than or equal to the rate at which it is
being heated.
5.10.2.
HEAT SHIELD DESIGN
There are many different heat shields used today. They come in varying shapes and sizes, and the
material they’re made from varies a lot. Heat shields rarely come as a module that can be bought “off-the-shelf”, but rather they have to be developed for the vehicle in question and so are usually
unique to the vehicle they’re designed for.
Depending on the vehicle shape, the heat shield has certain parameters that restrict it and that it
must meet to function correctly. The first parameter is the nature of the drag through the
atmosphere. As an object passes through the atmosphere at high velocities, the air particles ahead
of the leading surface are hit by the surface and bounce off. These then make contact with air
particles behind them, and this impacting causes the air to heat up and turn to plasma. This is an
unavoidable product of the motion of the vehicle through the atmosphere at high velocities.
However, the repercussions of this depend upon the shape of the object. This is because an object
at high velocities experiences either an attached or a detached shockwave. If an object has low
drag, the shockwave is situated on the surface because the air particles are not rebounded further
away from the surface, and this means the surface faces a lot of heating. If, however, the object
has high drag, the particles are rebounded heavily and so the shockwave is situated away from the
surface, and so the heating effect is considerably lower.
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The traditional capsule shape uses a shallow-curve domed base, and this has a fairly high drag
coefficient. This means the shockwave is detached, so the amount of heating the heat shield faces
is due to the friction the surface experiences and the barrage of air particles. Less-used shapes,
such as the Shuttle, feature areas with low drag coefficients to enable stable flight lower in the
atmosphere, so these areas move more towards attached shockwaves. Examples would be the
wingtips and the nose cone. These areas have to endure a lot more heating than others, and so
have to be made of a more heat resilient material. The material chosen is a carbon-carbon
composite. The benefit of this material is that it maintains good strength when heated, and can
withstand these high temperatures.
Typically tiles are used when making a heat shield. There are many benefits associated with using
tiles over a single constructed piece. The first benefit is that, with individual tiles, should any area
of the heat shield be damaged, the affected tiles can be removed and replaced. A single piece
would be rendered useless if an area was damaged, and so the entire piece would be wasted.
Therefore waste is kept minimal, as is cost through manufacture and labour. Another benefit is
that heat shield is easier to manufacture. By producing arrays of tiles, even though they have to be
crafted to fit to their exact grid location in the heat shield, they can be produced in batches or as
replacement pieces and are easier manufacture in workshops using smaller ovens and tools. A
single piece heat shield would require considerably larger ovens and tool arrays that are far more
expensive and harder to use. This would drive up manufacture costs with more labour required
and more expensive tools used.
Figure 44 Discovery's heat shield photographed from the ISS (KAUDERER, Amiko,
2012)
Other benefits of using tiles include easier transportation of initial and replacement parts, simpler
construction onto the vehicle, and easier integration with regards to moving parts (hatches etc.)
and future vehicle modifications. Transportation costs are kept lower because tiles can be
packaged into far more efficient spaces than a single piece heat shield, and fewer tiles required in
a shipment can be reflected in a smaller transport vehicle used. Construction onto the vehicle is
marginally more difficult in one sense in that it requires precise work to align and stick the tiles
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into their corresponding locations, but they can be handled individually and fixed individually,
meaning work is overall a lot more manageable than a large single piece that must be affixed to
the surface of the vehicle. The integration possibilities associated with tiles are very useful. Tiles
can be conformed around protrusions by constructing appropriately shaped tiles in those
locations, and they can be used around and on hatches and flaps because they do not hold a rigid
shape together, but align to the surface they are attached to. Modifications to the vehicle are also
more simply integrated, with only the tiles in the areas affected being modified or removed,
whereas a single piece heat shield would have to be removed and replaced with a new design
constructed to fit the most recent modifications made. Tiles obviously save a lot of cost that
would otherwise be associated with vehicle modifications for this reason.
The different materials currently available, as well as the obvious development possibilities, mean
that many different configurations of heat shield material are possible. Obviously it goes without
saying that more recently developed materials have superseded their older counterparts, and the
ever-changing field means advances are on-going. For this reason, finding a reliable value for cost
estimation is not very easy. Most companies do not publish their own cost breakdowns, and as
heat shields are generally not sold without contracts and project development, individual analysis
and quotations are needed to find a useful assessment.
5.11.
RE-ENTRY TRAJECTORY
A small range of angles are required in order to achieve successful atmosphere re-entry; the limits
of this narrow re-entry region are determined by the spacecraft’s trajectory, its rate of deceleration
and aerodynamic heating.
Figure 45 Boundaries of re-entry corridor (Image: From Aerospaceweb.org)
5.11.1.
BALLISTIC ENTRY TRAJECTORY
This is the conventional re-entry trajectory used by capsule vehicles because it is the simplest
and possibly the safety method. Very little aerodynamic lift (L/D<1) is generated as the
vehicle falls into the atmosphere under gravity and drag. The drag forces slow the vehicle’s velocity down to a velocity that allows parachutes to be deployed.
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Once the vehicle enters Earth’s atmosphere, there is no control of the vehicle until it lands
however it is possible to calculate and predict its ballistic trajectory, (SCOTT, Jeff, 2005). In
the past, manned space capsules used this trajectory including Mercury and Gemini, Soyuz
missions still re-enter with a ballistic trajectory in land in Siberia to a degree of 5km.
E.g. Manned space capsules of Mercury and Gemini flights used ballistic entry trajectory to a
splashdown at sea. Soyuz capsules still use the ballistic entry trajectory to touchdown on land
in Siberia.
Figure 46 Artist impression of ballistic trajectory (Image: NASA)
5.11.2.
GLIDE TRAJECTORY
Space vehicle glides through the atmosphere at an angle of attack of 40° generating
aerodynamic lift. The Space Shuttle had a high L/D ratio allows it to travel further than the
ballistic trajectory by gliding and circling around an airfield until it is able to land.
The main advantage is the reusability factor as the pilot can control the vehicle’s trajectory and can land on a runway. However there is a drawback which is the heat shielding required
would be expensive and lead to an increase in the overall weight of the vehicle, (SCOTT, Jeff,
2005).
Figure 47 Artist impression of a glide trajectory (Image: Aerospaceweb.org)
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5.11.3.
INITIAL RESEARCH
SKIP RE-ENTRY
Skip re-entry features both the ballistic and glide trajectories where the space vehicle enters
the outer atmosphere initially to generate drag to slow it down itself then it pitches up in a
controlled climb to exit the atmosphere again. This is repeated a few times until the space
vehicle reaches an acceptable speed to make a final ballistic re-entry. Space designed to do
this have an L/D ratio between 1 and 4 thus it is within their capability, (SCOTT, Jeff, 2005).
Figure 48 Skip re-entry schematic
Advantages
Disadvantages
The vehicle can achieve a greater entry Considerably higher aerodynamic heating
range
than
either
ballistic
trajectory.
or
glide required is required. Heavier shielding
would be required to project the vehicle
because the friction heat absorbed during
the skip entries grows at a high rate.
Can help dissipate huge amounts of heat Requires precise guidance performed on
compared to faster descents.
autopilot.
Never been used on manned spaceflight.
CONCLUSIONS
Based on the information about re-entry trajectories, a final decision was made between designing
a capsule or a lifting body vehicle for the human payload.
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5.12.
INITIAL RESEARCH
ENVIRONMENTAL IMPACT - PROPULSION SYSTEM
The propulsion technology has been enhanced over the history of space exploration. Apart from
traditional fuel, there is a few different propulsion systems been developed. Each fuel has their
own unique specification and requirement. However, the factor of the environmental issues and
sustainability cannot be neglect when selecting the technology. There is a brief environmental
comparison with the existing reusable space vehicle in Appendix 12.4.
5.12.1.
PETROLEUM
Kerosene is a combustible hydrocarbon liquid derived from petroleum. It is an elaborate and
energy intensive production process as it is a crude oil derivative. The process of mining from an
oil drilling can cause harm to the ecosystems and require large amount of energy for refining. A
much more refined petroleum, such as RP-1, is needed for rocket fuel. Engines that burn RP-1
have a limited operational lifespan caused by the residue produce from the fuel (CUNNINGHAM,
Jeff, 2012), where more engines will be need for the project lifespan. In addition, it often releases
toxic matters to the air and water. In the case of accident can cause long-term effect on the
ecosystems and public health. The last but not the least, the transportation of the fuel increases the
energy use, since most of the mining is from the seashore.
Kerosene is highly toxic to human and animal life. For limited exposure can cause irritation to
skin, eyes and mucus membranes (GREEN, Malachi Lloyd, 2012). Long-term exposure can cause
advanced toxicity symptoms or even fatal in the case of prolonged exposure to high concentration
(GREEN, Malachi Lloyd, 2012). The combustion of kerosene releases heavy concentration of
toxic particulate pollution, such as nitrogen dioxide, sulphur dioxide and carbon monoxide
(GREEN, Malachi Lloyd, 2012). The kerosene smoke emissions can also cause severe respiratory
diseases inhaled by humans or animals.
One of the main greenhouse gas produce by kerosene is carbon dioxide, like any other fossil fuels.
Greenhouse gases are directly linked to global warming, which raise the average temperature of
Earth’s atmosphere and ocean, creasing the rise of sea level. In addition, particulates such as
carbon monoxide can be immediately dangerous to life with a concentration of 1200ppm or
greater (ENVIRONMENT AUSTRALIA, 2011). Although carbon monoxide is not considered as
a greenhouse gas, it elevates the concentration of methane and ozone in the atmosphere. It will
however oxidise into carbon dioxide.
Kerosene has its advantage including cost and relative safety. It is a non-corrosive fuel, thus it is
safe to store for a long time and easy to be stored. It can be kept in a better condition by storing in
a controlled condition away from rain and sunlight.
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5.12.2.
INITIAL RESEARCH
HYBRID
Hybrid propulsion is a new form of propulsion method that combines properties from both liquid
and solid fuelled. Taking Virgin Galactic as an example, where part of the propulsion method is
using hybrid motor. The project is to delivery tourism to the suborbital flight. It produce
approximately 0.8 tonnes of carbon dioxide which is about 70% less compare to business class
flight from London to New York per passenger (VIRGIN GALACTIC, 2010). This is however a
more sustainable fuel in compares to kerosene.
There is a big draw back with this propulsion method, the combustion produce black soot. The
soot can be washed back to ground by rainfall in the lower atmosphere; however it accumulates in
the stratosphere (BABONES, Salvatore, 2012). According to a simulation study published in
Geophysical Review Letters, the global warming is 140,000 times more cause by stratospheric
soot that associated to carbon dioxide emissions for simulation of 1000 suborbital flight per year
(ROSS, Martin et al., 2010). In addition, this can increase polar surface temperatures by 1°C and
reduce polar sea ice by 5%-15% (NATURE, 2010).
5.12.3.
SOLID
Solid propellants are the simplest and oldest in history. It could be considered as the most reliable
propellants, as the fuel is very stable, easy to store and require no complicated pumps.
Taking the Space Shuttle as an example, the SRB NASA uses will produce chlorine, chlorine
oxides, nitrogen oxides, hydrogen oxides and alumina during combustion (AVR ENTERPRISE,
2005). However, it depends on the material mixture of the SRB production. As it has a significant
local exhaust deposition more likely to have a larger environmental impact compare to other
propulsion system, where NASA and the USA military produce about 725 tonnes of chlorine and
where natural sources produce about 75,000 tonnes and private industries produce about 300,000
tonnes chlorine. In another word, the SRB rate of usage will need to increase about 40 times to
match 10% of the private industries (WILLIAMS, Marcel, 2001). The deposited of the chlorine in
the stratosphere is a source of destruction of the ozone layer. However, the contribution of the
space industry might not be the main source of pollution. The space shuttle produces about 76.8
tonnes of greenhouse gas, carbon dioxide as a bi-product through each launch. Over 28 years the
contribution of NASA using Solid Rocket booster generate about 42,000 tonnes of carbon
dioxide, which is about 0.0003% of the automobile’s contribution in the USA (O'NEILL, Ian,
2009), where the auto industry is responsible for about 50% of the greenhouse green emission.
Furthermore, the Space shuttle launch produce about 56 tonnes of aluminium oxide, 35.2 tonnes
Hydrogen chloride and about one-third every 1,000 tonnes of Solid rocket fuel will become
hydrochloric acid (STARS WITH A BANG, 2009), where the aluminium oxide dust particles
form nuclei with water vapour cause rain and when the rain absorbs hydrochloric acid to produce
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acid rain (AVR ENTERPRISE, 2005). The pH value can be as high as 1 near the launch pad
initially. The pH value will drop as moving away from the launch pad. Thus, the ground station of
the Kennedy Space Center will monitor where the predict acid deposited with the prevailing wind
in different altitudes in every launch. As John Pike, president of Global Security stated "The
hydrochloric acid can pit the paint on your car if it is too close to the launch site" (SMEATON,
Zoe, 2005). In addition, NASA found "reduction in the number of plant species present and
reduction in total cover", which will have an impact nearby water lagoons and their wildlife
(SMEATON, Zoe, 2005).
5.12.4.
CRYOGENIC
The typical cryogenic propellants fuels are liquid hydrogen, methane and fluorine. These fuels
require a very low temperature for storage. For example liquid hydrogen requires being stored at 253°C (NASA, 2012). Thus, it needed to be handled with extreme care from evaporation and
boiling off. In addition, insulation from all source of heat is essential, for example air friction
during flight through the atmosphere and the sun when in space. This is because the fuel can
expand rapidly once heat is absorbed. Also it can leak through minute pores in weld seams
(NASA, 2012).
Liquid hydrogen has light with the lowest molecular weight and is an extremely powerful rocket
propellant. It has the lowest molecular weight and burn with extreme intensity of 3136°C (NASA,
2012). It has a highest specific impulse or efficiency in relation to the amount of propellant
consumed of any other rocket propellant.
Figure 49 - Comparison of emissions for kerosene and hydrogen of equal energy content (ZON, Nout Van, 2012)
The commercial aviation industry had been undertaken research of engine using liquid hydrogen
and kerosene. The experiment shows hydrogen emits no carbon dioxide and 80% less nitrogen
oxides, as shown in Figure 49. This will not cause harm to the atmosphere. Hydrogen had been
showing a great performance in compare to others. It requires three times less mass then kerosene
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to produce same amount of energy, but the drawback is that it requires four times the volume of
kerosene, as shown in Figure 50.
Figure 50 – Experimentation done with V2527-A5 engine with hydrogen (ZON, Nout Van, 2012)
The current method is using natural gas and electricity to produce liquid hydrogen and all current
method produce carbon dioxide. Thus it might be more expensive than using kerosene, which
does not provide any economic benefit. However, the price of kerosene is estimated to crossover
around 2037 with the oil prices rising and advances in hydrogen technologies. In the future there
will be more innovated and sustainable method available, which is photolysis, thermolysis and
biomass gasification (ZON, Nout Van, 2012). These technologies are expected to be available in
the year of 2040.
There are a lot of advantages of using liquid hydrogen, but due to it special properties it require to
store in special installation to prevent boiled off. Take NASA as an example, the space shuttle
used an insulation foam named chlorofluorocarbon (CFC) in 1997 will cause ozone depletion
(SMEATON, Zoe, 2005). Due to environmental regulation NSAS had replaced with more nonfreon-based foam and believe to be more environmentally friendly foam. This is by using this fuel
might cause other environmental impact in another part of the chain.
5.13.
ENVIRONMENTAL IMPACT – SPACE DEBRIS
5.13.1.
RISK OF IMPACT
There are two types of debris associated with the space environment. The first is naturally
occurring objects such micrometeorites and the second manmade objects. The second type of
debris can be split into three further sections that are mission-related debris such as rocket bodies
and telescope lens caps, surface degradation debris such as paint chips and debris from on-orbit
fragmentation (R.Wertz and Wiley J. Larson, 2008).
Space debris poses significant risks to space missions due to their high orbital velocities. If debris
in LEO is travelling in a retro-grade orbit collides with a satellite travelling in a pro-grade orbit
the resultant impact velocity could be as high as 15 km/s. The resultant energy of this impact
depends on the debris size and mass. A comparison of the size of natural and man-made space
debris is shown in Figure 51.
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Figure 51: Accumulated surface area flux as a function of particle diameter in 2010 (Ley, Wittmann and
Hallmann 2009)
Debris of a size between 1cm and 10cm is generally not observable and cannot be shielded
against due to size and speed. These will result in the catastrophic failure of any satellite they
collide with. The chances of such a collision occurring to a spacecraft in LEO with a cross
sectional area of 10m2 over a 10 years lifetime is approximately 1/500.
5.13.2.
REDUCING DEBRIS POLLUTION
All space users have a responsibility to include debris mitigation techniques in their mission
profiles (R.Wertz and Wiley J. Larson 2008).
Since the commercial development of space applications there has been a steady increase in
manmade orbital debris. The manmade contribution to space debris is shown in Figure 52. It
shows that rocket bodies account for approximately 18% of all debris. Debris associated with
launch vehicles is further increased when it is considered that the fragmentation debris can be
created by exploding fuel tanks that were not properly vented or surface degradation of fuel tanks
due to collisions.
Figure 52: Monthly number of catalogued objects in Earth orbit by object type (NASA 2011).
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Mitigation of debris production from the proposed launch vehicle can be achieved by ensuring the
majority of separations, such as the fairing or rocket stages, are conducted well before orbital
velocity is achieved. Separation systems should tend away from using pyrotechnic systems to
more passive systems such as band clamps. Bolt catchers should be used where applicable. Any
rocket bodies that do make it to LEO should be properly vented or conduct an idle burn (R.Wertz
and Wiley J. Larson, 2008). This is to prohibit tank over pressurisation that may build up over the
years leading to an explosion.
5.14.
CONCLUSION
5.14.1.
SUMMARY OF RESEARCH AREAS
Table 11 - Target Specification
Payload Capacity (LEO)
Passenger Capacity
Target Altitude
Fairing
Overhaul Period
$/kg to LEO
Vehicle Reusability %
Vehicle Lifecycles
Target
30,000t
6
300km
4m x 20m
1 month
$1000
90%
200
From the onset of the project a varied and in depth amount of research was ensued. This helped to
guide the design and bring up areas that may have needed further investigation. The project was
broken down into sizable allotments of interest and ideas were discussed. The main ideas coming
out of this research phase gave insight into the current space market, launch technologies the
limiting factors i.e. g loads, radiation to name a few, as well as heat shielding technologies and
Fuel options. Initially a market study was conducted on current and future markets which were the
main drivers for later decisions in terms of the overall context of the work. Also of particular
attention were the environmental impact, reusability, technical feasibility and viability.
5.14.2.
DOWN-SELECTIONS
Market research in the Inception Report led to the target capacity of thirty metric tonnes to LEO
with a price of $1000/kilogram solely to compete with the Space X systems. The large cargo
capacity was developed to account for some of the potential uses of space as outlined in previous
work such as the development of space infrastructure and industrialisation of space. On discovery
of the Skylon program which was being proposed to carry thirty people, further investigations
were pursued into the potential demands for such a large passenger capacity. Market research
showed a demand for space tourism but a lack of willingness to pay greater than $1M. The initial
Author: Group
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research suggests carrying 30 people to match Skylon or 6 people to replace the recently retired
space shuttle.
From cost per kg estimate, the 30-person module would assume a ticket price of $1M and proved
to be too high for personal interests such as tourism. In addition, no government or industry needs
to send that many people in one mission. However, it was pursued as a possibility should anyone
have a need for it in the future. Primary focus would be on the 6-person, in addition to some
mechanical payload to make 30T in total. Possible application of the system to deliver to the
moon, mars and perform round the world flights also considered. On orbit assembly is efficient
for the moon and mars. Round the world flights would be prohibitively expensive for a spaceship,
but the technology could be adapted from a space-plane and made cheaper for this.
Limitations involved in space travel to the humans seems plentiful, though many are present more
focus was given to the ones which seemed most significant these were limited to g loading,
radiation and environmental provisions. With various technologies outlined and choices for these
systems suggested for moving forward.
Launch options were discussed and research and design considerations were done to develop a
potential launch facility which would involve the use of the maglev technology which would be
employed within a tunnel to be constructed into a very tall mountain this was later down selected
for a more traditional vertical launch as the launch savings would not have been significant.
Fuel Types included liquid chemical propellants, solid and hybrid propellants, from the research
with solid propellants having a 10% reduction in specific impulse as compared to liquid
propellants. Though hybrid propellants tend to be safer to handle due to benign outputs which was
appealing their disadvantages led focus to be given instead to a combination of hydrogen, LNG
and RE-1 environmental impacts of these was then researched to support this as a positive choice.
Layout options considered for the fairing and fuel tanks included the Carrying packet Rocket,
Feeding packet Rocket which then led to the novel approach developed by the group to use a
feeding Packet with donut shaped tanks which was developed with the idea of storing the fuel and
oxidizer more efficiently than traditional cylinder shaped rockets however was not selected going
forward due to the complex separation of tanks and high production costs due to complicated
shape. More research was done as proceeding with the feeding Packet as research suggests it
could lower the environmental through reduced emissions and reduction in the amount of fuel
required. Stage recovery and engine recovery were highlighted as the next areas of investigation.
Potential Propulsion methods saw research from conventional to futuristic, including air
breathing, traditional rocket launchers as well as space planes. The outcomes of this were that the
Air Breathing technology is not designed for vertical launch and unit costs would be prohibitive
as they negate the potential financial benefits of using a simpler vertical rocket design as
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compared to a space plane. Additionally, analysis of the engines shows that air-breathing engines
that may be required would be heavier and more expensive – they are also less well understood,
making this risky. After thoroughly researching each of these areas some conclusions were made
about the benefits and drawbacks of air breathing systems, space planes and potential flaws in the
Skylon system. Further investigations were carried out into optimising traditional Rocket
technologies as these presented the most technological feasibility and years of trusted use. A
reusable space-plane is difficult from a thermal shielding perspective, inefficient in the mass it
takes to orbit, such as extra shielding or wings, and more costly, because of the added design
effort.
Author: Group
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6.
INITIAL DESIGN PHASE
6.1.
VEHICLE MASS ESTIMATIONS
6.1.1.
FIRST ESTIMATE OF TOTAL DELTA-V REQUIRED
The mass of a rocket is driven by the quantity of fuel required to achieve orbital velocity, and
overcome the force of aerodynamic drag and acceleration due to gravity. Drag and gravity can be
accounted for as additional velocities, resulting in the total delta-V that must be acquired by the
vehicle.
Penalties for drag depend on the vehicle size. A massive vehicle will be accelerated less by drag
than a smaller one, thereby resulting in a smaller delta-V. For a mid-range rocket mass, this is
0.2km/s. The delta-V due to gravity is the same for all vehicles following similar flight
trajectories, and is the integral of the component of acceleration due to gravity acting along the
flight axis – this is taken to be approximately 1.1km/s (Fortescue, Swinerd and Stark 2011).
𝛥𝑉 = 𝑉 + 𝛥𝑉 + 𝛥𝑉
Where: 𝑉 =
, 𝐺 is the gravitational constant and 𝑀 is the mass of the Earth. This results
in an orbital velocity of 7.73 𝑘𝑚/𝑠, for a circular orbit at 300km.
First-pass estimates for delta-V associated with Drag and gravity, 𝛥𝑉 and 𝛥𝑉 , are used. These
are,
𝛥𝑉 = 0.2 𝑘𝑚/𝑠 and 𝛥𝑉 = 1.1 𝑘𝑚/𝑠. (Fortescue, Swinerd and Stark 2011).
It is therefore seen that, in total, enough fuel must be carried to achieve a 𝛥𝑉 = 9.1 𝑘𝑚/𝑠.
Estimation of this nature is used because the acceleration due to drag on the vehicle can only be
calculated if the mass of the vehicle is known. The acceleration due to gravity is both dependent
on time and flight angle, which will also be unknown until a detailed trajectory is plotted.
6.1.2.
ESTIMATION OF FUEL MASS
Using the Detla-V requirement, fuel masses can be predicted using the rocket equation. This is
modified to allow for a structural mass element, proportional to the fuel, 𝑠. This is also estimated
based on existing rockets and text-book figures to be 10% (Fortescue, Swinerd and Stark 2011).
The ratio of fuel mass to dry mass 𝑅 is first calculated, where dry mass includes the mass of the
payload (𝑀 ), the mass of the structure (𝑀 ), including the engines initially.
Author: James Roper
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𝑚
𝑑𝑉
= 𝑒𝑥𝑝
𝑀
𝑔 ×𝐼
𝑅=
−1
𝑀 =𝑀 +𝑀
𝑀 = 𝑠 × 𝑚
𝑚 = 𝑅 × (𝑀𝑝 + 𝑀 )
𝑚 = 𝑅 × (𝑀 + 𝑠 × 𝑚 )
𝑚 =
The 𝐼
𝑅×𝑀
1−𝑅×𝑠
is taken from tabulated properties of the fuels that emerged from the feasibility study,
with modifications for non-ideal losses of 10% (Huzel and Huang 1992).
𝑑𝑉 is an equal fraction of the total 𝛥𝑉 for each stage. For equal specific impulses in each stage
(an approximation), this gives the optimum mass solution as each stage has the same ratio of fuel
to ‘payload’, where the payload is the mass that it is carrying (including the ‘upper’ stages).
Table 12 Vehicle take-off mass predictions for different numbers of stages, in tonnes, for a 30T payload.
Fuel and Specific Impulse
Stages
LNG (305s)
RP-1 (296s)
HYDROGEN (380s)
1
NOT POSSIBLE
NOT POSSIBLE
NOT POSSIBLE
2
1524
1790
595
3
1122
1271
516
Table 13 Payload fractions achieved by each fuel for different numbers of stages
Fuel
Stages
LNG
RP-1
HYDROGEN
1
NOT POSSIBLE
NOT POSSIBLE
NOT POSSIBLE
2
1.97%
1.68%
5.04%
3
2.67%
2.36%
5.81%
The data and supporting references showed a clear indication that there was indeed a benefit to
having three stages – that is, an improvement to the payload fraction of between 15% (hydrogen)
and 35% (RP-1). Hydrogen is seen to dramatically out-perform its competitors, with more than
double the payload fraction.
To perform a full comparison, the volume of propellant is the key parameter. A vehicle’s cost is proportional to its size. Thus, the volumes of oxidiser and fuel were also calculated, based on the
stoichiometric chemical mixture ratio and the density. It was initially expected that this would
Author: James Roper
| Page 78 of 262
reflect poorly on hydrogen, which has a very low density (67.8 kg/m3), however it was found that
the propellant volumes were relatively equivalent.
Table 14 Fuel and Oxidiser volumes in each stage of a 3 stage vehicle, for the 3 fuel options, and accompanying
chart displaying this data.
LNG
Stage
RP-1
HYDROGEN
3
2
1
3
2
1
3
2
1
Fuel Volume
28.4
95
318
18.3
63.9
223
70.7
182
471
Oxidiser Volume
44.8
150
501
46.2
161
562
33.6
86.7
224
Fuel
Volume
Oxidiser
Volume
0
200
400
600
800
1000
Oxidiser and Fuel volumes per stage, m^3.
At this stage, it becomes clear that hydrogen is the better candidate. As previously discussed, it is
the most environmentally compatible fuel, producing only water in its exhaust. It is also the most
widely available and is projected to reach the same cost as gasoline within the coming two
decades. The above calculations have shown that the vehicle using hydrogen has no drawbacks in
terms of volume, which would drive up the vehicle cost. The volumes for the three fuels, LNG,
RP-1 and LH2 are 1137 m3, 1074 m3 and 1068 m3 respectively. Additionally, a payload fraction is
achieved for a three stage rocket that will out-compete RP-1 and LNG fuels by 126%.
Hydrogen is thus pursued for more detailed analysis, in terms of both the engines and the tank
structure. Details of the calculations performed can be found in Appendix 12.4.
Author: James Roper
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6.1.3.
INITIAL TANK SIZING
Based on 10% dry mass fractions, the tank masses were estimated. The material chosen at this
stage was Aluminium-Lithium alloy 2195 which had the following properties Young’s Modulus (E) – 71 × 109 N/m2, density (ρ) – 2865 kg/m3 and maximum allowable tensile stress (σtu) –
609×109 N/m2. The maximum operating pressure used in these calculations was 2.34 bar which is
based on the Space Shuttle’s External tank since it was planned to use the same fuel and oxidiser.
This pressure was doubled as a safety factor.
The tank wall thickness was calculated using thin-walled pressure vessel formula for a cylindrical
section and the spherical ends (WERTZ, James R. and Larson, Wiley J., 2005) ,
𝜎=
(𝑐𝑦𝑙𝑖𝑛𝑑𝑟𝑖𝑐𝑎𝑙)
𝜎=
(𝑠𝑝ℎ𝑒𝑟𝑖𝑐𝑎𝑙)
Where σ is the allowable stress, p is the maximum expected operating temperature, r is the tank radius and t is the wall thickness.
Furthermore, the respective volumes were calculated for 3 lengths of tanks so the mass could be
tallied using the formulas below, (HUZEL, Dieter K. and Huang, David H., 1992);
𝑊 = 𝜌2𝜋𝑟𝑙 𝑡 (𝑐𝑦𝑙𝑖𝑛𝑑𝑟𝑖𝑐𝑎𝑙)
𝑊 = 𝜌2𝜋𝑟 𝑡 (𝑠𝑝ℎ𝑒𝑟𝑖𝑐𝑎𝑙 𝑒𝑛𝑑)
𝑇𝑜𝑡𝑎𝑙 𝑊 = 𝑊 + 𝑊
Results are presented in Table 36 of chapter 12.6 in the Appendix. The initial calculations
estimated a single tank without thermal protection system and fuel pumps, etc. would weigh about
1000 kg. This value is later used in the section Structural Design and Analysis.
6.2.
LAYOUT OPTIONS AND CONFIGURATION
At this point it has been decided to pursue 3 different fuel options. These options are the most
commonly used and give us the greatest leeway when studying the vehicle layout. In this section
the layout of the 3 primary components will be discussed. These are the payload fairing, the
engine and the fuel tanks. Where applicable an aerospike nozzle has been used to represent the
engine.
It has been decided that the payload fairing will be 4m in diameter and 20 long. This is to
accommodate Space Shuttle sized payloads such as space station segments. An additional fairing
has also been proposed that is 5.5m in diameter and 16m long to accommodate larger diameter
payloads such as the SpaceBus series by Thales Group. This will only be noted as a future
consideration and not designed at this stage.
Please note that at this stage none of the designs will incorporate an aerodynamic body. Only the
rough shape has been proposed.
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6.2.1.
FEEDING PACKET
The packet feeding design was first proposed by Mikhail Tikhonravov in
1947. It consists of a central support platform with the engine attached at
the bottom. The 4 first stage tanks and the 2 second stage tanks are
mounted on the outside of the central support platform in a hexagonal
configuration. The final third stage is mounted on top of the support
platform; it is obscured from view in Figure X. The payload fairing is
mounted at the top.
Advantages:

High payload mass fraction due to staging.

Fuel tank extension possible for higher payload masses.

Single nozzle reduces production costs.

Subsequent tank stages full of fuel at separation points.

Lighter tank stages simplify re-entry and reusability.
Disadvantages:

Complex pumping system required for cross-feeding.

Multiple stage separations introduce high failure risk.
Figure 53: Feeding Packet
Rocket
6.2.2.
CARRYING PACKET
Traditionally the carrying packet design is where each stage carries and
consumes its own fuel with no cross-feeding. The layout shown in figure X is an
Angara 3A by Khrunichev Research Centre.
Advantages:

Simple pumping as no cross-feeding required.

Booster stages can be used as independent rockets.

Good vehicle heritage and known reliability.
Disadvantages:
Figure 54: Carrying
Packet Rocket

Multiple engines increase the production cost of each booster stage.

Low payload mass fraction.

Complex reusability process.
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6.2.3.
FEEDING PACKET - DONUT TANK
This is an original design proposed by the group. It has been proposed with
the idea of storing the fuel and oxidizer more efficiently than traditional
cylinder shaped rockets. The Donut Tank employs a common bulkhead design
as shown by the splitting of the bottom tank.
Advantages:

Efficient fuel packing to reduce structure size and weight.

May provide a lower drag coefficient due to reduced structure size.
Disadvantages:

Complex tank separation procedure.

Long fuel lines from top tank to main engine.
Figure 55: Feeding
Packet - Donut Tank

Expensive production of unusual shaped tanks.
In conclusion the feeding packed design has been selected due to its simplicity and high payload
mass fraction. This will reduce the amount of fuel required thus reducing the vehicles
environmental impact through reduced emissions. A single nozzle will be used to reduce vehicle
construction costs. This saving comes at a price of introducing a complex pumping procedure that
requires the shut-off and opening of different stage valves during the launch phase. Having
multiple separation systems also poses a substantial risk to the launch phase as if separation does
not occur, or occurs at the wrong moment; the vehicle may be thrown off course and may
experience structural failure. For this design to be considered reusable, further analysis into the
stage recovery and engine recovery methods must be conducted.
6.3.
OPERATION
The amount of operation needed is heavily depending on how the space vehicle is design.
However, the operation needed compare to the past is significantly reduce as technology
improving constantly. The expected future is replacing the labours to computers and robots. This
will allow the reduction with labour force and human errors. Nevertheless, there are elements
require human to decide.
6.3.1.
MISSION PLANNING
The current planning process is one of the most time-consuming processes. The process could not
be replace and require the same amount of problem solved as the past. The planning before the
mission is known as strategic planning. The objective is to develop and end-to-end profile
meeting the mission requirement. This also includes all abort modes and contingencies of the
mission.
Author: Richard Fields / Norman Tang Fai Ng
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Flight-day planning is a mission profile will be developed based on the strategic planning on the
day of launch. Flight-day planning process focuses on the on the process of launch and re-entry.
This is to account for variables, such as the weather. The planned mission profile will be
constantly modified to the according constraints. The last but not least, the on-orbit re-entry
planning and coordination are the re-entry trajectories, which will be planned as part of the
strategic planning. This will be continually updated base on the weather and traffic conditions.
6.3.2.
GROUND OPERATIONS
The ground operation might be the key for each successful launch. This is because this operation
contains all assembled, maintained, repaired, serviced, fuelled and launching of the vehicle. This
process require large amount of labour force and could be the most time-consuming before the
launch process.
The ground segment needed to complete a series of tasks in the mission preparation stage to
ensure it had the capability. The existing infrastructure and mission system will be compared with
the project requirement. This will allow the preparation and execution of the operation process
can be carry out smoothly. A typical ground operation team structure is shown in Figure 56.
Figure 56 – The decision flow of ground operation team structure (Fortescue, Swinerd and Stark 2011)
The designs of the spacecraft should have its own dedicated manual for customers. In another
word, the payload will be design based on the spacecraft own specification. Thus, the payload will
be assembled and checkout before the vehicle integration. The process of the vehicle integration
begins after all spaceship components are ready from manufacture. This stage of the process is
very important to ensure all checks are made against the design. The processes require a control
environment, clean room. This typically includes receiving and processing payloads, preparing
mission cargo and testing for launch vehicle compatibility. The vehicle will then be move to the
launch platform when the whole assembly and checkout process is completed. The vehicle will
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then be prepared for launch operations. This stage is the last stage of the process for ground
operation. The fuelling process will begin after all final checks are complete on the day of launch.
The launch control centre will then authorise the vehicle to take off and monitoring process will
take over by the antenna from ground station around the world, shown in Appendix 12.5 for more
information about requirement and limitation of antenna.
Furthermore, for reusable launch vehicles will have an additional process, which is the recovery
and refurbishment operation. This is different to the vehicle with conventional design, Evolved
Expendable Launch Vehicle (EELV). EELV might not be design to have any reusable capability.
However, there are some EELV with the design of the SRB usage, where it can be reusable. The
process of the recovery of element will be further discussed in Chapter 6.3.4. The elements will be
bought back to the facility for inspection after the recovery process. There might be section of the
elements need to be refurbish before additional launch.
6.3.3.
GROUND OPERATIONS ARCHITECTURE
Figure 57 – Ground operations architecture (NASA, 2006)
The cost of the ground operation architecture is highly dependent on the management process
control and complexity of the flight-to-ground interfaces, an example of the ground operation
architecture shown in Figure 57. Thus the simpler the operation the quicker the cycle time can be,
such as the removal of the use of hazardous and toxic commodities. The use of commodities
required expensive infrastructure and special suit for personnel operation under safe conditions.
This overall lead to a lower logistics support cost.
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Figure 58 –Operation Concept (NASA, 2006)
An efficient operation concept is allowing the ground segment to produces routine launch for both
human and cargo payload, shown in Figure 58 as an example of operation concept. This required
safe and efficient ground based process by minimizing the number and complexity of the flight
elements. Thus, the operation require is minimize to allow quicker turnover time, therefore reduce
the cost of the launches.
6.3.4.
VEHICLE ELEMENTS RETRIEVAL
The biggest problem is to retrieve elements of the vehicle. The elements normally are not design
to land or glide back where the facilities are. This is because such design tends to be heavier than
tradition ones. The weight of the system is always the driving factor of launch vehicle design. The
elements retrieved are normally the SRB and the human capsule. The traditional method is still in
use, which landing in the sea with an aid of parachutes. However, in the future this will change.
Space X is developing a reusable rocket prototype name grasshopper. The main goal of the
development is to fully use the rockets and spaceships.
NASA had retrieval teams and custom made their two ships for the SRB retrieval operation. Each
ship is design to retrieve one booster and require 10 crews, one retrieval supervisor and observer
and nine crews for the SRB retrieval operation. The team will conduct visual assessment of the
flight hardware when the SRB is located. The parachutes will be brought aboard before any
further action. The SRB is design to buoyant with the aft skirt of 33m under water. Then two
small boats will be deployed with nine retrieval divers. The first team will install an Enhanced
Diver-Operated Plug (EDOP) in the nozzle of the booster, where the EDOP is 7 m in length and
weights of 500 kg. While the second team check the EDOP and the aft skirt before the de-water
process with the air hose connected with the ship. Finally, the SRB will fall horizontally and tow
back to Port Canaveral for disassembly and refurbishment process (DISMUKES, Kim, 2003).
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The biggest disadvantage is the requirement of an extra operation. The STS mission carry by the
Space Shuttle in 1983 cost about $53.36M per launch for the retrieval if the SRB (RASMUSSEN,
Anker, 2012). This could be cheaper if the employment of a private salvage company due to the
rate of launches by NASA. However, private salvage will not be an ideal solution if there are
more frequent launches. Creating own operation team will be a more reliable approach. This
return to the same question addressed at the beginning of this chapter, will the grasshopper
technology be available in time of the project. Alternatively, design the element with the
capability to land on lands.
6.4.
STAGE RECOVERY
A key aspect of this project is to produce not only a launch vehicle but one that is also reusable to
some degree. To achieve this requirement it has been deemed necessary to recover and reuse as
many flight stages as possible.
The preliminary design proposal involves the use of 3 stages to achieve close to optimal launch
efficiency. The 1st stage consists of 4 fuel tanks; the 2nd stage consists of 2 fuel tanks; and the 3rd
stage consists of the payload, 1 fuel tank and the engine. This section will investigate the possible
methods of stage recovery and to quantify their reusability.
The tanks have been estimated to cost USA$2M each. If the total cost of recovery and
refurbishment minus the economic impact of the additional weight comes out less than the
production cost reusability would be considered worthwhile from an economic standpoint.
Another aspect of reusability is the environmental impact. In the scope it has been decided to try
and achieve a vehicle turnaround time of 4 weeks, this would lead to 13 launches per year. If the
tanks are not reusable this would require the production of 91 tanks, if they can be made
recoverable and reusable there is potential to produce just 7 tanks that would last for the year and
possibly longer.
6.4.1.
1ST STAGE RECOVERY OPTIONS
As described in the propulsion section, the optimal moment for 1st stage tank separation is 120
seconds into the launch, at an altitude of 60km and with a vertical speed of 1800m/s and a
horizontal speed of 1300 m/s. Momentum will carry the tanks an extra 30km resulting in a peak
altitude of 90km and carrying the tank approximately 150km downrange. The fuel tank sizing has
been estimated to be 31m long and 3m in diameter with a Haack series nose cone.
6.4.2.
PARACHUTE
There is only one vehicle in history that employed stage recovery; this was the Space Shuttle that
used a parachute system to recover each SRB. Approximately 4.5 km above sea level the pilot
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chute would be deployed to ensure the SRB was falling tail first. At 2 km above sea level the
drogue chute would deploy followed shortly by the 3 primary parachutes which decelerate the
SRB to 22.5 m/s allowing a relatively soft landing in the ocean. Subsequently a vessel will find
the SRB, seal the nozzle and pump out any water inside. The SRB is then towed back to shore for
a full inspection and refurbishment. This process is approach to reusability is simple but not cost
or time effective. For the Space Shuttle program it was estimated to cost USA$500M per year to
run recovery and refurbishment operations (NASA 1988) the vast majority of this cost was for
maintaining the recovery fleet. An image a SRB splashdown is shown in Figure 59.
Figure 59: Space Shuttle SRB Parachutes and splashdown.
With relation to the design proposal it would be possible to utilise the exact same method. The
recovery cost by ship would be vastly reduced since the tank has no nozzle to let water into, the
port for fuel transfer could be mechanically sealed via a non-return valve. Since pumping would
not be required it would be possible to pick up the tanks directly from the water, either by boat or
preferably a helicopter such as the Sikorsky S-64 Skycrane. This helicopter can be rented for
$5000 per hour (Erickson Air-Crane 2012) or bought outright for a longer term solution.
6.4.3.
PARAFOIL
Another possible method would be to utilise a parafoil rather than a parachute. A parafoil differs
from a parachute in that it produces substantially more lift. It achieves this by using the passing
air to inflate a non-rigid cellular structure into an aerofoil shape. This method would allow a
guided decent to a more desirable location, such as on landing on a traditional runway. There has
already been some development into this area for NASA’s X-38 although this is for the landing of
a manned vehicle and not a fuel stage. An example of how this might look has been provided in
the Figure 61.
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Figure 60: A pair of parafoils carrying a Space Shuttle SRB (Shoer, 2009)
The system is believed to be a lightweight and relatively cheap solution. It would minimally
impact the launch procedure. Reliability would be essential if the system were to be used over
land as a chute failure could cause the booster to land in a populated area causing collateral
damage.
Further investigation reveals that a parafoil cannot be fully deployed until it is travelling at subtransonic speeds. This is due to supersonic shock waves forming on the surfaces causing the
parafoil to collapse (NICOLAIDES, John and Tragarz, Michael, 1971).
6.4.4.
WINGS
Another possible solution would be to attach wings to the fuel tanks. This could provide a vastly
extended range compared to the parafoil approach due to the solid design of the wings. A wing
provides good stability as high winds have little effect on the wings shape in comparison to the
parafoil. This would make the system safer and more reliable thus making it easier to certify for
landing on land. An example of a winged cylinder would be the Pegasus launch vehicle shown in
Figure 61.
Figure 61: Pegasus Launch Vehicle by Orbital Sciences Corporation
There are also issues with a winged recovery system. If the wings are deployed during launch they
will be creating lift and drag as the rocket accelerates. Depending on which way round the wings
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is, they could be pulling against the central body inducing a large stress on the attachment points
or they could be pushing against the central body causing issues when separation is required. One
solution to this would be to store the wings during lift off and separation and then deploy them for
descent. This system is a heavy approach and would impact the launch vehicles efficiency
requiring larger tanks to carry the same payload weight. An example of this approach is shown in
Figure 62.
Figure 62: Russian Baikal Booster with Rotatable Wings
6.4.5.
POWERED RETURN
A newer concept in the space industry involves the 1st and 2nd stages carrying additional fuel that
can be burned through the main engine to slow their decent and possibly provide a soft landing.
One of the most recent USA patent applications by Jeff Bezos et al details “Sea Landing of Space Launch Vehicles and Associated Systems and Methods”. An image of the general system is detailed in Figure 63.
Figure 63: Proposed Landing at Sea of Spent Booster Stages (P.BEZOS, Jeffrey, 2010).
This system is not as readily applicable to the current design proposal as the 1st and 2nd stages
have no engines and thus cannot prove thrust intrinsically. Additional rockets, such as those used
in the rocket-assisted take off of aircraft, could be strapped to the base to provide this thrust
needed for this manoeuver. This method could provide a very rapid refurbishment time since the
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stage will have landed in a predefined location with known conditions compared to landing in the
sea or desert where conditions may vary. Traditionally these rockets are not re-usable and costly
to produce, typically reserved for military applications.
6.4.6.
1ST STAGE RECOVERY CONCLUSION
Table 15: A quantifiable overview of the proposed recovery systems for the 1st stage fuel tanks. Values range
from 1 as the worst outcome to 5 being the best.
Technology
Cost
Environmental
Recovery
Technical
Impact
Time
Feasibility
Reliability
Health &
Safety
Parachute
2
4
1
5
5
4
Parafoil
4
4
3
5
4
4
Wing
2
5
5
3
4
5
Powered
2
2
4
1
2
3
Tallying the results from Table 15 it is possible to see that the parafoil and wing recovery options
score the highest, with both getting 24. Both of these options have their individual benefits and
drawbacks. A compromise may be to pursue a para-wing approach with a semi-rigid parafoil but
this was found to not decrease the weight compared to a wing nor increase the L/D ratio of a
parafoil (NAESLETH, Rodge, 1970).
The final decision is to pursue the parafoil approach due to it being a lightweight approach which
does not affect the launch aerodynamics drastically. It is also a proven technology that may be
adapted or combined with other technologies to improve the potentially longer recovery times.
6.4.7.
2ND STAGE RECOVERY OPTIONS
The fuel tanks of the 2nd stage pose a much larger challenge for reusability. This is due to the
release velocity being 4000 m/s, approximately 50% of orbital velocity at an altitude of 150km,
again being carried up by its momentum to approximately 300km. This poses significant
problems as it must effectively be designed as a re-entry vehicle, requiring heat shielding.
Once the re-entry phase has been completed, exactly the same recovery method as used for the 1 st
stage fuel tanks will be applied.
6.4.8.
3RD STAGE RECOVERY OPTIONS
The 3rd stage may be recovered in a similar fashion to the 2nd and 1st stages but it may prove that
the cost of carrying the additional weight of the recovery system to orbit far outweighs the cost of
replacing the tank. Even if this is not the case the tank will require special refurbishment due to
the high thermal gradient involved around the structure.
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A novel solution to this may be to use the tanks as in-orbit structures, since it has already been
carried all the way to a full orbit. Below are some of the proposed solutions:
Retrofit into a space station – This idea has been inspired by Skylab, the first space station built
by the USA. It was launched using a Saturn V rocket where the upper stage was not filled with
fuel but left dry instead. It was fitted with basic support systems which were added to with 2
subsequent launches. This proposal may not be feasible due to the upper stage of the proposed
launch vehicle being required to carry fuel which would greatly complicate the retrofitting
procedure.
Propel upper stages to the moon – This proposal involves boosting the spent upper stages to the
moon to be used to form a lunar space station and elevator. A rough estimation has put the upper
stage tank mass to be approximate 1500 kg. Assuming a Hohmann transfer would suffice to put
the tank into lunar orbit and an Ion Jet space tug was used to do so it would require an additional
500kg of fuel. If a dedicated lightweight Space Tug and Fuel were attached to the upper tank it
would minimally impact the potential payload to LEO. A in depth analysis of this proposal has
been conducted by Dave Hunt of Embry-Riddle Aeronautical University (HUNT, Dave Randall,
1998).
Once the tanks are in orbit they could be grouped together to form a large hexagonal body that
surrounds a vacuum core. This vacuum core would provide the space for constructing a habitable
section to house scientists, this section would have to be sent in a dedicated launch.
Further consideration has been taken to this idea. The space environment outside of LEO contains
deadly radiation that requires heavy radiation shielding to be carried and even then will not fully
protect astronauts. To reduce the requirement of carrying heavy long term radiation shielding it
has been proposed that a lunar elevator be built, which would carry dirt from the moon’s surface up to the space station and fill the tanks. This would be conducted in a manner similar to that
proposed in by Ranko Artukovic (ARTUKOVIC, Ranko, 2002). The cable length for this
compared to an Earth based space elevator would be significantly less due to the decreased
specific gravity of the moon. This results in the cable stress being significantly less and could be
made of current materials such as Zylon or Kevlar.
The final stage of this proposal would involve sending the launch vehicle engines, at their end of
life, to the moon space station via Ion-Jet Space Tug. In recent years there has been increasing
evidence that the moon contains water on its surface (AMOS, Jonathan, 2009), this could be
mined (STONE, William, 2011) to produce the LOX/LH required by the launch vehicle engines.
The engines could be used to turn the lunar space station into a long term interplanetary
transportation vehicle for human travel to Mars and beyond.
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6.5.
6.5.1.
CONCLUSION
DOWN-SELECTIONS
Design reviews were held within the group to assure the satisfaction of the design requirements
and working efficiently. Through effective communication between the group members and
documentation of the informal meetings as well as formal, the programme is as planned and meets
the requirements.
Based on initial calculations, a 𝛥𝑉 requirement of 9.1 km/s was required. This allowed the fuel
masses to be estimated for RP-1, Methane and Hydrogen. The results showed that a 3 stage rocket
was the best option key result and another key result that emerged was that Hydrogen fuel was the
most suitable fuel. Hence more detailed analysis led to initial tank dimensions to be estimated as
well as engine options.
Aluminium-lithium alloy 2195 had been chosen as the best material candidate for the tank
structure simply because of its high specific strength and better cryogenic ductility compared to
other alloys. Initial tank mass calculations performed gave an estimate of about 1000 kg per tank
excluding any fuel pump management devices and insulation. Another key result from the initial
mass calculation was a 6% tank: fuel mass fraction, this led the optimisation of the overall
vehicle.
In order to achieve complete reusability, the vehicle’s layout configuration should not be over complicated thus the feeding packet design was chosen for its simplicity as well as high payload
mass fraction. This design also met the environmental, cost and technical feasibility requirements
the group had set out to meet. Further investigation was carried out on stage recovery methods to
complement the feeding packet configuration with the aim to achieve complete reusability.
After comparing the different stage recovery methods in terms of cost, environmental impact,
reliability, recovery time, technical feasibility and health & safety, parafoil technology was
chosen as the best approach to achieve reusability for the 1st and 2nd stage tanks. In order not to
completely abandon making the 3rd stage tanks reusable, it was determined that further
investigation into the idea of constructing space habitats on the Moon via Ion-Jet space tug
technology was necessary. The construction of the space tug would also allow the engines to be
transported to the Moon via the tug where resources such as fuel could be mined for use of the
engines; beneficial for future interplanetary missions.
It has been advised that developing in-house space operations would be more beneficial in the
long term. The operations site could be used a future commercial airport once LH2 became a
commonly used fuel by civil aircraft. This would bring in additional revenue to the investor
instead of renting say NASA's Kennedy Space Centre as a launch site. Another was to minimise
Author: Group
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cost per launch was to design an operation system that allowed much quicker turnover time. It
was suggested that Space X’s Grasshopper programme if successful would offer unsurpassed
reusability therefore in terms of vehicle retrieval, the design should either be done to incorporate
Space X’s technology or better since they are a future competitor.
Author: Group
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7.
FINALISED DESIGN PHASE
7.1.
7.1.1.
SECOND ITERATION OF MASS ESTIMATION
ENGINE MASS PREDICTION
In the fuel mass estimation, the engine was included as a part of the structural mass. This would
be true of a conventional rocket, where engines are dropped along with the rest of the structure at
each staging event.
Figure 64 Cross sectional view of a typical aerospike engine layout
However, in the previous analysis, it was identified that the pressure compensating engine should
be used throughout the flight and returned as a single unit, thus improving operational efficiency.
As such, the engine should be considered more as a part of the payload than the structure.
To size the engine, the maximum thrust it would need to produce was calculated. The maximum
thrust is the take-off thrust, calculated from the prior mass-estimation. The acceleration (including
against gravity) can be estimated. For the predicted take-off mass of 516T, a thrust of 8.10MN is
required to accelerate at a total of 0.7G (or 1.7G including the effect of gravity).
The average T/W is calculated from a set of first stage engines as 78.6. This is very conservative
compared with the high thrust to weight ratios claimed by modern engine manufacturers, such as
the T/W of 160 for the Merlin-1D engine (Space X 2012), but reflects the fact that the aerospike
design is non-conventional. The use of a conservative factor reduces the risk, should the engine
prove to be heavier than modern engines. It also allows more flexibility in safety factors applied
to the design for the enhancement of reusability.
Author: James Roper
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This yields an estimate of 10T for the engine mass. Because this ten tonnes must be added to the
payload mass, an additional 5T is added due to the increased fuel mass that must be carried,
bringing the engine mass to 15T.
7.1.2.
FINAL RE-SIZING
The vehicle is now re-sized based on the engine mass and improved predictions of the tank
structural mass fraction, which has been estimated as 6% (50% greater than the External Tank of
the Space Shuttle, allowing for re-usability).
Table 16 Performance characteristic and data for the finalised vehicle sizing
Total Launch mass:
654T
Payload mass:
31.6T
Payload Fraction:
4.83%
Launch Thrust (0.7G):
Engine Mass prediction:
Stage Breakdown:
10.3MN
13.4T
1
2
3
Totals
Full Tanks Mass (tonnes)
386
158
64.8
608.8
Dry Tanks Mass (tonnes)
21.9
8.96
3.67
34.53
Fuel Mass (tonnes)
40.5
16.6
6.80
63.9
Oxidiser Mass (tonnes)
324
133
54.4
511.4
Fuel Volume (m3)
597
245
100
942
Oxidiser Volume (m3)
284
116
47.7
447.7
The reduced payload fraction is not unexpected. This is a penalty associated with designing-in
reusability and applying safety factors to unknown structures. If a more optimistic figure were
used for the engine T/W, the payload fraction could be increased. For example, an engine T/W of
130 gives and engine mass of 7T and a total launch mass of 538T. This would yield a payload
fraction of 5.58%, for a 30T payload.
It should not be diminished however that the payload fraction, even with increased factors of
safety, is significantly higher than that of competitors and existing launchers. The Flacon Heavy is
predicted to launch 53T to LEO, and has a lift off mass of 1400T (SpaceX 2012), giving it a
payload fraction of only 3.78%, which is 21% less than the proposed launch vehicle.
The engine mass based on the new rocket mass is lower than the safety-factor of 5T applied to the
previous prediction. A further round of iteration using a smaller engine mass is one possibility.
However, it seems more prudent to allow the 1.6T of additional payload to be used elsewhere, for
example within the payload deployment structures, orbital manoeuvring systems and so on. Any
Author: James Roper
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further modification to the engine mass, as a result of detailed design and development, could also
be used to boost the payload capacity.
Based on the fuel volumes, tank sizes were calculated. As previously mentioned, it may be
desirable to give all tanks the same diameter, in order to allow parts of them to be interchanged
and to improve overall modularity. On this basis, a preliminary scaled sketch of the resultant
launch vehicle is given below, with 3m diameter tanks.
This layout contains a first stage of four tanks (31m length), a second stage of 2 tanks (25m
length) and a final central stage with a single tank (21m length). The engine remains with the
vehicle throughout the launch. The payload bay is 18m by 4m diameter. Height is approximately
Figure 65 Preliminary layout drawing for a parallel staged, common tank-diameter launch vehicle
64m.
7.1.3.
ENGINE DESIGN FOR RE-USABILITY.
Heat transfer rates within the engine have a strong influence on the design. The heat transfer is
highest at the throat and in the combustion chamber. The chamber temperature for an LH2/Lox
engine is lower than all the majority of fuel combinations, at 3132°𝐾. Hydrogen also has
excellent thermal conductivity (including at supercritical conditions). Because hydrogen
combustion does not result in any soot, it also gives lower adiabatic wall temperatures (Huzel and
Huang 1992).
As a result, using hydrogen for regenerative cooling is an almost inevitable choice. The only
drawback of this type of cooling is a limitation to the throttling ability of the engine. However,
this is circumvented by the use of multiple chambers, which can be shut down independently for
high throttle control. The benefit is that materials with high conductivity can be used, such as
copper used in the Space Shuttle main engine. It is noted that NASA studies have shown coppercoatings to be effective in preventing hydrogen embrittlement (Harris 2010).
Author: James Roper
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Despite this, high thermal loading of the interior wall will exceed the yield point and initiate
cracks (Sutton and Biblarz 2001). As a result, the internal wall is made thicker than necessary. A
destructive testing and safe-life program would be required to determine how long this would
survive for. If crack growth is sufficiently small, it could be repaired by machining away the
chamber surface and rebuilding it with a plasma coating.
The second role of the thrust chamber is to contain the chamber pressure. This is performed by the
outer layer of the thrust chamber, with the inner layer acting as a heat-shield only. Suitable
materials may include steel, nickel super-alloys, or carbon-carbon composites (which are the
lightest). These materials can hold their structural strength up to high temperatures – for example,
carbon-carbon composites are still structurally sound at 3700°𝐾 (Sutton and Biblarz 2001).
The nozzle throat typically incurs significantly heavier heat loads, and degrades faster than the
rest of the engine. To combat this, ablative, replaceable carbon-carbon composite inserts can be
used. Such inserts will burn in an oxygen-rich environment, so fuel-film or transpiration cooling
is employed over the surface. This reduces both the ablation and the oxidation, extending the life.
A typical estimate of 1-6% throat diameter increase is suggested. For the film-protected insert,
this could be 3%.
The increase in diameter would reduce the engine thrust by the same 3%. A drop in the specific
impulse would also be noted, of 0.7% (Sutton and Biblarz 2001). Calculations have been made to
show that the engine could lose this level of performance on each of ten flights before it became
unacceptable. To combat the drop in thrust and specific impulse, the payload would have to be
reduced by 3% per flight. An overhaul of the engine, including replacement of the insert, would
then allow for another ten flights. It is believed that this could be repeated four times, giving a
total of fifty flights.
The nozzle similarly experiences high heat flux, though not as high. The flux is not constant, and
is highest at the shock-wall interfaces, which move depending on the pressure-compensation
conditions. The XRS-2200 included a regeneratively cooled ramp section (Sutton and Biblarz
2001). As the conditions here are more benign, the liquid oxygen could be used for this task.
The turbo-pumps are given the most robust-possible design, to prevent creep and fracture. Axial
designs typically suffer more from creep than centripetal, as the direction of loading for an axial
turbine is also in the direction of the clearance to the wall. In civil and military turbo-jets, active
cooling of turbines is employed. In this case, fuel or oxidiser could be used to accomplish the
cooling. It is known that many hours of flight are possible in turbo machinery, so design of the
turbo-pump should follow the same design ethos and is expected to last for many flights.
Author: James Roper
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The electrical control unit of the engine would be placed in a low-temperature environment.
Having reached space, it would be subject to the space environment. The simplest way to keep the
control electronics safe is to encase them in a hermetically sealed and shielded vessel. This would
ensure that with minor testing, the control unit could be re-deployed. Such testing would be
automated and carried out by a computer diagnostic program.
This covers the sources of major expense within the engine. Other low cost items include
injectors, electrical looms, fuel and oxidiser pipes/hoses and valves. In addition, the sub-structure
of the engine is a low cost, but also low-stress item that would typically be easily recovered with
minimal inspection.
The conclusion reached is that a 10-flight reusability between overhauls is reasonable, and
following this an overhaul would be performed to address cracking within the combustion
chamber walls and to replace the carbon-carbon composite insert. The engine design would be
suitably modular in order to allow easy disassembly and access, and mechanical fasteners used
wherever possible.
7.2.
FINANCIAL ANALYSIS FOR THE PROPULSION SYSTEM
Development of the engine for reusability will entail a single year design study, followed by four
years of ground testing and two years of flight testing, making it feasible in the stated time period.
The first year of this study will cost approximately $17m, and each subsequent year (six in total)
will cost $85m (Parmalee 2002).
The setup of a factory for overhauling such engines could cost in the region of $400m, based on
the cost of a plant for engine manufacture planned by Toyota (RABELLO, Maria L & Mukai,
Anna, 2012) and scaled down for the lower production volume expected, but respecting the high
cost of advanced machinery and equipment. This facility would also perform the engine overhaul.
The maintenance, operations and technical support activity associated with the engines is
estimated to be $5.7m, per engine-launch, from NASA’s operational expenses on the Space Shuttle main engine (McCleskey 2005). The workforce required to build and overhaul the
engines, as well as for support activities, will consist of 54 people and have a combined salary of
$7.1m per year.
Table 17 Division of labour for engine production and overhaul
Author: James Roper
Work area
Number employed
Strip and Build Team
10
Movements Team
2
Accumulation and Warehousing
2
Engineering Team
3
Repair Team
10
Inspectors
5
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NDT Inspectors
2
Production Team
3
Materials Team
5
New Build (Replacement parts)
10
Technical Records Team
2
The cost of the engine itself is estimated to the $15m new, based on the oft-quoted figure that jet
engines are ‘worth their weight in silver’ (DERBY, 2009) (SILVERPRICE, 2012) and the initial
mass estimate of 15T for the engine.
It would be realistic to expect that the engine could make fifty flights during its lifetime, with
overhaul every ten flights (four overhauls). For example, to operate a flight each month, each year
for thirty years, this leads to a requirement for eight engines to be constructed.
The cost comparison is illustrated on the chart below. For reusable engine, the cost per flight is
calculated as follows: The initial outlay cost is divided by the number of flights. Initial outlay
costs are the development and set-up, plus the number of engines required multiplied by the
engine unit cost. The operational costs per flight are then added.
For the non-reusable data, the engine cost is added to the development and set-up costs per flight,
as well as the operational costs.
50
Reusable
Cost per flight, $m's
45
Non-Reusable
40
35
30
25
20
15
10
5
0
0
100
200
300
400
500
Total number of flights during program
600
Figure 66 Comparison of cost-per-flight for reusable and non-reusable engines over varying launch numbers
A cost per flight of approximately $9.02m is estimated for the engine, if used for 390 flights. That
is, 13 flights per year for 30 years, with eight engines. If the engines were not reused, this would
be $17.42m per flight, thus there is a cost saving associated with reusability of 48%.
Author: James Roper
| Page 99 of 262
7.3.
LAUNCH TRAJECTORY
Having defined the stage masses, initial estimates of the mass flow-rate required to achieve the
thrust necessary were made. These were then used as preliminary guesses in a Matlab program,
which calculated the flight angle and accounted for accelerations due to drag and gravity
numerically, over the calculated flight-time of each stage. See Appendix 12.5.
The results of this indicated that orbit would be achieved, with a 330km altitude, for the payload
of 30 tonnes. The vehicle would have accelerations of between 1.5G (at stage ignition) and 4G (at
stage cut-out). In total, the launch from ground to orbit takes approximately 6 minutes.
The main outcomes of this simulation were the cut-off speeds and altitudes of the stages. This
data was necessary to select appropriate re-entry and recovery methods, as well as appropriate
thermal protection. A summary of the relevant data is provided in Table 18.
Table 18 Vehicle trajectory for each flight staging event.
Stage
1
2
3
2.26
2.85
7.69
Cut-off Altitude (km)
54
203
328
Cut-off Range (km)
81
441
1431
Burn time(s)
100
120
130
Angle at burn out
47°
18°
7°
Cut-off velocity (km/s)
The other important outcome from this simulation is to confirm that the estimates made for DeltaV resulting from drag and gravity were indeed sufficient. As such, it can be stated with reasonable
certainty that the proposed launch vehicle is feasible.
From this exercise, the recovery and reusability of the spent tank stages will be determined.
7.4.
FUEL TANK DESIGN
Material selected for the tank design is an important aspect in terms of weight-saving since tanks
contribute a large percentage of the overall vehicle structural weight and compatibility with the
type of propellants chosen in terms of fuel storage.
During launch, tanks experience various loading so the tanks were designed to withstand the
dynamic loadings on the vehicle in the axial and lateral directions. This tank will be designed to
be stabilised against buckling by keeping the tank walls under tension loads always at a specified
pressure level to be maintained during storage and handling. Because huge tank structures are
sensitive to external buckling loads, two ways used to stabilise large booster stage systems are
(Huzel and Huang 1992);
Author: James Roper / Charles Ofosu
| Page 100 of 262
i.
Pressure stabilisation

This is when the tank pressure is maintained above the specific minimum by use
of elaborate controls. Usually thin-walled monocoque which requires special
handling procedures.
ii.
Self-supporting

In this system, tanks wall are reinforced with skin stringers or use of waffle grid
patterns.
The self-supporting method was chosen to increase buckling strength with the tank designed as a
semi-monocoque structure.
Whilst considering which fuel type to go with, tank design was a factor also considered as well as
reliability, environmental factors, etc. Various tank design problems caused by use of cryogenic
propellants include thermal gradients, need for insulation and need for constructional materials
which are capable of remaining ductile at really low temperatures, (WERTZ, James R. and
Larson, Wiley J., 2005). However, after detailed analysis and consideration, hydrogen was chosen
was mentioned above in section 6.1.2.
7.4.1.
MATERIAL CHOICE
In order to design the tank structure, there was a need to compare different materials and their
suitability to the fuel choice as well as trade studies to compare weight implications and risks.
Since the aim was to design a heavy lift launch vehicle, the structural masses were driven by
mainly strength requirements and the material chosen must have the highest ultimate strength-todensity ratio.
Typical materials used include

Aluminium alloys 2000 & 6000 series

Steel alloy AISI 300 series

Titanium alloys

Filament-wound graphite/resin composites.
For operating temperatures up to 177°C, aluminium alloys are compatible with cryogenic
propellants. Liquid hydrogen operates at temperatures as low as 20 K. Titanium alloys and the
austenitic and semi-stainless steels are known to possess good mechanical properties at cryogenic
temperatures. Toughened carbon fibre reinforced plastics (CFRP) are being investigated for use as
the main structural materials of cryogenic propellants because their low densities would result in
drastic weight reduction required for efficient reusable space transportation vehicles, (Morino
2001). Results from (Morino et al, 2001) show that these materials have good cryogenic
properties.
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Aluminium-Lithium alloy 2195 was chosen as the tank’s material based on its high strength to
density ratio, good weldability and it exhibits better cryogenic ductility and much greater strength
than the conventional alloy for cryogenic tanks, Al-Li 2219, (HARTUNIAN, R. A, 1995). It also
offered 4% weight saving when used to manufacture the Space Shuttle’s External Tank compared to conventional alloys. More conclusive results in using toughened CFRP for cryogenic tanks
would be required to make a change to it but at this stage, it was deemed too risky.
7.4.2.
THERMAL PROTECTION SYSTEM
The choice of fuel for the launch vehicle was liquid hydrogen with liquid oxygen as the oxidiser
therefore tank insulation was necessary. Insulation is mandatory for cryogenic propellants to
prevent ambient-air liquefaction which leads o high heat transfer rates (WERTZ, James R. and
Larson, Wiley J., 2005). In order to prevent high boil-off rates, cryogenic insulation is required on
both sides of the tank wall.
(Morino et al, 2001) evaluated two types of plastic foam used at cryogenic temperatures and the
results has illustrated by Figure 67 shows that Airex foam has better elongation properties
compared to Rohacell foam particularly at low temperatures. Therefore, at this stage of the design
process it is proposed that Airex foam should be used as the main insulation material for the
propellant tank of the launch vehicle. The insulation must reduce the heat influx into a hydrogen
system significantly and reduce the impact of hydrogen embrittlement.
Figure 67 Comparison of two typical cryogenic insulation materials. (Courtesy of Morino et al.)
Figure 68 below was obtained from (Johnson et al, 2005) and illustrates that an Al-Li 2195 alloy
with Airex form insulation results in no cracks and no disbands when the tank was subjected to
combined thermal and mechanical tension loading, (JOHNSON, Theodore F. et al., 2005).
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Figure 68 Experimental results for tanks subjected to both thermal and mechanical tension loading. (Johnson et
al, 2005)
7.4.3.
STRUCTURAL DESIGN AND ANALYSIS
Once the vehicle was re-sized based on a design review and optimisation, the final estimate of the
tank masses were to be calculated. The process used to estimate the mass of the fuel tanks is
outlined below and was obtained from (WERTZ, James R. and Larson, Wiley J., 2005). It should
be noted that the calculations used assume a thin-walled monocoque structure without skin
stringers.
Table 19 Process for estimating the mass of the tank.
Step
Description
1
A structural approach was selected detailing the type and shape of the structure,
arrangement and load paths.
2
An initial rough estimate of the mass distribution for the tank structure and all equipment
was taken.
3
Using the information from the above steps and the axial and lateral frequencies
experienced during launch acceleration, estimates for the mass were calculated. Iteration
took place until the desired structural design was obtained.
4
Load factors (axial, lateral and bending) that would be experienced during launch were
combined and applied to calculate the maximum design load.
5
The structural capability was computed and compared with the applied loads to
determine a margin of safety. The design was then iterated and optimised to obtain the
necessary margin of safety.
| Page 103 of 262
The fuel tanks would be semi-monocoque cylindrical shapes with spherical end caps. They would
consist of a first stage of four tanks (31m length), a second stage of 2 tanks (25m length) and a
final central stage with a single tank (21m length) with a constant diameter of 3m. The uniform
mass distribution of a single cylinder was assumed to be 1000 kg based on initial tank sizing.
As described in the previous section, Material Choice, the material chosen was AluminiumLithium alloy 2195 which has the following properties;; Young’s Modulus (E) – 71 × 109 N/m2,
density (ρ) – 2865 kg/m3 and ultimate tensile strength (σtu) – 609×109 N/m2.
7.4.4.
RIGIDITY SIZING
Since the tanks would have to withstand vibrations during the launch phase, sizing for rigidity to
meet the natural frequency requirements were performed. The axial frequency and lateral
frequency of 25 Hz and 10Hz respectively were chosen based on the Titan II fundamental
frequencies (BARTER, Neville J. and Thompson, Tina D., 1992).
The natural frequency equation
𝑓=
1 𝑘
2𝜋 𝑚
Where m is mass, k is the stiffness (spring constant).
This equation becomes the one below when applied to axial beam becomes according to
(WERTZ, James R. and Larson, Wiley J., 2005).
𝑓 = 0.25
𝐴𝐸
𝑚 𝑙
Where A is the cross-sectional area of the cylinder, 𝑚 is the uniformly distributed mass, l is the
length of the cylinder and E is the Young’s Modulus.
𝐴 = 2𝜋𝑅𝑡
Again when the natural frequency equation is applied to a lateral beam, it becomes;
𝑓 = 0.56
𝐸𝐼
𝑚 𝑙
Where I is the moment of inertia of the cylinder’s cross-section;
𝐼 = 𝜋𝑟 𝑡
Solving for axial rigidity gives the following required thicknesses for the various lengths,
| Page 104 of 262
Table 20 Axial rigidity sizing for all 3 tank dimensions.
Length, l (m)
Area, A (m2)
Required thickness, t (m)
31
4.37×10-3
0.463×10-3
25
3.52×10-3
0.374×10-3
21
2.96×10-3
0.314×10-3
Solving for lateral rigidity gives,
Table 21 Lateral rigidity sizing for all 3 tank dimensions.
Length, l (m)
Moment of inertia, I (m4)
Required thickness, t (m)
31
13.4×10-2
12.6×10-3
25
7.02×10-2
6.62×10-3
21
4.16×10-2
3.92×10-3
Thus the lateral (bending) mode is much more critical and their thicknesses were used in further
calculations. For instance, the 31 m long tank’s cross-sectional area is
𝐴 = 2𝜋 × 1.5 × 12.6 × 10
7.4.5.
= 0.119 𝑚
APPLIED AND EQUIVALENT AXIAL LOADS
The load factors used here were obtained from the Delta rocket family series (BARTER, Neville
J. and Thompson, Tina D., 1992), during the launch phase the tank must be designed to survive
the sum of steady-state and dynamic accelerations in the axial and lateral directions including a
bending moment from the centre of mass location. The tank initial weight estimate was multiplied
by the load factors to obtain the limit load. The limit load is the maximum load expected in each
critical period, with an allowance for statistical variation, (WERTZ, James R. and Larson, Wiley
J., 2005).
Table 22 Applied launch acceleration loads to the tank structure.
Loading
Weight (N)
Axial
Distance (m)
Load factor
Limit load
9807
3.4
33344 (N)
Lateral
9807
2
19614 (N)
Bending moment
9807
2
304017 (Nm)
15.5
At this stage, the equivalent axial load, Peq was calculated. This evaluates the combined axial,
lateral and bending loads on a thin-walled cylinder and is based on the fact the bending stress will
be greatest at the two points farthest from the neutral axis. It is an axial load on the tank that
would result when uniform stress becomes equal to a peak stress created by the sum of an axial
| Page 105 of 262
and bending moment. According to (WERTZ, James R. and Larson, Wiley J., 2005), the peak
stress would occur along a cylinder’s circumference leading to the formula below,
𝑃 =𝑃
+
2𝑀
𝑟
For the 31 m long tank,
𝑃 = 438700 𝑁
The ultimate/allowable load is a product of the limit load and ultimate factor of safety and under
this level of load, the structure must not collapse, rupture or undergo gross deformation. The
typical ultimate factor of safety used for spacecraft design is 1.25 (WERTZ, James R. and Larson,
Wiley J., 2005).
Therefore the ultimate load = 438700 × 1.25 = 548375 N.
7.4.6.
TENSILE STRENGTH SIZING
Using Al-Li 2195’s maximum allowable tensile stress of 609×109 N/m2, the thickness required for
the tank to survive the ultimate load was calculated;
σ = P/A where A= 2πrt, therefore
𝑡=
7.4.7.
548375
= 0.0955 𝑚𝑚
2𝜋 × 1.5 × 609 × 10
SIZING FOR STABILITY
The equation for a cylinder buckling stress
Since a thin walled cylinder is susceptible to crinkling and collapse by local buckling, it is
important to design against this. For a thin cylinder of wall thickness, t, the buckling critical
stress, σcr is given by elasticity theory as (CRUISE, A.M. et al., 1998):
σcr = Et/r × [ 3 (1 - n2) ]-1/2 = 0.6 E t / r for n = 0.3 (which is valid for most materials).
In practice however, this overestimates test results by a factor up to 3x, thus the more
conservative formula used was;
𝜎
= 0.2
𝐸𝑡
𝑟
This was equal to 119 MPa for the 31 m long tank leading to a critical buckling load of 14.16 MN
based on the formula below.
𝑃 = 𝐴𝜎
| Page 106 of 262
Clearly, the applied ultimate load Peq is much less than the critical buckling load and this
demonstrates the structural integrity of the design. A method of showing structural integrity of a
component is margin of safety (MS) which is defined as;
𝑀𝑆 =
𝐴𝑙𝑙𝑜𝑤𝑎𝑏𝑙𝑒 𝐿𝑜𝑎𝑑
−1
𝑈𝑙𝑡𝑖𝑚𝑎𝑡𝑒 𝐷𝑒𝑠𝑖𝑔𝑛 𝐿𝑜𝑎𝑑
The margin of safety must be greater than or equal to zero to be accepted. The initial iteration
generated a margin of safety shown below;
𝑀𝑆 =
14.16 × 10
− 1 = 25
548 × 10
This MS represents 2500% thus further iterations were performed with the aim to minimise this
drastically in order to meet the tank mass requirements. The table below shows a summary of
sizing the cylinder for stability.
Table 23 Iteration to find the optimum thickness to meet stability requirements.
Thickness (m)
σcr (MN/m2)
Area (m2)
Pcr (kN)
MS
12.6×10-3
119
0.119
14.16×103
25
2.5×10-3
23.7
0.0236
558
0.017
2.6×10-3
24.6
0.0245
603
0.100
2.7×10-3
25.6
0.0254
650
0.186
2.0×10-3
18.9
0.0188
357
-0.349
The thickness chosen was 2.6×10-3 m as it conformed to the tank mass requirements as well as
offering a more than adequate margin of safety, 10% which may lead to higher reliability.
Now, the hoop stress of the cylindrical tank was calculated using a critical pressure of 2.34 bar
(WERTZ, James R. and Larson, Wiley J., 2005).
𝜎 =
=
.
×
. ×
× .
= 140 𝑀𝑃𝑎
The Al-Li alloy’s maximum allowable stress is more than four times higher than the hoop stress which proves its suitability.
7.4.8.
MASS CALCULATION
Lastly the mass of the cylindrical part of the tank is the product of the density, ρ and the volume, 2πrtl.
| Page 107 of 262
𝑚 = 𝜌2𝜋𝑟𝑡𝑙 = 2685 × 2𝜋 × 1.5 × 0.0026 × 31 = 2040 𝑘𝑔
The spherical end masses were calculated using the same equation stated in the chapter Initial
Tank Sizing. Exact calculations were performed for all the tank lengths shown in chapter 12.6 of
the appendix however, a summary is provided in the table below.
Table 24 Total tank mass calculations
Length
Quantity
Cylinder Mass
Spherical end
(kg)
mass (kg)
Total
(kg)
31
4
2040
99
8556
25
2
1645
99
3488
21
1
1382
99
1481
Thus the total mass for all the tanks is 8556+3488+1481 = 13,525 kg.
From (WERTZ, James R. and Larson, Wiley J., 2005), 20-30% of overall tank mass must be
added to account for propellant management devices, pumps and mounting hardware resulting in
a total mass of approximately 17,583 kg.
7.4.9.
COMMON BULKHEAD
The propellant tank is proposed to have a common bulkhead to reduce its weight. Usually, the
fuel tank would be separated from the oxidiser tank to avoid excessive heat transfer between
them, however to reduce the total structural weight of the tank considerably, a common wall
(bulkhead) is introduced between the fuel and oxidiser tanks (Logsdon 1998). Extra spray-foam
insulation is then used between the two fuels to decrease the heat transfer between them to an
acceptable degree; this technique resulted in weight saving of 3600 kg of the Saturn V Mon rocket
(Logsdon 1998).
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7.5.
LAYOUT DESIGN
Aerodynamic Body:
The Launch Vehicle will need an
aerodynamic body to reduce its drag
coefficient. It is recommended that a
Haack series nosecone is used.
Fairing:
The Launch Vehicle fairing will
be made of 4 quarter segments.
This is to improve separation
reliability.
Fuel Lines:
Contrary to the image shown the fuel
lines, marked in orange, should be
Heat Shielding:
facing in towards the core structure.
The
This will add protection and simplify
experience strong heating due to
the pumping system.
friction during launch. Each nose
Launch
Vehicle
will
cone will include heat shielding
as shown in black.
Separation System:
The separation of the fuel tanks will
be a major design challenge. Stage
separation is one of the primary
causes for launch vehicle failure;
this system should be failsafe and
have redundancies.
The chosen design has one
noticeable limitation which is the
requirement for a heavy support
structure between the engine and
the 3rd stage. Further design
analysis could be done to reduce
this sizing, perhaps by the usage
of alternative tank diameters.
This section could also be used to
house secondary payloads such
as cube satellites.
| Page 109 of 262
7.6.
IN-ORBIT TRAJECTORIES AND PROPULSION
In the early days of space exploration just being able to put a satellite into any orbit was
considered a great success. As space flight has moved on from exploration to being commercially
utilised it has become a requirement to place satellites into very specific orbital trajectories.
7.6.1.
ORBIT DEFINITIONS
The aim of positioning satellites is not to escape the Earth’s gravitational pull but rather to balance it so it adds complexity to the launch trajectory as it conflicts with the aim of reaching to
deep space.
POLAR ORBITS
This type of orbit is typically in LEO with an inclination of 90°. Much of the Earth’s surface is observed in these orbits making them ideal for mapping applications and Earth observation. Some
examples of spacecraft that operate in these environments are the USA Air Force surveillance
satellites of the DMSP series, or the series of French Earth-resources spacecraft SPOT (STERN,
David P and Peredo, Mauricio, 2006). However it should be said that trajectories for human
spaceflight should avoid passing through these orbits as they would receive increased radiation
doses when passing through the aurora (STERN, David P and Peredo, Mauricio, 2006). It also
should be noted that due to safety due to safety constraints, polar orbits cannot be achieved from
KSC, and such requirements have been met by expendable launches from Vandenberg AFB
(VAFB) near Lompoc, California. (GRIFFIN, Michael Douglas and French, James R, 2004)
SUN SYNCHRONOUS ORBITS
Also typically in LEO, they have an inclination of approximately 98°.It is a polar orbit which
provides consistent sun lighting condition along the ground track (GT) of the orbit. (BOAIN,
Ronald J., 2004) For some applications it may be important to have the maximum amount of
sunlit exposure. Calculations may be done to determine the launch windows required to achieve
this. There are two opportunities a day where for the right ascension angle of the sun angle α
which requires that- 𝛼 − 𝛺 = ± . The season or possible launch time of the orbit is influenced
by both the inclination of the orbit (related to the orbital altitude for Sun-synchronous orbits) and
the fact that the apparent solar motion is from latitude 23.4°N during summer solstice to 23.4°S
during winter solstice (Fortescue and Stark 1995)
GEOSYNCHRONOUS ORBITS
These are placed in GEO and have an inclination of close to 0°. Geostationary orbits are ideal for
weather satellites and communications satellites as they are orbits in which the satellite remains in
the same location above the Earth allowing for optimal ground station locations. It is typically at
| Page 110 of 262
an altitude 35764 km and the strategy for launching satellites requiring these types of orbits is
usually to place them in a roughly circular type of orbit at about a few hundred kilometres at an
inclination of 28° and then an orbital transfer propulsion unit will fire to bring about the required
orbital acquisition (Fortescue and Stark 1995).
MOLNIYA ORBITS
At an altitude of 39000km, these are highly elliptical orbits and inclined 63.4° from the equator.
They are used as an alternative to Geosynchronous Orbits for high latitude communications,
popularly adopted in Russia and additionally with interests elsewhere. The perigee of
approximately 600km is chosen such that it is avoids excess drag (KIDDER, Stanley Q and
Vonder Haar, Thomas H., 1990)and its altitude is well above the altitude of comparable
geostationary satellites with a difference of 4000 km between the highest.
From the space shuttle system if the angle of inclination used for launch needs to be higher than
28.5° the payload would have to be lighter to achieve the same altitude. The proposed system was
analysed using a polar orbit consideration or worst case scenario so this criterion may change
slightly but for considerations Figure 69 and Figure 70 shows how the shuttle’s payload capacity
would change based on the altitude needed at 28.5° inclination.
Figure 69 Cargo weight vs Circular orbital altitude for KSC launch. (GRIFFIN, Michael Douglas and French, James R, 2004)
Figure 70 Cargo weight vs circular orbital altitude-VAFB Launch (GRIFFIN, Michael Douglas and French, James R, 2004)
| Page 111 of 262
PARKING ORBIT
Depending upon the final payload destination a parking orbit may be used. These destinations are
typically MEO, GEO and beyond. This is because satellites destined for these higher orbits often
require a precise position within the specified orbit; if this position was to be achieved by
launching directly from Earth then the launch window would be very small, in the order of
seconds each day. To mitigate this issue a parking orbit is used where the tight launch window is
replaced with a transfer window, which could occur up to 16 times a day. Another benefit of a
parking orbit is that it allows time for system tests to check the payload was not damaged during
launch. For LEO destination no parking orbit is used, instead the payload is placed directly into
the correct orbital altitude.
7.6.2.
ORBITAL ELEMENTS
To achieve a desired orbit it is necessary to characterise it first. It is possible to describe simple
circular and elliptical orbits using six Keplerian orbital elements.
In real world applications there are perturbations that cause slight changes to an orbit. An example
of this would be atmospheric drag, solar radiation pressure, Earths equatorial bulge and the
higher-order potential harmonics of the Earth (Logsdon 1997).Each of these perturbations
requires an extra orbital element to model it. An example of this would be the GPS satellites
which use 16 orbital elements.
7.6.3.
ORBITAL MECHANICS
To reduce the ΔV requirement to gain orbit it is possible to get assistance from the Earth’s rotational velocity. The maximum rotational velocity of the Earth occurs at the equator. The
maximum assistance occurs when the launch vehicle has an azimuth of 0°, thus going due east.
The ΔV assistance will equal the Earth’s rotational velocity at this point thus:
∆𝑉
=
2×𝜋×𝑅
24 × 60 × 60
Where 𝑅 denotes the radius of the Earth. This gives us a ΔV assistance of 463 m/s. To calculate the assistance of different launch latitudes:
∆𝑉
= 𝑐𝑜𝑠 (𝑙𝑎𝑡) × 463
The rest of this section will identify the manoeuvres and ΔV required to achieve a specific orbit
from a variety of parking orbits. One of the fundamental equations used in space flight is Isaac
Newton’s vis viva (living force) equation expressed below:
| Page 112 of 262
𝑉=
𝜇×
2 1
−
𝑟 𝑎
Where 𝑉 denotes the velocity of the satellite, 𝜇 represents the gravitational parameter of Earth, 𝑟
is the radial distance between the centre of the Earth and the centre of mass of the satellite and 𝑎
is the semi-major axis of its elliptical orbit.
For a circular orbit, 𝑎 is equal to 𝑟 thus giving:
𝑉
𝜇
𝑟
=
Using the vis viva equation it is possible to calculate the velocity required to escape the
gravitational field of Earth, this is often known as the C3 value. The equation and is derivation for
this is given below:
𝑉
=
𝜇×
𝑉
=
𝑉
2 1
−
𝑟 ∞
2𝜇
𝑟
= √2𝑉
Thus to escape the Earth a 41.4% increase of a circular orbit’s velocity is required.
The kinetic energy and potential energy of a circular orbit remains constant. In an elliptical orbit
these are not constant, which causes the elliptical motion. Only the only common orbit that
requires an elliptical motion is the Molniya orbit, for the sake of simplicity only circular orbits
will be analysed in this section.
7.7.
7.7.1.
POWERED MANOEUVRES
HOHMANN TRANSFER
The Hohmann transfer involves two engine burns to either raise or lower a satellite or vehicles
altitude relative to the Earth. A single burn would place the vehicle into an elliptical orbit with the
kinetic and potential energy constantly shifting between one another, the second burn is required
to balance this shifting process. Assuming the increase in velocity is instantaneous it is possible to
calculate the ΔV required for the initial burn and second burn:
∆𝑉
𝜇
𝑟
2×𝑟
−1
𝑟 +𝑟
| Page 113 of 262
∆𝑉
𝜇
2×𝑟
1−
𝑟
𝑟 +𝑟
Where 𝑟 is the initial distance from the centre of the Earth and 𝑟 is the desired distance. Thus the
total ΔV required to change altitude using the Hohmann transfer is:
∆𝑉 = ∆𝑉
+ ∆𝑉
In reality the increase in velocity is not instantaneous, for low thrust engines such as ion thrusters
it may require days or weeks to achieve the necessary total ∆𝑉 for the transfer. The effect of this
is that the ∆𝑉 will be increased due to gravity losses. This must be taken into consideration when
choosing the type of thrusters to be used in the manned spaceship and any unmanned vehicles that
may be built.
The ∆𝑉 requirement to transfer a satellite from a 330km circular parking orbit with 0° inclination
to a higher altitude circular orbit with 0° inclination has been plotted in Figure 71.
Figure 71: ∆V Requirement from the parking orbit where the altitude of RapidEye is 630km, MEO is 23,222km, GEO is 35,786km and the Moon is 384,400km.
7.7.2.
PLANE CHANGE MANOEUVRES
The launch vehicles often deliver multiple satellites into orbit but not all of these satellites will
have the same desired orbital inclination. To change a satellite’s inclination requires a plane
| Page 114 of 262
change manoeuvre that can either be carried out by the launch vehicle itself or by the satellite
itself. The launch vehicle will sometimes perform the plane change manoeuvre for the primary
satellite being carried but it can also be achieved by the satellites own propulsion system.
The ∆𝑉 requirement for a plane change manoeuvre can be calculated using the following
equation:
∆𝑉 = 2 × 𝑉
× 𝑠𝑖𝑛
∆𝜃
2
For commonly used altitudes the ∆𝑉 requirement for a plane change manoeuvre of angle 𝜃 has
been plotted in Figure 72.
Figure 72: ∆V Requirement for a plane change manoeuvre
7.7.3.
COMBINING HOHMANN TRANSFER WITH PLANE CHANGE
It has been found that the optimal transfer manoeuvre between different altitudes and inclinations
is to combine the 1st burn of the Hohmann transfer with a small part of the plane change
manoeuvre and the 2nd burn to correct for the remaining plane change required. The ∆𝑉 saving
when transferring from a circular orbit at 28.5° inclination to a geostationary orbit at 0°
inclination is approximately 70 m/s, about 0.5% of the original ∆𝑉 requirement. For simplicity
this will be ignored in further analysis and all plane change manoeuvres will occur at the
operational altitude.
7.8.
GTO DELIVERY
The currently proposed launch system is capable of delivering a payload of 30,000 tons to a LEO
of 330km. For additional manoeuvres, such as a Hohmann transfer or station keeping, the payload
will have to carry its own propulsion system and fuel.
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In traditional launch philosophy a payload going to LEO will be deployed in the launch vehicles
parking orbit and make its own way to its desired orbit. For a payload going to MEO or GEO the
launch vehicle would provide the additional ΔV required for the 1st part of a Hohmann transfer
with the payload providing the remaining ΔV with its own rocket. This method would result in the launch vehicle engine being left in GTO and thus not being re-usable.
7.9.
SPACE TUG
To overcome this issue the use of a Space Tug is proposed. This could be used to initiate the
transfer of a payload to a desired higher orbit and inclination. The Space Tug would then be on a
free return elliptical trajectory using its own power to decelerate to a LEO parking orbit and
receive its next payload for delivery plus additional fuel. This fuel could either be carried by the
launch vehicle or retrieved from an in-orbit fuel depot. This leaves the launch vehicles engine free
to return to Earth after the initial launch.
The size and capability of the space tug will determine the largest payload mass possible to
transfer to a higher orbit such as GTO or to the Moon. The fuel mass of this transfer vehicle can
be calculated using the rocket equation discussed in Chapter 6.1.2.
The critical design choice of the Space Tug is the propulsion method to be used. A method with a
lower ISP such as a LOX and LH rocket can provide a much greater thrust than a method with a
higher ISP such as an Electrostatic Ion Jet. A higher thrust allows mitigation of losses due to
gravity but requires a substantially larger mass of fuel. Both of these methods will be compared
using an ISP of 475s and a vehicle dry mass of 4 Tonnes for the LH2/Lox system (DAVIS,
Richard, 1988) and an ISP of 2,000s and a vehicle dry mass of 1.5 Tonnes for the Ion Jet system
(R.Wertz and Wiley J. Larson 2008).
An example of the maximum deliverable payload to GTO has calculated assuming that:

The Space Tug is already placed in orbit so only additional fuel must be carried in the
launch vehicle.

The thrust provided by the LH2/Lox Space Tug is instantaneous thus losses due to
gravity can be neglected.

The thrust provided by the Ion Jet Space Tug is not instantaneous and will be required to
provide 200% of the ideal∆𝑉.(Logsdon 1997)

Only 1st burn of the Hohmann transfer and a 2nd deceleration burn will be required giving
a ∆𝑉 requirement of 3.09km/s for the 1st burn and the same again for the deceleration
burn.

There is no additional ∆𝑉 requirement for a plane change manoeuvre.
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The calculation will work backwards from the empty return of the Space Tug to LEO. This will
allow the total fuel required to be predicted using the rocket equation stated in Equation 1.
∆𝑉 = 𝑔 𝐼 𝑙𝑛
𝑊
𝑊
Table 25 states the specifications of taking the heaviest class of geostationary satellites,
approximately 6 tonnes to GTO.
Table 25: Comparison of Space Tug propulsion methods for delivery of a 6T payload to GTO
LH2/Lox Space Tug
Ion Jet Space Tug
Payload Mass
6,000 kg
6,000 kg
LEO Return Dry Mass
4,000 Kg
1,500 Kg
Pre GTO - LEO burn Mass
7,763 Kg
2,055 Kg
Post LEO - GTO burn Mass
13,763 Kg
8,055 Kg
Pre LEO - GTO burn Mass
26,711 Kg
11,037 Kg
Fuel Required
16,711 Kg
3,537 Kg
Delivery Time
Hours
Weeks/Months
Cost/Kg to GTO
USA$4,451
USA$1,839
To get an accurate comparison of the vehicle against identified competitors, in this case SpaceX,
the payload mass to different elliptical transfer orbits has been calculated. Figure 73 represents
this data.
Figure 73 also gives a comparison of what the Space Tugs capability would be if it was used in
conjunction with Space X’s Falcon 9, which is capable of delivering 13,150 Kg to LEO. A
comparison was made against the official Falcon 9 GTO delivery data given in Figure 74. This
comparison showed that with a Ion Jet based Space Tug it would be possible for the system to
deliver approximately 480% more cargo to GTO than Space X. This drops to approximately
280% increase if a LH2/Lox based Space Tug is used.
If the Ion Jet based Space Tug were used in conjunction with the current Falcon 9 system it could
improve Space X’s delivery performance by up to 200%. A LH2/Lox based Space Tug in
conjunction with the current Falcon 9 would decrease its GTO performance by 42%.
This gives an interesting trade-off. If a LH2/Lox based Space Tug is chosen to be developed in
parallel to the main project then it could provide a good heritage in LH2/Lox propulsion
technology resulting improved reliability and efficiency. It would also simplify the development
procedure as only one type of propulsion system would be developed.
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Alternatively, if an Ion Jet based Space Tug was pursued it could vastly improve the delivery
performance, at a sacrifice of delivery time. If developed early during the project lifetime it could
be used in conjunction with competing launch vehicles to provide a source of revenue before the
launch vehicle was operational.
Figure 73: Payload mass deliverable to GTO with Space Tug and its performance in conjunction with Falcon 9
Figure 74: Payload mass deliverable to GTO using Falcon 9
In conclusion it has been decided that it is best to pursue a both an Ion Jet and a LH2/Lox
powered Space Tug. The efficiency of Ion Jet propulsion is enticing and is a worthy asset for a
company that may be interested in expanding into further roles in space. The anticipated
development cost of a LH2/Lox Space Tug of this size is approximately $500M (DAVIS,
Richard, 1988) and should be pursued in conjunction with the primary project to ensure it is fully
developed and implemented in time for peak launch vehicle usage.
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7.10.
RE-ENTRY PHILOSOPHY
7.10.1.
ATMOSPHERIC DENSITY
The critical feature behind the amount of deceleration a vehicle undergoes, and the overall profile
of its descent, is the density of the atmosphere. Obviously, the density increases as the vehicle
descends. The density at sea level is not anywhere near the same as the density 100km up. Gravity
is obviously the reason for this, as the atmosphere has far less mass above it high up than it does
at sea level. This causes the air particles to be spread out at high altitudes, having the effect of
limiting the amount of friction that can be gained, and the number of particles that collide and
form plasma. On the one hand, limited friction at high altitudes means launch becomes easier as
the vehicle ascends. On the other, descent is harder because the vehicle, with little opportunity for
deceleration, gains velocity farther out in the atmosphere, and by the time it reaches a thicker
atmosphere, the velocity it has causes plasma to form readily in the air around it, and means the
vehicle experiences high friction.
For the purposes of this report, an appropriate model for atmospheric density had to be used. A
NASA atmospheric model was used, which modelled the atmosphere based on varying
temperature and pressure with altitude. It also uses three tiers, the first for atmospheric density up
to 11km, the second from 11km to 25km, and the third for altitudes above 25km. This makes for
more accurate calculations of deceleration, and also heating, in the MS Excel simulation.
The specific formulae NASA suggests are as follows:
Table 26 NASA Earth atmospheric model
h
< 11,000
11,000 < h < 25,000
h > 25,000
15.04 − 0.00649ℎ
−56.46
−131.21 + 0.00299ℎ
(Alt. in m)
T
(Temp. in ˚C)
P
101.29
(P es. in
×
KPa)
𝑇 + 273.1
288.08
.
2.488
22.65
× 𝑒(
.
.
)
×
𝑇 + 273.1
216.6
.
These boundaries have been put into the following equation:
Equation 3 NASA equation for atmospheric density
𝝆 𝑫𝒆𝒏𝒔. 𝒊𝒏 𝒌𝒈⁄𝒎𝟑 = 𝑃⁄ 0.2869 × (𝑇 + 273.1)
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In the context of the simulations created for this report, this model is implemented over every
time increment to give as accurate a representation of the atmosphere as the vehicle descends.
7.10.2.
PARACHUTES
The conventional, most tried-and-tested, method still employed today for deceleration is the
parachute, or in some cases parachutes. More recent advances in the field have endeavoured to
improve the deceleration of the vehicle during descent, but the guaranteed choice is the parachute.
The range of parachutes is wide however, and whilst parachutes are all designed to greatly
increase the air resistant properties of the falling object, the differing conditions between high
altitude descent at great velocities and low velocity descents means that a one-fits-all approach
cannot be taken. For a start, different deceleration requirements mean that the force applied to the
parachute material can vary greatly, and this in turn means the requirement for material strength
may vary. This in turn affects the parachute weight and shape.
Certain materials used resist air flow better than others. Certain materials fill out and stretch,
providing the best diversion of air into the dome of the parachute. Others remain fairly rigid,
maintaining a fairly constant deceleration amount by not filling out more at greater velocities. The
structure of the material also extends to its porosity. More porous materials allow partial air flow
through the skin of the parachute, and this limits the force the parachute experiences, which
improves its survivability at higher velocities.
Figure 75 Key features of a ballistic parachute
(ESDU, 2009)
Another key feature relying on magnitude of force applied is the shape of the parachute. A
parachute that focuses a lot of the air flow into the dome of the parachute will undergo a large
amount of force under high velocities, and this could well mean the parachute tears or breaks
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away from the chords attaching it to the descending body. To overcome this problem, a more
aerodynamic shape is used, that lets more of the air flow around the parachute. Although this
sounds counter-productive, it is a trade-off to permit the use of parachutes at these higher
velocities.
The parachutes chosen for the simulations in this section are laid out below:
Parachute type
Function
Typical CD
Deployment velocity
Ballute
Deceleration chute
0.9
0.8 < 𝑀 < 4.0
Ringsail
Landing chute
0.78
𝑀 < 0.5
The ballute parachute has been chosen for its high deployment velocity, and if needed can provide
valuable deceleration at a critical stage in the descent. The ringsail is a reliable and stable
parachute, and can be used effectively right to the point of landing. Ringsail parachutes have also
been used with other space capsules and have proved effective (ESDU, 2009).
Figure 76 A ballute and a ringsail parachute (ESDU, 2009)
7.10.3.
VEHICLE SHAPE CHOICE
After some preliminary analysis, it is obvious that the two main different vehicle types, winged or
capsule, have their own strengths and weaknesses. The winged vehicle is more controllable, in
that it can descend whilst controlling its trajectory and direction fairly simply, and is capable of
returning to a runway. Whilst this is true, it does suffer from higher heating levels on the wing
leading edges and on the nose cone. It also becomes heavier with wings and extra control features,
and landing gears etc. that force the vehicle to be larger. A capsule has less control, generally
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returning under ballistic travel and landing under parachute, but the heating it faces is fairly
regular and can be controlled in such a way as to focus it onto the specific heat-shielded surface
designated as the base of the vehicle. Despite the lack of control over the exact place of landing,
the capsule is the more reliable of the two types, and can more easily be implemented and adapted
in future program alterations. Therefore, this is the vehicle type that analysis has been carried out
upon.
7.10.4.
HEAT SHIELD MATERIAL ANALYSIS
From preliminary analysis in the Inception Report, it was concluded that a heat shield would be
far more viable than some of the alternatives (e.g. magnetic field plasma particle deflection,
aerogel expulsion forming a regenerative layer, etc.) because of the relative simplicity and
reliability of the technology involved. The alternatives have either not been developed to a viable
stage yet, and have little forecast to mature within the allotted timeframe, or are more conceptual
and have no structural technical analysis behind them. For these reasons, a heat shield will be
employed in this report.
The majority of heat shields used are formed mainly from an array of tiles shaped to the contours
of the vehicle to create a body with a desired drag coefficient, and maintaining a heat resilient
surface to protect the contents of the vehicle. The benefits of the tile system have been laid out
above, and these reasons all contribute to the general acceptance of a tile-based heat shield on all
major space vehicle projects today. The durability of a tile, combined with its modular nature,
make them adaptable and more stable than other forms.
The tiles can either be fitted to a sub-layer already built onto the vehicle, or to a shell which is
fixed to the vehicle to allow for a consistent surface on which the tiles can be glued. This layer
has some thermal insulative properties, but also acts as a strong layer which bonds well with the
adhesive to the tiles, and therefore forms a strong tiled surface. The surface is required to be
durable, not only with regards to the heat, but the vibrations that the surface endures during
turbulent travel and take-off. Some of this comes from the dragging effect of the atmosphere, and
this energy vibrates the surface, which puts the adhesive under stress and can cause the tiles to
dislodge from the sub-surface. Another source of vibration that can stress the adhesive is the noise
during travel. As the vehicle travels, both when it takes off, and during turbulent flight, there is a
lot of noise and this causes the air particles to vibrate. The levels of noise can be very high; to
stress this point, when a rocket takes off, large volumes of water are pumped at great speed
beneath the rocket, and this is not just for counteracting the exhaust temperatures. One of the main
reasons is to allow the water to absorb the sound energy. Otherwise the vibrations could severely
damage the concrete surface of the launch pad, and the surrounding structures. With this level of
sound energy around the vehicle, the need for a strong resistant surface is obvious, and therefore
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the adhesive, and the bonding surfaces the adhesive works between, must be of sufficient quality
to withstand this stress.
With some heat shields, an external coating is applied to further enhance the performance of the
tiles, and reduce the wear they face during use. This aids in lengthening the lifespan a tile has, and
therefore reduces the maintenance costs associated with replacing and fitting tiles. This external
coating can be applied by either painting or spraying it onto the ready tile surface, and this seals
the surface. During descent, this extra layer is burnt away, and the tiles provide the durable
insulated layer beneath.
Another type of heat shield, which has only recently
really been developed and tested, and has not yet
reached general use, is the inflatable heat shield.
NASA’s HIAD system, as explained in the Inception Report, is a heat-resistant material formed of a series of
layers that have been developed and combined for their
different properties, to give an overall strong surface
capable of withstanding high temperatures. It is also
able to maintain its shape under the drag force the
vehicle encounters as it enters thicker atmosphere, and
it is a lot lighter than conventional heat shields. The
shield can be carried deflated, and just require canisters
containing compressed gas for the inflation process.
The shield is held together by a series of Kevlar belts
attached to Kevlar tubes lined with silicon (TALBERT,
Tricia, 2012). However, this design concedes a little on
Figure 77 IRVE-3 artist's concepts
composite (TALBERT, Tricia, 2012)
the maximum temperatures it can withstand, due to the
requirements it has to be inflatable and lightweight.
To start, some of the currently available materials for heat shields are the Space Shuttle’s various heat shield constituents. The Shuttle had a heat shield that was built around a number of key
vehicle points, such as landing gears, flaps and doors. The belly of the vehicle was covered in a
mixture of tiles and filler, shaped around the contours and features of the vehicle, because the reentry of the vehicle focussed a lot of heating on this region of the vehicle. However, because the
surface was fairly flat, the heating was spread out over the entire area. Therefore, the heating was
high but not too intense that tiles couldn’t withstand it. The tiles used were NASA’s HRSI (Hightemperature Reusable Surface Insulation) type, which came in two different densities depending
on strength required, and these tiles were coloured black. This was done to improve the ability of
the surface to radiate heat, so as to cool the vehicle faster and maintain an internal temperature
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more easily. This is necessary to counteract in some way the high heating phase of re-entry,
reducing the maximum temperature rise by radiating heat energy. The low density version of the
HRSI tiles were used over most of the underside of the vehicle, and the higher density HRSI tiles
were used where stresses were high, such as around landing gear doors and windows. Filler was
used between tiles to provide a firm seal that could be replaced when necessary, and that would
also fit around tiles without pushing or dislodging them when the orbiter airframe and the tiles
heat up, because they have different thermal expansion rates. Other parts of the Shuttle were
covered in LRSI (Low-temperature Reusable Surface Insulation) tiles, FRSI (Felt Reusable
Surface Insulation), FIBs (Flexible Insulation Blankets), RCC (Reinforced Carbon- Carbon)
pieces, or were exposed heat resistant glass and metals that didn’t come under excessive heating and so didn’t require high shielding. The LRSI tiles were sometimes used where heating rates weren’t so high, up to 1200 F (649 ˚C),
as they were generally lower density than HRSI tiles and were coloured white. All the tiles were
coated in reaction-cured glass made from glass powders, thickeners and pigments. FRSI was used
for low heating areas where peak temperature was below 600 F (316 ˚C), as were FIBs.
Figure 78 Orbiter thermal protection systems (NASA - JFK SPACE CENTER, 2008)
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The RCC pieces were used on the wing edges and the nose because these areas cut through the
air, and were therefore exposed to far greater heating, as they can withstand over 3000 F (1649
˚C). (NASA - JFK SPACE CENTER, 2008)
Further advances in the Shuttle program saw the tile compositions change. The first upgrade was
the development of FRCI-12 (Fibrous Refractory Composite Insulation) in 1981, which reduced
the weight of the heat shield, from tiles with a density of 22lb/cub ft (352kg/m3) to tiles with only
12lb/cub ft (192kg/m3). This material had a slightly higher thermal conductivity and lower
thermal shock resistance than the previous tiles, but these were within flight limits. These tiles
replaced the high and low density tiles in a number of areas of the Shuttle. Then in 1996, another
tile material was developed. AETB-8 (Alumina-Enhanced Thermal Barrier) took the FRCI tile
material and increased the thermal stability and conductivity whilst negligibly affecting strength
of weight. This development also came with a new coating that produced tiles known as
toughened unipiece fibrous insulation (NASA - JFK SPACE CENTER, 2008). These tiles had
higher strength with minor weight impact. The most recent advance, in 2005, was a new material
called BRI-18 (Boeing Rigid Insulation-18lb/cub ft). It was developed for its strength, which is
the highest of all the developed Shuttle tiles when coated to produce toughened unipiece fibrous
insulation. Its high impact resistance was very desirable in light of the Columbia tragedy, to be
fitted in areas with high impact risk (NASA - JFK SPACE CENTER, 2008).
A reasonably mature material is Avcoat. It has been used before, on a number of vehicles. This
material is an ablator type, in that it is designed to withstand large amounts of heating, but at the
cost of some of the material charring and falling away. This has its benefits, in that the charring of
the material uses a lot of the heat energy that is applied to the surface of the material, and whilst a
relatively low proportion of the material is lost, the amount of heat energy rejected from the
surface is high enough to be of great benefit to the vehicle and its contents. Avcoat takes the form
of silica fibres in an epoxy-novalic resin, filled in a fiberglass-phenolic honeycomb (DUNBAR,
Brian, 2009).
A more recent development is the material PICA. Developed by NASA, it has been used on craft
such as Stardust, and has proven to be effective at keeping the vehicle at a stable temperature
without requiring substantial weight. The main advantage overs Avcoat is that it has lower
weight, and better values of thermal conductivity. The name is an acronym for PhenolicImpregnated Carbon Ablator. The material can be formed in varying compositions to give it
different properties and make it more suitable to the situation in which it is employed. NASA
developed a number of variations with different material properties. However, the densities of
PICA vary mostly in the region of 0.26-0.28g/cm3 (THORNTON, Jeremy et al., 2011). Combined
with this, PICA has low thermal conductivity, which makes it very suitable as a heat shield
material. This is because low thermal conductivity means heat energy is poorly transferred
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through the material, and thus it is a good insulator. PICA has been used successfully, and as
such, it has only taken a little time for another party to take the work NASA did on PICA, and in
collaboration with NASA, develop a new variant of the material for private commercial interest.
The Space X variant PICA-X has very little available data, but it has been implemented on
vehicles in the Space X collection, and has been used successfully to deliver payloads to LEO.
For the purposes of analysis, general PICA values have been used to analyse their base suitability
for a vehicle descending from LEO.
7.10.5.
HEAT SHIELD MATERIAL CHOICE AND STUDY
After comparison of these materials, the PICA variant PICA-X looks to have the most promise. It
has been noted that PICA-X, developed by Space X in partnership with NASA, should have
superior material properties to the previous PICA variants. However, without released data on the
material, further analysis has been carried out with more generic PICA material properties, and
will be assumed to be similar or better with PICA-X tiles.
To start with, the size of the PICA-X tiles has not been particularly listed, although they are
described as being “as large as a cafeteria tray, but over 8cm thick, and weighing only about a kilogram each” (SPACEX, 2010). However, an estimation of the number of tiles required has
been made by comparing this vehicle with the proposed Orion capsule heat shield. On this heat
shield, 200 PICA tiles are used (NASA, 2008). As Orion is fairly similar in profile to the
proposed capsule, this value has been selected as a reasonable estimate of the number of PICA-X
tiles that would be required for the capsule.
The Orion capsule also features 1300 tiles that are the same as those used on the underside of the
Shuttle (SICELOFF, Steven, 2012). These protect the sides and top of the capsule, and as these
areas will come under some level of heating, this is appropriate. The tiles used are lower in
density than the PICA-X (assumption based on PICA generic values) tiles, at around half the
density. This makes them useful on the sides and top, by adding the lowest amount of weight that
can be managed, whilst still shielding these areas. A conservative estimate for the capsule
developed for this report has been set at 1600 tiles around the capsule sides and top to account for
the slight difference in diameter, and to add a cost safety factor.
An analysis using PICA generic density, thermal conductivity, thermal diffusivity and emissivity
was done through the simulation created in Microsoft Excel using time steps of 0.1 seconds. The
initial velocity of the system was taken to be the orbital velocity at the initial altitude, and then
enters as a decaying orbit profile, with an initial attack angle of 0˚. For the purposes of this
analysis, heat flux has been determined at each point during descent, and from this surface
temperature and internal temperature has been calculated. This provides a method to monitor not
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only the tiles’ survivability against the heating, but also the implied temperature change within the
capsule.
The actual simulated vehicle comes in four variants. For the purposes of showing all the study
considerations, the first vehicle is a purely-ballistic simulation of a 30 tonne capsule; the second is
a parachute-landed capsule of 30 tonnes. The third vehicle is a 9 tonne purely-ballistic capsule,
and the fourth, the main focus of this report, is the 9 tonne capsule with parachute-landing.
7.10.6.
SIMULATION RESULTS AND CONCLUSION
The first point to understand about the simulations is that, being constructed in Microsoft Excel,
they are limited and will not be anywhere near as comprehensive as a final project simulation. The
simulation lacks a changing drag coefficient, and this is because tabled values of changing drag
coefficients for parachutes and blunt bodies were unable to be obtained for the purposes of this
report. However, it should be noted that for project progression and complete analysis, drag
coefficients should be obtained from data tables for the appropriate body type in question.
Another inaccuracy with the simulation is the lack of attitude control procedures into the descent,
only having initial attack angles and velocities and falling uncontrolled from that point. Lastly, the
drag calculations are made based on the cross-sectional area of the vehicle, and this is maintained
throughout suggesting that the vehicle enters with the surface always directly perpendicular to the
line of travel. This will not necessarily be the case, but for the purposes of these simulations, it is
a simplification that will not have a great effect on the overall conclusions drawn.
The landing velocities of the vehicles, along with the maximum temperature the heat shield
reaches, and the maximum temperature reached beneath the heat shield, are important values
within the simulations. There are values of deceleration (and therefore ‘G-force’) that are calculated within the simulations, but these are based on no attitude-controlled orientation change.
The value for deceleration climbs drastically during entry into the effective atmosphere at around
120km altitude, and is sustained at these high values for around 15 seconds. This would be
problematic for the capsule, as the human payload would be in a critical situation, and may not
survive the heavy g-force. However, with correct attitude control at this point in the descent, the
problem is counteracted. The important values are outlined in the first table in Appendix 12.7.
The key points to note are that the maximum temperature the proposed capsule’s heat shield
reaches is 1348K and internally 298K if the heat shield is 150mm thick. This is very good, and
should more than adequately protect the contents of the capsule. Factoring in the carbon
composite carrier structure that sits behind the tiles, the heating experienced inside the capsule is
negligible. The total mass of the required quantity of PICA-X, taking the density to be 280kg/m3
as an approximation, would be 824.7kg. This value, when combined with the 1600 tiles (at an
approximate average of 8x8 inches or 413cm2 each, and a thickness similar to the Shuttle at 2
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inches or 5cm) that protect the rest of the capsule, with a density of 9lb/cub.ft (144kg/m3), brings
the total heat shield mass to approximately 1310kg in tiles.
7.10.7.
HEAT SHIELD FOR THE ENGINE
The viability of this project centres on an appropriate level of reusability, and it has been
determined that the engine is a highly desirable feature to reuse if possible, reducing launch costs.
Therefore, a system proposal has been conceived to allow for this. The idea developed involves a
deployable heat shield. Unfortunately, the HIAD system developed by NASA will not be capable
of protecting the engine because of the nature of the descent: the engine will descend nozzle-first,
and so this means that any shield used has to be capable of being deployed in front of the engine.
This means, in turn, that it must otherwise be stored away so as not to restrict engine function
normally, or be damaged during launch. The HIAD system deploys from a central point, when gas
is fed into the centre of the shield, filling out the sides and resulting in a shallow blunt-cone shape,
rather similar in shape to the conventional capsule base. However, the impracticality of deploying
the shield is evident. The complexity of moving the central block into position under the engine,
and then pumping gas to it, not only makes integration difficult, but does little to reduce the risk
of losing the engine entirely. Therefore, a better alternative was developed in which a solid heat
shield, much like that used on the capsule, is manufactured to be deployable in such a way that it
is normally concealed to the sides and mostly behind the engine during flight in two parts, and
then releases from the locking positions allowing the two parts to come together and lock into
position underneath the engine. Whilst requiring some level of development, the idea is sound in
that it uses tried and tested material, and functions pretty simply as a mechanism, whilst becoming
as strong as a single-piece heat shield once locked into place. The shield might look something
like this:
Figure 79 Possible abstract design of the engine heat shield
7.11.
COSTS ASSOCIATED WITH THE HEAT SHIELDS
Unfortunately exact numbers are hard to acquire for commercially-sensitive materials, so a
number of reasonable assumptions have been made with regards to the pricing. To start, the cost
to develop has been set with the knowledge that PICA-X, as a viable heat shield material, is
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already developed and could be purchased from SpaceX under contract. The other tiles, those
originally used on the Shuttle, are also pre-developed and only require production, possibly inhouse to save on transportation costs. Therefore, an estimate for development of the heat shield
has been suggested as $4,000,000 based on development of the mechanised heat shield for the
engine, and the configuration of tiles on the heat shields. Testing equipment and tools will also be
needed within this, and may cost somewhere in the region of $500,000.
The carbon composite carrier structures used by the heat shields are already used, but figures are
not available for the cost of such a piece. There may be the possibility of production in-house, but
estimated costs of these carbon composite carrier structures stand at $500,000 for the single-piece
heat shield, and $800,000 for the split-part heat shield. These pieces must be crafted precisely, as
they dictate the layout of the tiles that are laid upon them and precision and accuracy means these
pieces are fairly expensive.
The PICA-X tiles are hard to value. They are currently used by their own developer company,
SpaceX. For this report, a moderate value has been estimated based on the Shuttle. The tiles used
on the Shuttle are estimated to have cost $1000 apiece (EGGERS, Mike, 1997). At the same time,
despite being supplied by another company, the price for the tiles would be competitive.
Therefore, a reasonable estimate would be an average of $700 a tile (based on a breakdown of
$600 for the simpler tiles, and $800 for the more complex tiles). With regards to the former
Shuttle tiles, taking into account the age of the technology, and the less-restricted budget that
NASA had to hand, rather than $1000 a tile, a sensible estimate for the HRSI tiles would be
something in the region of $400 apiece.
The maintenance required, according to Space X’s chairman Elon Musk, is minimal as PICA-X
tiles could last hundreds of returns from LEO (CHAIKIN, Andrew, 2012). However, this is
probably exaggeration and a safer limit would be around 25 launches maximum per tile. On top of
this, the landing may cause damage to tiles, so this must be accounted for. A NASA estimation
says that, of 23,000, the Shuttle only had to replace on average 50 tiles from damage each flight.
This works out to be 0.2%, and this would make a reasonable estimation of the number of tiles
that could be replaced each launch. For the PICA-X tiles, this works out to be only one tile, but as
the tiles are the main contact surface with the ground on landing, a safety factor might be to say
four tiles a flight might need replacing. Of the 1600 tiles surrounding the capsule, 0.2% comes to
four tiles a flight, but again a more precautionary estimate would be ten tiles a launch.
The carbon composite carrier structures should last somewhere in the region of 25 launches, as
they are durable. However, for safety, as they come under heat loading, repeated adhesive
exposure, fatigue loading etc. and is a critical part of the heat shield, it is recommendable to
replace these after 25 launches.
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A likely workforce requirement would be 20, and the adhesive and coating might annually come
to $75,000.
To summarise, the total cost breakdown is:
(24 × 4) + 200 × 600 = $177,60
in capsule PICA-X tiles over 25 launches,
(24 × 4) + 200 × 800 = $236,800
in engine PICA-X tiles over the same period,
(24 × 10) + 1600 × $400 = $736,000
in capsule HRSI tiles over 25 launches, $1,075,000 in workforce, adhesive and coating costs
annually, and $4M in development costs initially. This comes to a starting cost of $4.5M
(development and tools/equipment), $2.45M every 25 flights on the heat shields, and an annual
cost of $1.07M
FINAL DESIGN SUMMARY
To summarise, the final design for the heat shields is a solid carbon composite carrier structure for
the underside of the capsule, fixed and with 200 PICA-X tiles arranged and fitted to it. The
surrounds of the capsule are covered in 1600 HRSI-type tiles, of the lower of the two original
densities. The engine heat shield is formed of a split carbon composite carrier structure, also
arranged and fitted with a total of 200 PICA-X tiles, and with a support and locking structure
fitted to the rear of it that allows for the two parts to slide out of the way of the engine under
launch conditions, and then to slide into position and lock in place and together to form a solid
PICA-X shield. The heat shield on the capsule should be sufficient to negate the requirement of a
separate heat protection or regulation system between the shield and the cabin. A tile thickness of
150mm is sufficient to achieve this. As a safety consideration, a backup heat protective
honeycomb layer, of appropriate material choice such as aluminium or similar, could be used to
prevent critical risk in the event of partial heat shield failure. The mass of the capsule heat
shielding amounts to a total of 1307kg, and the engine heat shield weighs a total of 825kg.
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7.12.
INFRASTRUCTURE
7.12.1.
LOCATION SELECTION
The position of the location launch facilities is very important in a space project. As the vehicle
parts will be delivered on site for assembly and checkout. The increase of the distance might lead
to the increase in pricing and greater environmental impact.
Commercial companies always intend to maximise the efficiency to provide better service. Sea
Launch is the only company launches payload from the Pacific Ocean and they provide their
launches from the equator. Due to the location, the vehicle can take advantage of the centrifugal
force (SUNGENIS, R., 2010) allowing 15%-20% more payload mass with the same amount of
fuel. Space X is not any different to Sea Launch. Space X intend to construct their complex at the
southernmost county in Texas, allowing the launches to be more efficient. However, Space X is a
good example being rejected by the local environmental authority. The Texas Parks and Wildlife
Department (TPWD) is concerned about the direct impact of noise, heat, vibration, fencing and
hazardous material to the surrounding areas. Unfortunately, three sides of the site is surrounded
south Texas Park and wildlife refuge (PARK, Minjae, 2012). On the other hand, a great deal of
the public support the propose plan, as the launch site could create job to the local and other
income to the city (PARK, Minjae, 2012), as discuss in Chapter 5.1.6. However, there will be a
portion of the public be against the idea. Moreover, this could vary in another country, since the
laws and regulation differs.
The propose design of the vehicle only uses the aid of liquid oxygen and liquid hydrogen, which
do not affect the environment as much another propulsion fuel, as discuss in Chapter 5.12.4.
However the project does not create any physical pollution but will still create noise, heat and
vibration problem. This will limit the launch location site available for selection.
7.12.2.
CONSTRUCTION
Figure 80 – Typical construction cost distribution curve (CRM TUBOR, 2010)
Launch facilities do not require excessive amounts of development compared to the vehicle. The
modern launch method is more or less the same as if in the past. However the construction
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technology nowadays is more advance and requires fewer labours. The cost of construction can be
further reduced compare to the pass. However, the type of infrastructures used is not the key
driving factors for the project cost.
The main difference between the old and the modern construction is the speed of construction
process and the reduction of risk occurring. Therefore a new design of launch facilities might not
be needed and the new designs might not provide enough advantages to overcome the risk that
could occur. In addition, the most expensive stage of construction industry is the construction
stage and not the design stage, where the labour force being one of the main components, Figure
80 show the cost distribution of a typical construction project. Thus, using the same or even
similar infrastructure design as previous project can further reduce the cost of the infrastructure.
Furthermore, the vehicle should be design to or could be modify easily to adapt to any launch pad.
This is because the launch system proposed is to sell as many customers as possible and the
vehicle is proposed with the capability to launch from anywhere on Earth. The customer might
already have the launch facility available, such as NASA. If the vehicle is design with wider range
of specifications will further increase the customer range. Thus, the infrastructure should be
similar to other launch infrastructure such as Complex 39.
7.12.3.
INFRASTRUCTURE REQUIREMENTS
The require launch facilities are govern by how the launch system is designed. For example many
current space Launch Companies operate slightly different. The Space Shuttle needs the aid of a
runway for landing, whereas the Dragon is retrieved from the sea. Therefore, there is some
essential infrastructure needed to provide the launch service.
Orbiter/payload processing facility is to provide control environment for assembled and checkout
before the vehicle integration. The integration of the vehicle is carry out from the vehicle
assembly building, where it also require a control environment for assembling the launch vehicle.
The Launch Control Centre is basically the command centre of the operation. Last but not the
least, the launch pad is to carry out the action during launches, Appendix 12.6.2 shows more
information of what is the purpose and what is included in the buildings.
There are some other infrastructures are less important in comparison. This is because it is to
design to support the main infrastructures. For example, voltage electrical substation,
communications and electronics, cable terminal building and propellant system components
laboratory are the buildings that provide supports to the entire launch complex. Launch equipment
shop, ordnance storage facility and instrumentation building are the building that provide the
support the vehicle for assembled and checkout. A series of supporting infrastructure needed
before the vehicle can be transported to the launch pad to carry out its mission. There will need to
be a method to transport the vehicle from the Vehicle assembly building to the pad and with the
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aid of the launcher, which hold the vehicle in the launch position. In addition, a service structure
is needed to access the vehicle when on the launch pad either for final checking or loading human
in to the vehicle before the mission.
The main purpose of the infrastructures can be considered as a tool to carry out all the essential
tasks before the launches and the main objective is allowing smoother operation for the ground
segment. NASA is one of the leading organisations in the space industry. Most of the launch
complexes are still in use, such as the Launch Complex 39 named Kennedy Space Center (KSC)
with a history of 50 years (NASA-KSC, 2012). The launch pad 39A is under modification for
Falcon Heavy (CLARK, STEPHEN, 2012). Thus, this complex is one of the successful spaceport,
as it can provide service to other design. Thus, the require infrastructure need will be similar to
KSC and in Appendix 12.6.2 show all the essential infrastructures from complex 39.
7.12.4.
STORING AND TRANSPORTATION OF HYDROGEN
The major concern of dealing with liquid hydrogen is liquid boil-off. It is a cryogenic liquid and
any heat transfer can cause hydrogen to evaporate. This is including rtho-to-para conversion,
mixing or pumping energy, radiant heating, convection heating or conduction heating can all
cause loss (AMOS, Wade, 2000).
Storing the fuel in cryogenic tanks can prevent this matter. The tanks are constructed with double
wall and the space in between is evacuated to further reduce convection and conduction loss.
Multiple layers of reflective, low-emittance heat shielding are installed to prevent radiant heat
transfer (AMOS, Wade, 2000). Moreover, some vessels have an additional outer wall with liquid
nitrogen in between to further reduce the heat transfer by lowering the temperature different.
Majority of the tanks are spherical. The shape has the lowest surface area for heat transfer per unit
volume. It is more efficient to store in a larger diameter tanks, because the volume increases faster
than the surface areas. However, the Space Shuttle burned about 132M kg of liquid hydrogen with
63.1M kg depleted by storage boil off and transfer operations (NASA-CLIMATE, 2012). Thus,
each flight there is 33% fuel wasted.
Table 27 – Price comparison of trucker and factory of hydrogen, data collect from (AMOS, Wade, 2000)
On site storage
Liquefaction Capitals Cost
Liquefaction Operation Cost-Compression equipment
-Liquid hydrogen tank
-Electrical energy
Transportation Operation Cost
Transportation Capitals Cost
Boil-off losses
Trucker
$31-$700/kg
N/A
N/A
N/A
N/A
$1.8-$2.1/kg
$350,000/4080kg
10%-20% up to 50%
Factory
$31-$700/kg
$25600
$0.08
$0.13
$0.99
$0.39-$1.2/kg
$620,000/km
Re-liquefy
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A research shows 10% to 20% are loss during transportation and it could be as high as 50%
(AMOS, Wade, 2000). However, setting up factories on-site can solve this problem. The boil-off
fuel can be re-liquefying letting the system more environmentally friendly. Producing about
8390kg can overcome the factory cost, by using the case study figure from Table 27. The analysis
is calculated from the price of $5.5/kg for liquid hydrogen versus the production from $1.25/kg
using the aid of hydrogen gas. The 8390kg production cost 55% in capitals, 22% in operation and
23% in the purchase of hydrogen gas. The cost per unit can be further reduced as the amount of
liquid hydrogen needed. However, the cost does not include the transportation of the hydrogen
gas. If the location is far away from the source of the supply can have a huge impact of the factory
setup cost. This suggests the solution of manufacturing liquid hydrogen might not be ideal for all
applications.
7.12.5.
FUTURE PROOFING
The basic infrastructures are needed to provide launch services are described in the previous
discussion. The landing facility is not needed for this project and will not be constructed in the
initial stage. However, there are more and more spaceports built currently and to compete with
others will need to provide more and better service. As suborbital flight is one of the emerging
industries and most of the design to land on runway, thus this market cannot be ignored and close
monitor of this market is needed. Therefore, if the demand of such infrastructure is great enough
to eliminate the annual maintenance cost, there will not be reason not providing the service.
On the other hand, suborbital flight is not the main target, but the launch system will be targeting
the orbital flight when the price per launch decreases. Thus a landing facility will become an
essential infrastructure in this launch system.
Moreover, the use of liquid hydrogen in commercial flight will soon become the replacement fuel
of kerosene, as discuss in Chapter 5.12.4. Thus, the spaceport will not require any modification as
a normal airport, which provided a great opportunity to enter the commercial flight industries. The
Hong Kong international airport has 2 runways with a length of 3.8 km (HONG KONG
INTERNATIONAL AIRPORT, 2012) and produced a turnover of $1B in 2010 (HONG KONG
INTERNATIONAL AIRPORT, 2010). The commercial runway can handle about 290 km/h of
160 tonnes for flapless landing and about 225 km/h of 180 tonnes for typical landing with normal
condition, according to a conversation with a senior pilot from Dragon Air. The requirement of a
space vehicle will be even higher. Thus the construction of a spaceport runway will have the
capacity to deal with the commercial industry without any issue. The runway for Complex 39 is
about 4.6 km (Dumoulin 1993) where the Hong Kong Airport is only 3.8 km (CIVIL AVIATION
DEPARTMENT, 2013). This can further decrease the cost per launch, as the commercial flight
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could spread out the cost of the runway. In addition, if the spaceport can provide cheaper fuel by
creating on-site liquid hydrogen manufacture will attract more investors for the spaceport.
7.12.6.
CONCLUSION
The design of the vehicle is more or less a traditional rocket system. Thus the infrastructures
required could be simple. Complex construction designs are becoming less of a challenge in
comparison to other elements of the project, as the construction technology is becoming more
advance. In addition, the space launch industry does not require exceedingly large amounts of
complex infrastructure. The improvement can be achieved by being more energy efficient and
making innovations on the design. This can further reduce the price of construction and
maintenance cost. Furthermore, this could cut down the operation cost, since the infrastructure
could reduce the operation require.
Runways for the complex will not be needed till later in the project, because the tourists’ space vehicle will not be in service in the earlier stage. Furthermore, the commercial flight industry will
be governed by the location of the infrastructure. In another word, this will not be advice to all
potential consumers for the project, since not all locations are ideal. The last but not least, the
construction of liquid hydrogen plans can reduce the cost per launch and increase the potential
usage of the spaceport as an airport for commercial flight industry. However, this will be also
governed by the location of the spaceport, as the transportation of hydrogen gas becomes a
challenge. On the other hand, this could change as the more innovated and sustainable methods
expect to be available in 2040, as descried in Chapter 5.12.4.
7.13.
FINANCIAL ANALYSIS
The financial analysis will be able to prove the project is financially viable and increase the
confident for investors or customers. The analysis will show the require amount of money needed
to the according years. The income statement will illustrate the money needed each year. The
balance sheet shows the financial condition of each year of the project. An example will be shown
in Appendix 12.7.3 and Appendix 12.7.4 respectively, which is predicted by analysis undertaken
in Chapter 136. Some basic income statement and balance sheet analysis can also undertaken for
customer, as illustrated in Appendix 12.7.3 and Appendix 12.7.4 respectively. In addition, an
example of the timeline costing throughout project will be shown in appendix 12.7.2.
One of the longest and the most expensive element of the project is the vehicle. It takes six years
with $532M to develop and take one year with $41.82M to manufacture. It needed about $8.21M
to overhaul after every launch. However, the engine can be maximally reused 50 times and the reentry system can be maximally reused 25 times, which will have a replacement cost of $15M and
$0.96M respectively.
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The human module can cost up to $550M to develop and about $246M to manufacture. It costs
about $37.5M to maintain. However, to deliver mechanical payload modules requires much less
in comparison. It cost $0.5M for the fairing and $1.4M for the integration per launch. This is
because the fairing is not design to be reusable. The infrastructure cost about $217M to construct
and take three years head start before the launches. It requires annual maintenance of $57M. The
overall employee for operation cost about $302M annually when the service is in place. Each
launch will cost about $0.32M for fuel. The costs estimation of each element is based on the
previous chapters and the price will be illustrated in Appendix 12.7.1.
The vehicle is design to deliver 30 tonnes of payload. It can either deliver full 30 tonnes of
mechanical payload or 20 tonnes payload with human module, which is 10 tonnes. The vehicle
can generate revenue of $390M each year after development by delivering payload, if the project
can match the design requirement and achieve $1,000/kg and a launch every 4 weeks. The
analysis in Chapter 7.13.1 and Chapter 7.13.2 are carried out without the use of discount rate,
which is the sum of the interest rate and the risk factor. This is because it will be simpler to
compare with other analysis in this chapter. In addition, the unforeseen risk can affect the project,
however cannot be confirmed due to the stage of the design. This will be further discussed in the
Chapters 8 and a sensitivity analysis will be analysing the impact of the discount rate in Chapter
7.13.3.
7.13.1.
30 TONNES MECHANICAL PAYLOAD
As described in previously the cost of fairing production and integration is $2M per launch. This
will not produce any profit even if two vehicles are manufactured and it met the target vehicle
lifecycles of 200 launches. This is due to the relatively high costs related to development,
construction and operation, as shown in Appendix 12.7.3 Figure 89. However, increasing the
price to $1500/kg or manufacturing three vehicles with price of $1000/kg can allow the project to
generate profit, as show in Appendix 12.7.3 Figure 90 and Figure 91 respectively.
7.13.2.
HUMAN MODULES
The development of cargo transfer module should be finish within a year; however the vehicle is
design to deliver human modules with extra payload in the same time, which will take extra two
year for development. The human module weighs approximately 10 tonnes and can deliver about
3 tonnes of cargo along with 6 people.
For delivering humans to space requires a 100% success rate of the vehicle. This means
development of the human module is not required until later in the project when this reliability
has been achieved. Nonetheless, the vehicle will need to prove its capability to the investors and
potential customers, to further invest in the project or purchase the service of the human modules.
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This can reduce the initial development budget require and allowing the project to generate small
amount of revenue by the mechanical payload at the initial launch and reduce the risk of the
investors.
Nevertheless, the previous scenario will be ignored to determine the profit margin for the financial
analysis purpose. The require amount of delivery by human module is assume to be 13 time a year
and it can be reused 13 times, which is one vehicle launches a year. Replacement will be need
after each year. The usage of the human module to deliver 3 tonnes of cargo and 20 tonnes of
payload by fairing could generate profit shown in Appendix 12.7.6 Figure 92, which cost
$7000/kg and $1000/kg respectively. The price of the human module delivery can decrease to half
with $3500/kg shown in Figure 93, if three vehicles are in operation. This price could be expected
for the delivery of human; however it can further reduced where the cost of the other system is
also included.
7.13.3.
SENSITIVITY ANALYSIS
The space launch industry can be affected by many factors. The vehicle itself and all other
supporting components are a complex design system and for a successful mission none of the
elements can be missing. If one of the delegated elements fails completely or even partially, this
could have an impact of the project. If fail to deliver human to space could have a part of the
project immediately terminated or even huge financial crisis, where mechanical payload tends to
only have financial problem. However, the project is more or less an integration of different
current technology that minimise such catastrophic event by lowering the risks. The following
sensitivity analysis will be carried out based on the example of three vehicles initially launch at
year 7, 12 and 13, including 13 human modules launch each year for the entire design life of the
vehicles, as shown in Figure 93. The analysis will be only changing variable one at a time.
Reducing the reusability of engine can impact the net revenue, as shown in Figure 94. The project
will stop producing profit by the reusability reduce below 10 times. The net revenue only decrease
about 10% even the reusability of the tank reduced 80%, 40 time reusable. The decrease of 80%
reusability for the fuel tank will have about 10% decreases in the revenue, as shown in Figure 95.
The net revenue decreases about 6% as the re-entry system decreases its reusability by 80%, as
shown in Figure 96. The biggest cryogenic fuel customer was probably is the Space Shuttle of
NASA at the time. The decrease of 75% in fuel price can increase the net revenue by about 20%,
as shown in Figure 97. The infrastructure construction cost can have up to 50% increase in the net
revenue by decreasing the cost by half, as shown in Figure 98. This could be possible as the cost
is base from NASA’s information and Russian tends perform the similar project with half of the cost. In addition, the maintenance cost of the will be able to reduce as a result and will produce
| Page 137 of 262
further revenue. The last but not the least, the human module will bring 70% less net revenue if
the module is manufacture every launches, as shown in Figure 99.
The typical interest rate is 3.5% and the risk will vary with technology according to the year of
usage. For a new and innovative technology will have a risk factor as high as 9%. The pricing for
the delivery may be set too low. The net present value (NPV) suggests the project can be gain
profit only with the rate around 3%, as shown in Figure 100. This propose to achieve the
according price will need a source of low interest grant and relative low risk.
7.13.4.
CONCLUSION
The project can match the $1000/kg specification and generating profit, when 3 vehicles in
service. Meanwhile, the human module will be able to provide service with a price of $3500/kg.
Some costing of elements is suggested to be lowered such as the infrastructure, if the project
includes the discount rate. However, the rate could be relatively low. This is because the present
of government grant and the usage of the existing technology, lowering the discount rate.
However, it will depend on where the project is to take place, as this will have an impact.
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8.
RISK ASSESSMENT
RISK ASSESSMENT
Assessment of the risk was conducted throughout the whole design procedure and collectively
identified from the final design proposal. The risks were split into the causing and effect. The
likelihood and severity the risk were calculating using Appendix 12.8.1 Table 28 and Table 29. A
risk control measure, RCM, such placing fire extinguishers in fire hazard zones, was then devised
to lower the likelihood and/or the severity of the risk. This risk was given a rating as described in
Table 30.
A post-RCM analysis was conducted to see if the proposed RCM would help reduce the
likelihood and severity. If the risk was still rated as medium or high a further RCM was suggested
that might further reduce the risks involved.
A total of 26 priority risks were identified of which 18 were technical based, 5 were H&S based
and 3 were environmental based. Implementation of RCM’s to these resulted in a medium risk to technical aspects, a low risk to H&S and a low risk to the environment.
Moreover, the project will need to be carried out with a further analysis from NASA/Air Force
Cost Model (NAFCOM). This is analysis will run by base on the historical aerospace project, as
shown in Appendix 12.8.2 for the application form applied. This will indicted more risks during
the design stage, such as the finical risk of the project. Hence, the project is in place is ready for
carrying out next stage of the design.
Author: Group
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Causes
Tiles dislodging
Engine Failure
Engine Failure
Turbopump Fails
Engine Exhaust Overheat
Engine Maintenance Delayed
Failure Engine Inspection
Shield Failure
Shield Failure
Faulty tile
Faulty carbon composite carrier structure
Manufacture
Trajectory
Leakage of Tank
Boil-‐off
Boil-‐off
Human capsule
Human capsule
Human capsule
Fuel
Location
Pollution
Single Event Latchup
Power source failure
Stage Separation Failure
3 GP1 H&S
4 GP1 H&S
5 GP1 Tech
6 GP1 H&S
7 GP1 Tech
8 GP1 Tech
9 GP1 Tech
10 GP1 Tech
11 GP1 Tech
12 GP1 Tech
13 GP1 H&S
14 GP1 Tech
15 GP1 Tech
16 GP1 Tech
17 GP1 Tech
18 GP1 Tech
19 GP1 Tech
20 GP1 Tech
21 GP1 H&S
22 GP1 Env
23 GP1 Env
24 GP1 Tech
25 GP1 Tech
26 GP1 Tech
Technical 18 No. Medium
Incorrect re-‐entry profile
Production fault
Production fault
Equipment Misuse
Lost control of the landing location
Fuel leakage
Residual fuel boiloff leading to explosion
Lost of fuel leading to explosion
Failure of life support systems
Ammonia leakage
On-‐board fire
Cryogenic skin burn
Impact to the local environment
Polluction cause by launches
Radiation induced short circuit
Full spacecraft system failure
Launch Failure or Abort
Engine fails
Engine: Combustion Chamber Failure
insufficient thrust to continue flight
Tank punctured by debris causing fuel leak
Rocket explodes
Rocket explodes and falls to land in a populated area
Insufficent fuel flow for engine causing shut down
People or equipment nearby may ignite
Rocket turn around time affected
2 GP1 Tech
Hazard
Engine: Combustion Chamber Failure
By
1 GP1 Tech
Risk No.
H&S Env Tech
Owner
3
2
2
3
3
2
2
3
2
4
2
3
2
1
1
3
2
2
2
3
2
2
2
2
2
2
5
4
4
4
4
5
5
3
5
5
5
5
3
4
5
4
5
5
3
2
5
4
5
5
5
4
Risk Control Measure(RCM)
Pre-‐re-‐entry checks, and regular maintenance
Inspection and backup guidance systems
Detailed mission planning
Detailed mission planning
Limiting the toxic materials and fuel used
High
Environmental 3 No. Low
Failsafe system with redundancies
Medium Emergency Generator
Radiation shielding and hardware Medium
redundancies
High
High
Inspection and backup guidance systems
High
Inspection circults system
Medium Wear personal protection equiment
Choose location with lower environment High
impact
High
High
High
Detailed mission planning and backup guidance systems
Medium Full testing of tiles
Medium Detailed inspection prior to fitting
High
H&S inspections and floor manager
Detailed mission planning and backup Medium
guidance systems
High
Detailed inspection before refuelling
High
High
1
2
1
3
3
1
1
3
1
2
1
2
1
1
1
1
1
1
1
1
1
1
Multiple turbopumps used, with redundancy and failure resistant design
1
1
1
1
Clear site prior to launch
Post ROM's
5
1
2
2
2
5
5
2
5
5
5
5
3
3
5
4
5
5
3
2
1
2
1
2
2
2
Residual Hazards
Desing allowing redundancy
Desing allowing redundancy
Medium
Low
Low
Statistical risk analysis of specifc modes of failure
Medium Montioring the local polluction level
Medium Limiting elements cause the impact
Medium Desing allowing redundancy
Medium equired fire extinguisher on-‐board
Medium Health and safety training
High
High
Low
Low
Medium Health and safety training
Medium Allow on-‐board patch repair system
Low
Low
Low
Low
Low
Low
Low
Low
Likehood Severity Risk Rating
Do not fly over populated areas
Escape tower for human payload
Shielding between combustion chambers and tankage
Rocket overhaul time requires sufficient float -‐ Medium
3 weeks
Take care during inspection, provide rigid Medium
inspection process and manual checklists
High
Medium
High
High
High
Multiple combustion chambers with engine Medium
out capability
Likelihood Severity Risk Rating
Health & Safety 5 No. Low
Launch Vehicle
Human Payload
Human Payload
Public
Public
Human payload
Human payload
Operators
Human payload
Space Environment
Spacecraft
Spacecraft
Spacecraft
Space craft/Engine
Space craft/Engine
Operators
Space craft/Engine
Space craft/Engine
Spacecraft
Spacecraft
Staff/equipment
Spacecraft
The public
Crew
Launch Vehicle
Customers
Comments/ Persons Affected
Pre RCM's
Owner
9.
TECHNICAL REVIEW
TECHNICAL REVIEW
The launch vehicle propulsion system involves a 24 chamber, fully pressure compensating
aerospike nozzle that will use LH2/Lox as the propellant. This combination provides a high thrust
to weight ratio and increased flight performance in comparison to traditional bell nozzles. A
single nozzle will be used to decrease operational complexity and increase efficiency. The main
engine and primary components will be fully reusable up to 10 times by utilising a deployable
heat shield and parachute system for re-entry.
The vehicle layout and propellant feeding system will be a feeding packet design where the
propellant is cross-fed only from current stage tanks. This means that at the end each stage the
remaining tanks are full and offers the optimal fuel mass fraction throughout the launch. The
design involves 3 separation stages with the 1st dropping 4 tanks, the 2nd dropping 2 tanks and the
final stage being only a single tank.
The vehicle trajectory will see the payload, 3rd stage tanks and engine reach an orbital altitude of
330km and an orbital velocity of 7.69km/s. A minimum acceleration of 1.5G will occur at stage
ignition and a maximum acceleration of 4 G at stage cut-out. It will experience an axial load of
3.4G and a lateral load of 2G. The maximum deliverable payload to a LEO of 330km will be 30T
and to GTO will be 9.5T.
In depth tank design has been conducted and finalised; the tanks will be made from Al-Li and be
3m in diameter having standard lengths of 21m, 25m and 31m to decrease production costs whilst
allowing increased modularity. The proposed design should be very similar to a final design and
gives a further estimation that the final structural mass will be approximately 17 tonnes. This is
well below the estimated structural mass value of 34.5T used in the final engine mass calculations
thus providing a large margin for increasing the vehicle mass structure if needs be.
It has been decided that heat shielding will use 2 varieties of insulating tiles which when used in
conjunction with the capsule with result in a leading edge temperature of 1348K and an internal
temperature of 298K. The total weight of this shielding system is 1310kg which is highly
comparable to other capsules in the market.
To allow for human passage to space a pressurised capsule has been proposed. It will be carried
on top of a modified fairing that may also carry additional cargo to offset the launch costs. This
will be capable of sustaining a habitable environment over a period of days and will be primarily
used to transport scientists and eventually tourists to LEO destinations such as the ISS. Once the
in-orbit mission has been completed it will initiate re-entry using on board thrusters. To cope with
the immense thermal stress of re-entry the vehicle will be fitted with 200 Pica X tiles on the lower
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surface and 1600 HRSI tiles on the upper surface. It will deploy a ballute parachute once its
velocity drops below Mach 4 to assist with high altitude deceleration and finally deploy a ringsail
landing chute to provide the capsule with a soft oceanic landing.
Hardware used in the launch vehicle will be either COTS for non-direct mission related items
such as batteries, antennas and filtration for the fairing. Hardware items directly related to the
mission such as the propellant feeding system, payload separation devices and launch abort
systems will be developed in-house to reduce production cost and ensure a fully reliable and
sustainable solution has been achieved. This will also allow lower level hardware modifications to
be achieved without intervention from suppliers.
Reusability has been a key driver in the design of the proposed launch vehicle. As it stands the
proposed design could be up to 92.5% reusable. The key components currently identified as being
non-reusable are the 3rd stage tank due to weight issues, the fairing body and parts of the
separation system. This would be a market leading system in terms of reusability. Proposals to use
the 3rd tanks as a frame for space stations in LEO or lunar orbit may further increase the overall
reusability and provide a new source of revenue.
The reusability of the 1st and 2nd stage tanks will be achieved using a parafoil packed around the
wall structure. This will deploy once the tanks have fully re-entered the atmosphere, at an altitude
of approximately 8km. This proposal will allow the tanks to be guided to a soft landing either on
land, pending flight certification, or sea with the use of suitable sleds. Due to the lightweight tank
structure and short downrange separation distance, the 1st stage tanks could be flown back by
helicopter and ready for re-use in a short amount of time. The 2nd stage tanks, which travel a much
greater distance, could be transported back over a longer period of time or even be sold to a local
launch agency for re-use in closer launch pad.
The required infrastructure for the related launch vehicle will be adaptable to a multitude of
environments. This will open the market up to more potential customers perhaps using a standard
modular site plan to allow cheap facility production. One of the largest concerns noted was the
production of the propellants, especially liquid hydrogen due to the extremely low temperatures
required. It has been decided that for all but the smallest scale launch facilities would benefit from
in-house production of liquid hydrogen.
An in depth financial analysis has been conducted and has found that the project can match the
$1000/kg to LEO price target. This requires a minimum production of 3 vehicles over the project
life time and sustaining a minimum of 13 launches per year. Profitability will be reached 10 years
after the project initiation. The proposed capsule would be able carry humans to orbit at a price of
$4M per head. This is substantially less than the current market price and will be highly
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competitive with the primary competitor, Space X. For deliveries to GTO the cost will be
$3100/kg but will potentially result in the loss of an expensive engine.
Environmental consideration has been taken in every aspect of the project. Using LH2/Lox as the
propellant significantly reduces the nitrous oxide emissions in comparison to kerosene based
systems. The high reusability percentage further decreases the environmental impact respective to
current usage. One aspect of making such an appealing launch system is that it may result in very
high vehicle production rates and thus result in an increase in pollution due to heavy usage.
Investment in green energy sources for fuel production will not only help further reduce the
systems environmental impact but may also spread to other sectors helping provide momentum to
further advances in green technologies.
Another consideration throughout the design process has been the ever growing concern over
space debris. The proposed vehicle has not quite met the design requirements for this as the 3 rd
stage could potentially be left in orbit for long durations. Further research should be conducted
into lightweight de-orbiting systems which could be attached to the vehicle. Other than the 3 rd
stage tank the vehicle would not further contribute to space debris.
There are many aspects of future proofing that are available in the area of launch vehicles and
payload delivery. Not all of these are directly project related but should be considered for further
expansion into the aerospace sector or to further human knowledge of the universe around us.
The primary future proofing target should be a heavy lifting launch vehicle. Market research has
shown that there are proposals by competitors to launch up to 120T into LEO. Due to the
modularity of the proposed vehicle it is not unfeasible that an increase in launch capacity just
needs more or longer tanks coupled with multiple engines or a single larger one.
Bigelow Inflatable Structure: The requirement for a permanent human presence in space is ever
growing. Further investigation should be taken into working in conjunction with Bigelow
Aerospace. The proposed launch vehicle is estimated to be large enough to accommodate the
deflated structure and could be serviced by the proposed human capsule.
To better service the delivery market outside of LEO, a space tug system has been proposed.
Primarily, further research should be conducted into the LH2/Lox based version as this would be
technically similar to the proposed launch vehicle. This type of vehicle would enable a fully
reusable GTO delivery method.
As a long term investment, a space tug using ion jet propulsion this could be developed and used
not only with the proposed launch vehicle but also service other launch providers. It’s possible
uses range from a very efficient GTO delivery method to enabling further 3rd stage reusability.
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A key aspect to producing a sustainable space based economy is reliability. If the proposed launch
vehicle has a poor track record of launch success then potential customers may shy away to more
expensive but reliable solutions. Throughout the whole design process careful consideration has
been taken to include large margins of safety. Key areas identified that are especially at risk to
being unreliable are the propellant management systems; the stage separation systems and the reentry systems. It is essential that further research into reliable engineering solutions in these areas
is undertaken.
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10. CONCLUSION
The resulting spacecraft and company name was developed (C3x rocket and APEIRON capsule
after the Greek word for infinite) the technologies used have led to a system that has met the
aforementioned targets. It has an estimated developmental time of 7 years and employs a three
stage conventional rocket design and single 24 chamber aerospike nozzle to improve flight
performance and operational complexity.
The proposed launch vehicle offers greater than 92.5% reusability (by dry mass, excluding
payload). The rocket would achieve a benign launch environment for both crew and payload and
delivers its third stage to an altitude of 330km at an orbital velocity of 7.69km/s and does not
exceed 4Gs. The propellant selected was LH2/LOX and has a large application/ potential revenue
in the future space port provisions as well as being more readily available as compared to the
declining amounts of fossil fuels.
The C3x rocket has a cost/kilogram price tag of $1000 better than its direct competitor the Falcon
9 Heavy ($2340 per kilogram). This makes it possible for financial break-even by year 15 and
$2.8 b net profit realisation by year 28. Confidence can be placed in these numbers due to
sensitivity analysis showing an 80% decline in the predicted reusability as acceptable.
Developmental costs have been minimised; in comparison to Skylon, which also offers $1000/kg
at a cost of $12b, this proposal will cost only $1.8b and carries significantly reduced technical
risks.
The tanks are re-usable, with inspection, for the 200 flight vehicle life. The engines are reusable
for 50 flights with refurbishment at 10 flight overhauls. The APEIRON capsule will be re-usable
13 times due to its Pica-X tiles. The course of the project was diverted slightly as new ideas and
integration considerations arose. There were also some changes done to the Gantt chart due to
coursework deadlines for other engagements and these can be seen in the Deadline and Meeting
Calendar and Project Management sections. Recommendations have been suggested in each of the
subject areas considered to carry the investigation further and develop in years when technology
and or markets develop.
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Figure 81 Vehicle Comparison against identified competitors
Ariane 5
1st
stage Vulcain
Falcon 9 v1.1
Falcon
9 LWT
MED
Heavy
Proposal
Proposal
Merlin 1D
Merlin 1D
Aerospike
Aerospike
engine
Diameter
9m
3.6m
10.8
9
9
Stages
3
2
3
2
3
Mass to LEO
21T
10.5T
53T
14T
30T
LEO $/kg
$10,000
$4,100
$1,500
$2,000
$1,000
Mass to GTO
10.5T
4.5T
12T
8.5T
22T
GTO $/kg
$20,000
$9,475
$6,625
$3,300
$1,300
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| Page 162 of 262
APPENDIX
12. APPENDIX
12.1.
MARKETING
12.1.1.
2011 COMMERCIAL LAUNCH EVENTS
Figure 82 – 2011 worldwide commercial launch events (FEDERAL AVIATION ADMINISTRATION, 2012)
| Page 163 of 262
12.1.2.
APPENDIX
2011 NON-COMMERCIAL LAUNCH EVENTS
Figure 83 – Payload type delivery by non-commercial launches (FEDERAL AVIATION ADMINISTRATION,
2012)
12.1.3.

ORBITAL CLASSIFICATION
Geosynchronous Earth orbit (GSO): A spacecraft in GSO is synchronized with the
Earth’s rotation, orbiting once every 24 hours, and appears to an observer on the ground to be stationary in the sky.

Geostationary Earth orbit (GEO): GEO is a broad category used for any circular orbit at
an altitude of 35,852km (22,277 miles) with a low inclination (over the equator).

Non-geosynchronous orbit (NGSO): NGSO satellites are those in orbits other than GEO,
including:
o
Low Earth orbit (LEO): lowest achievable orbit, about 2,400 kilometers,
o
Medium Earth orbit (MEO): 2,400 kilometers to GEO,
o
Elliptical (ELI): a highly elliptical orbit,
o
External (EXT): used for trajectories beyond GEO (such as interplanetary
trajectories), and
o
Sun-synchronous orbit (SSO): an orbit that passes over the same part of the Earth
at roughly the same time each day.
| Page 164 of 262
12.1.4.
PAYLOAD WEIGHT CLASSIFICATION

Micro: up to 91kg

Small: 92 to 907kg

Medium: 908kg to 2,268kg

Intermediate: 2,269kg to 4536kg

Large: 4,537kg to 9,072kg

Heavy: Greater than 9,072kg
12.1.5.

APPENDIX
PAYLOAD USAGE CLASSIFICATION
Classified: Any system whose purpose is officially deemed classified or cannot be
officially verified.

Communications: Any system designed to receive and transmit data for purposes of
facilitating communications. This includes fixed satellite services, mobile satellite
services, military communications, store-and-forward systems, asset tracking, and similar.

Crewed: Any system designed primarily to transport humans into, through, or back from
space.

Development: Any system whose purpose is to advance hardware design as part of a
research and development program.

ISS: Any system designed primarily to transport cargo into, through, or back from the
International Space Station (ISS).

Meteorological: Any system designed to monitor the Earth’s weather for forecasting and issuing weather watches and warnings.

Navigation: Any system designed to provide signals for accurate timing, positioning, and
navigation.

Remote Sensing: Any civil and commercial system designed to gather data by means of
optical (panchromatic, multispectral, or hyperspectral) or radar sensors.

Scientific: Any system designed to gather data about astrophysics, astronomy, biology,
cosmology, celestial bodies, physics, and the space environment. This designation also
includes systems designed to monitor the Earth, except those systems designed
specifically for meteorology.

Test: Any system designed to provide telemetry and data on launch vehicle performance.

Unknown: Any system whose mission is unknown.

Other: Any system whose purpose does not fit in any of the provided categories.
| Page 165 of 262
12.1.6.
APPENDIX
ECONOMY IMPACT BY INDUSTRIES
Commercial space transportation and enabled industries (CST&EI) and enabled industries include
the following:
 The sales of all commercial satellites constructed manufacture by USA
 Satellite services such as direct-to-home television (DTH TV), very small aperture
terminal (VSAT) services, satellite data services, transponder leasing, satellite digital
audio radio services (DARS) and mobile satellite telephony
 For Ground equipment such as satellite-related hardware: gateways and satellite control
stations, which also include mobile uplink equipment, VSAT terminals and consumer
electronics used with satellite service
 Sataellite remote sensing cover the raw satellite imagery, but excluding geographic
information systems (GIS)
 The last sector is distribution industries being the wholesale, retail trade and transit costs.
Distribution industries is consider as one of the components because it delivered require
part from the manufacture and transportation for the launch site
Figure 84 – Economy impact cause by the industries affect (ADMINISTRATION, FEDERAL AVIATION, 2010)
| Page 166 of 262
RAMP DESIGN
Figure 85 – Ramp Design with steel structure
12.2.
APPENDIX
| Page 167 of 262
APPENDIX
Figure 86 – Calculation and assumption of the Ramp Design
| Page 168 of 262
APPENDIX
Figure 87 – Refining the Ramp Design
| Page 169 of 262
12.3.
APPENDIX
EXISTING TUNNELS
The longest tunnel – Railway tunnel (ALPTRANSIT GOTTHARD AG, 2010)
Gotthard Base Tunnel 2016





Length 57km
2 km below mountain (Maximum overburden)
Diameter 8.83–9.58 m
Total project cost $10.3B
Amount of excavated rock: 28,200,000 t
Drilling tool (TBM)






Total length: 440 m (1,440 ft) (including back-up equipment)
Total weight: 3,000 t (3,300 short tons; 3,000 long tons)
Power: 5 MW
Max. excavation daily: 25–30 m (82–98 ft) (in excellent rock conditions)
Total excavation length by TBM: about 45 km (28 mi) (for each tube)
Manufacturer: Herrenknecht, Schwanau, Germany
The Largest tunnel – Highway tunnel (WALLIS, Paula, 2010)
Orlovski Highway Tunnel under the Neva River - St Petersburg, Russia






Length 6.0km
Under river
Diameter 19.25 m
soft ground including sand and clay
Manufactured by Herrenknecht AG
The Largest Hard-rock Tunnel (ONTARIO POWER GENERATION, 2013)
Canada's Niagara Tunnel Project






Length 10.4km
140m below ground (Maximum overburden)
Diameter 14.4 m
1.6 million cubic meters of rock (hard rock)
Manufactured by The Robbins Company
| Page 170 of 262
!
!
!
Boeings CST-100
capsule
Dream Chaser
space plane
!
Boeings CST100 capsule
Dream Chaser
space plane
Extra 10-20 ton
LH!
Utilizes!carbon!
neutral!liquid!
hydrogen/oxygen!
fuel!
!
Orion capsule
Utilizes!carbon!
neutral!liquid!
hydrogen!
(LH)/oxygen!fuel!
!
Orion capsule
Environmental
Impact
Extra
1!
Zero CO2!
Greenhouse!Effect! 1!
CO2 pollution!
Zero!CO2!
Shuttle'derived' Shuttle'derived'core'
core'vehicle'(SD0
vehicle'(SD0CV)'
CV)'with'ACES'41'
with'stretched'
Service'Module'
hypergolic'Service'
(SM)'upper'stage' Module'(SM)'upper'
stage'
Safety!
1!
2!
Launch!Reliability! Launch!Abort!
Launch!Abort!System!
System!(LAS)!
(LAS)!
!
!
Two stage to orbit Two stage to orbit
vehicle!
vehicle!
!
!
Engine!out!
Engine!out!capability!
capability!in!both!
in!second!stages!
stages!
Dream Chaser
space plane, rear
position LAS
require extra SRB,
lower reliability
Orion capsule
!
5!
5!
Highest CO2
Highest CO2!
840 tonnes!
Utilizes!greenhouse!gas!polluting!
RPO1!(Refined!Petroleum!1)!
2!
Launch!Abort!
System!(LAS)!
!
Two stage to
orbit vehicle!
!
Engine!out!
capability!in!
first!stages!
2!
Launch!Abort!
System!(LAS)!
!
Two stage to orbit
vehicle!
!
Engine!out!
capability!in!
second!stages!
!
Falcon'9'
!
Atlas'V'with'ACES'
41'Service'
Module'(SM)'
upper'stage'
'
Upper!stage!
uses!Carbon!
neutral!liquid!
hydrogen/oxy
gen!fuel!
!
3!
Minor CO2!
3!
!
!
!
Two stage to
orbit vehicle!
!
Engine!out!
capability!in!
second!stages!
!
Ares'I'
4!
Launch Abort
System (LAS)!
!
Three boosters
to orbit!
!
No!engine!out!
in!the!SRB!
!
LAS!on!the!side!
of!the!external!
tank!
3!
Minor CO2!
!
Sidemount'
Shuttle'
Capability!of!
crew!plus!40O
50!ton!
payload!
Utilizes!carbon!
neutral!liquid!
hydrogen/oxyge
n!fuel!
3!
Zero CO2!
4!
Launch Abort
System (LAS)!
!
Three boosters
to orbit!
!
No!engine!out!in!
all!stages!
!
Extra!10O20!ton!
fuel!
Dream Chaser
space plane
3!
Minor CO2
28 tonnes!
Core!booster!
uses!Carbon!
neutral!liquid!
hydrogen/oxyg
en!fuel!
!
5!
!
!
!
Three boosters
to orbit
!
No!engine!out!
in!the!SRB!
!
Man0rated'Delta' Space'Shuttle'
IV'Heavy'
'
Higher!capacity! Orion capsule
Core!booster!uses!Carbon!
neutral!liquid!hydrogen/oxygen!
fuel!
!
3!
Minor CO2!
4!
Launch!Abort!
System!(LAS)!
!
Three
boosters!
!
No!engine!out!
in!the!SRB!
!
Man0rated'
SD0HLV'
12.5.











APPENDIX
ANTENNA
Mechanical limitation
Topographic feature – protection against external RF interference
Robust - pointing accuracy
Stable foundation –pointing accuracy
Weather
o Precipitation can affect RF reception
o Cloudy skies can inhibit the use of Laser Communication Terminals
Usage of S-band
o Launch and Early Orbit Phase (LEOP)
o Emergency (safe mode)
Usage of very high data rate – frequencies (X, Ku, Ka)
o Payload data – higher transmit bandwidth
Controlled by Antenna Control Unit
Spacecraft’s ephemeris
o Describing the orbit which is computed off-line (program tracking)
Different motion of antenna
o Auto-tracking mode – full motion mono-pulse antennas
o Step tracking mode – fixed position antennas pointed to geo. Satellites
Acquisition aid antenna
o Aid to control the main dish
o Scan of a wider region of the sky
o Poor communication link, spread over larger radiated energy beam
| Page 172 of 262
12.6.
APPENDIX
INFRASTRUCTURE OF KENNEDY SPACE CENTER
12.6.1.
THE LOCATION OF THE INFRASTRUCTURE
Figure 88 – The location of infrastructure of Kennedy Space Center (NASA-COMPLEX39, 2010)
| Page 173 of 262
Length [m] Width [m] Height [m] Note
218.4
157.9
160.3 Including appurtenances
115.2
55.3
23.5 4 story concrete structure
4572.0
91.4
0.4
60.0
45.7
29.0 Spectrometer room
Logistics storage area
Launch processing system checkout area
Mechanical and electrical shops
Communications and electronics -‐ operational intercom system
2000 stations
12 local communication area 50 to 200 station per local area
112 channels
Communications and electronics -‐ operational television system
114 cameras
255 monitors
Propellant system components laboratory
propellant components lab building
propellant transporter repaor and maintenance shed
petroleum product storage building
deionized water plant
Barge terminal
Launch equipment shop
Ordnance storage facility
Storage area 1 -‐ 2 overburdened
Storage area 3 -‐ 6
Lab, shipping and reciving building
Instrumentation building
8.4
5.5
5.3 antenna tower with 18.9m
Pad water pumping station
11.9
4.9
2.8
115,000 Volt electrical substation
6 transformers
Cable terminal building
Ground Structure
Vehicle Assembly Building -‐ VAB
Launch Control Center -‐ LCC
Orbiter Landing facility
Orbiter processing/payload facility
Detail Description of the Infrastructure
Essential Facilities of Complex 39
USD 575,000
USD 1,909,000
USD 818,000
USD 200,000
USD 955,000
USD 1,500,000
USD 671,000
USD 2,400,000
USD 5,000,000
USD 8,000,000
Cost
USD 117,000,000
USD 10,000,000
USD 27,300,000
USD 9,900,000
Launch pad B
Mobile service structure – MSS
Launch Structure
Launch pad A
Cost
2
0.7km
12 pedestal resistance of 36,287,400kg
USD 21,500,000
refractory brick surface withstand
flame trench
flame deflector
environmental control system room
pad terminal connection room
high pressure gas storage facility
emergency egress dome
RP-‐1 storage
liquid oxygen storage
liquid hydrogen storage
2
0.7km
12 similar specification
USD 20,300,000
122.5 4,445,210kg
USD 11,600,000
5 platforms
2 high-‐rise elecators
1 base work elevator
1 manlift elevator
2 booms and hoists
base buildings
4 support columns
operational windspeed free standing -‐ 103.0km/h
pad position with holdiclamps -‐ 136.8km/h
park position with holddown clamps -‐ 201.2km/h
Crawler way
Mobile Launchers
Launch Structure
Transporter
40
34.7 6.1-‐7.9
2,721,550kg
load capacity 2,443,110kg
Manual or automactic operation
hydraulic system
DC power system
AC power system
auxiliary power system
51859
39.6
2.1m (Thickness)
Pedestal resistance of 36,287,400kg
USD 7,500,000
USD 33,963,000
Cost
USD 13,600,000
Title
Topic
Spaceship Captial,cost
Type
Breakdown Note
Development Engine
Tank
ReHentry
Nozzle
Head4Unit
Engine
Thrust4Chamber
Other
Material
Fuel4Tank Insulation
Connection
SubHsystem Flight4Computer
Equipmet
ReHentry Tiles
Carbon4composites
Operation,cost Maintenance
anomaly4resolution
hardware4refurbishment
Engine new4hardware4spares
flight4support
inventory4management
Fuel4Tank Overall
Critical4tiles4replacement
ReHentry
Adhesive4coating
Period
Price
Factor
First4year 13,000,000
64year4each 85,000,000
14year
5,000,000
14year
4,000,000
Price4(USD)
1 13,000,000
1 85,000,000
1 5,000,000
1 4,000,000
Manufacture
504launch 15,000,000
1
15,000,000
2004launch 5,000,000
1
5,000,000
2004launch 20,000,000
2004launch
860,000
160,000
254launch
254launch
800,000
per4launch 5,700,000
per4launch
per4launch
per4launch
2500000
3,200
2,000
1 20,000,000
1
860,000
1
160,000
1
800,000
1
5,700,000
1
1
1
2,500,000
3,200
2,000
8,205,200
Title
Topic
Infrastructure Captial/cost
Type
Breakdown
Construction Vehicle4Assembly4Building4C4VAB
Launch4Control4Center4C4LCC
Orbiter4Landing4facility
Orbiter/payload4processing4facility
Communications4and4electronics4C4operational4intercom4system
Communications4and4electronics4C4operational4television4system
Propellant4system4components4laboratory
Launch4equipment4shop
Ordnance4storage4facility
Instrumentation4building
Pad4water4pumping4station
115,0004Volt4electrical4substation
Cable4terminal4building
Launch4pad4A
Launch4pad4B
Mobile4service4structure4–4MSS
Transporter
Crawler4way
Mobile4Launchers
Operation/cost Maintenance Vehicle4Assembly4Building4C4VAB
Launch4Control4Center4C4LCC
Orbiter4Landing4facility
Orbiter/payload4processing4facility
Communications4and4electronics4C4operational4intercom4system
Communications4and4electronics4C4operational4television4system
Propellant4system4components4laboratory
Launch4equipment4shop
Ordnance4storage4facility
Instrumentation4building
Pad4water4pumping4station
115,0004Volt4electrical4substation
Cable4terminal4building
Launch4pad4A
Launch4pad4B
Mobile4service4structure4–4MSS
Transporter
Crawler4way
Mobile4Launchers
Note
Period
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
Construction4Length
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
per4year
might4not4require
per4year
per4year
Price
Factor
3 117,000,000
1 10,000,000
3 27,300,000
3
9,900,000
1
8,000,000
1
5,000,000
1
2,400,000
1
1,500,000
1
671,000
1
575,000
1
1,909,000
1
818,000
1
200,000
2 21,500,000
2 20,300,000
1 11,600,000
3 13,600,000
1
7,500,000
2 33,963,000
4,300,000
600,000
25,000,000
5,000,000
Price4(USD)
1 117,000,000
1 10,000,000
0
0
1
9,900,000
1
8,000,000
1
5,000,000
1
2,400,000
1
1,500,000
1
671,000
1
575,000
1
1,909,000
1
818,000
1
200,000
0
0
1 20,300,000
1 11,600,000
1 13,600,000
1
7,500,000
1 33,963,000
1
4,300,000
1
600,000
0
0
1
5,000,000
13,500,000
1
13,500,000
600,000
5,200,000
600,000
600,000
600,000
600,000
600,000
9,100,000
10,400,000
6,600,000
1,600,000
7,500,000
6,600,000
1
1
1
1
1
1
1
0
1
1
1
1
1
600,000
5,200,000
600,000
600,000
600,000
600,000
600,000
0
10,400,000
6,600,000
1,600,000
7,500,000
6,600,000
Title
Topic
Payload'Modules Captial'cost
Type
Breakdown
Development
Manufacture Fairing
Operation'cost Intergation
Note
Period
each4launch
Price
0
500000
1400000
Factor
Price4(USD)
1
1
500,000
1,400,000
Title
Topic
Human&Modules Captial&cost
Type
Breakdown
Development Human>Modules
Manufacture
>Capsule
SubUsystem
ReUentry
Operation&cost Maintenance
Capsule
ReUentry
Note
Quantity
Price
Factor
Price(USD)
1 550,000,000
1 550,000,000
Environmental>Control
1
Food>Preparation>Control
1
200,000,000
1
200,000,000
Waste>Management
1
Safety>System
1
Launch>Abort
1 30,000,000
1 30,000,000
Flight>Computer
2
2,000,000
1
2,000,000
Motion>and>Navigation>Control
2
500,000
1
500,000
Onboard>Measurement>System
2
2,000,000
1
2,000,000
Star>tracker
3
1,000,000
1
1,000,000
Attitude>Control>Thrusters>(ACT)>Small
16
3,000,000
1
3,000,000
ACT>Large
4
2,000,000
1
2,000,000
ACT>Fuel>
1
5,000
1
5,000
Power>(Fuel>Cell)
3
3,000,000
1
3,000,000
Communication>Antenna
2
1,000,000
1
1,000,000
Tiles
25>launch
160,000
1
760,000
Carbon>composites
25>launch
800,000
1
500,000
Overall
37,500,000
1 37,500,000
Critical>tiles>replacement
per>launch
3,200
1
3,200
Adhesive>coating
per>launch
2,000
1
2,000
Title
Topic
Type
Operation Captial,cost
Employee
(average,salary,$132000)
Operation,cost
Breakdown
Engine
Engine
Engine
Engine
Engine
Engine
Engine
Engine
Engine
Engine
Engine
Tank
Payload
Ground7Operation
Ground7Operation
Ground7Operation
Fuel
Fuel
Note
Strip7and7Build7Team
Movements7Team
Accumulation7and7Warehousing
Engineering7Team
Repair7Team
Inspectors
NDT7Inspectors
Production7Team
Materials7Team
New7Build7(replacement7parts)
Technical7Records7Team
Price/unit
132000
132000
132000
132000
132000
132000
132000
132000
132000
132000
132000
132000
132000
Mission7Operations7Facilities
n/a
Mission7Planning7&7Operations
n/a
Program7&7Doc.7Support7Management n/a
Liquid7Hydrogen7[kg]
3.82
Liquid7Oxygen7[kg]
0.15
Quantity
Price
Factor
1320000
264000
264000
396000
1320000
660000
264000
396000
660000
1320000
264000
1848000
2640000
per7year
148,400,000
per7year
74,900,000
per7year
69,400,000
63900
244098
511400
76710
10
2
2
3
10
5
2
3
5
10
2
14
20
Price7(USD)
1
1,320,000
1
264,000
1
264,000
1
396,000
1
1,320,000
1
660,000
1
264,000
1
396,000
1
660,000
1
1,320,000
1
264,000
1
1,848,000
1
2,640,000
1 148,400,000
1 74,900,000
1 69,400,000
1
244,098
1
76,710
320,808
Captial cost
Year
0
1
0
2
0
3
0
4
USD 0
USD 0
USD 85,000,000
0
5
USD 0
USD 40,960,000
USD 94,000,000
0
6
USD 106,667,600
USD 960,000
13
7
USD 106,667,600
USD 0
13
8
USD 106,667,600
USD 15,960,000
13
9
USD 106,667,600
USD 0
13
10
USD 106,667,600
USD 41,920,000
13
11
USD 213,335,200
USD 41,920,000
26
12
USD 320,002,800
USD 16,920,000
39
13
USD 320,002,800
USD 15,960,000
39
14
USD 320,002,800
USD 16,920,000
39
15
USD 320,002,800
USD 960,000
39
16
USD 320,002,800
USD 16,920,000
39
17
USD 320,002,800
USD 15,960,000
39
18
USD 320,002,800
USD 16,920,000
39
19
USD 320,002,800
USD 960,000
39
20
USD 320,002,800
USD 960,000
39
21
USD 254,361,200
USD 15,960,000
31
22
USD 213,335,200
USD 15,960,000
26
23
USD 213,335,200
USD 960,000
26
24
USD 213,335,200
USD 960,000
26
25
USD 213,335,200
USD 0
26
26
USD 147,693,600
USD 0
18
27
Vehicles Launching at year 7, 12 and 13 with 13 Human Modules each year
Launches
USD 0
USD 85,000,000
Spaceship
USD 0
USD 85,000,000
USD 0
USD 0
USD 0
USD 28,100,000
USD 0
USD 84,300,003
USD 0
USD 105,404,500
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 57,400,000
USD 0
USD 0
USD 0
USD 85,000,000
USD 0
USD 0
USD 0
USD 13,000,000
USD 0
USD 0
Development
USD 0
USD 0
Manufacture
Operation cost Maintenance
USD 0
Infrastructure
Construction
Captial cost
Development
USD 0
USD 0
USD 0
USD 0
USD 0
USD 7,128,000
USD 0
USD 0
USD 0
USD 7,128,000
USD 0
USD 0
USD 0
USD 8,976,000
USD 0
USD 0
USD 245,765,000
USD 550,000,000
USD 4,170,504
USD 304,316,000
USD 18,200,000
USD 6,500,000
USD 37,505,200
USD 245,765,000
USD 4,170,504
USD 304,316,000
USD 18,200,000
USD 6,500,000
USD 37,505,200
USD 245,765,000
USD 4,170,504
USD 304,316,000
USD 18,200,000
USD 6,500,000
USD 37,505,200
USD 245,765,000
USD 4,170,504
USD 304,316,000
USD 18,200,000
USD 6,500,000
USD 37,505,200
USD 245,765,000
USD 4,170,504
USD 304,316,000
USD 18,200,000
USD 6,500,000
USD 37,505,200
USD 245,765,000
USD 8,341,008
USD 304,316,000
USD 36,400,000
USD 13,000,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 12,511,512
USD 304,316,000
USD 54,600,000
USD 19,500,000
USD 37,505,200
USD 245,765,000
USD 9,945,048
USD 304,316,000
USD 43,400,000
USD 15,500,000
USD 37,505,200
USD 245,765,000
USD 8,341,008
USD 304,316,000
USD 36,400,000
USD 13,000,000
USD 37,505,200
USD 245,765,000
USD 8,341,008
USD 304,316,000
USD 36,400,000
USD 13,000,000
USD 37,505,200
USD 245,765,000
USD 8,341,008
USD 304,316,000
USD 36,400,000
USD 13,000,000
USD 37,505,200
USD 245,765,000
USD 8,341,008
USD 304,316,000
USD 36,400,000
USD 13,000,000
USD 37,505,200
USD 245,765,000
USD 5,774,544
USD 304,316,000
USD 25,200,000
USD 9,000,000
USD 245,765,000
Human Modules
Captial cost
Manufacture
USD 0
USD 0
USD 7,128,000
Captial cost
Manufacture
Development
Payload Modules
USD 0
USD 0
USD 7,128,000
USD 540,000,000
USD 0
USD 617,500,000 USD 1,007,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,397,500,000 USD 1,157,500,000 USD 1,007,500,000 USD 1,007,500,000 USD 1,007,500,000 USD 1,007,500,000
USD 7,128,000
Employee
Operation
Operation cost Intergation
Captial cost
Operation cost
USD 617,500,000
USD 906,094,737
USD 617,500,000
USD 90,477,592
USD 617,500,000
USD 979,328,697 USD 1,069,806,289 USD 1,161,243,881
USD 617,500,000
USD 75,477,592
USD 0
USD 888,851,105
USD 0
USD 173,347,552
USD 0
USD 0
USD 813,373,513
USD 0
USD 0
USD 344,939,488
Revenue
USD 640,025,961
Playload
USD 344,939,488
-‐USD 255,149,144
USD 295,086,473
USD 91,437,592
-‐USD 49,853,015
USD 328,979,488
USD 90,477,592
USD 329,939,488
USD 344,939,488
-‐USD 378,832,503
USD 328,979,488
USD 328,979,488
USD 329,939,488
-‐USD 708,771,991
USD 328,979,488
-‐USD 204,944,304
USD 49,517,592
-‐USD 163,024,304
-‐USD 92,128,000
-‐USD 178,984,304
-‐USD 20,128,000
-‐USD 163,024,304
-‐USD 20,128,000 -‐USD 112,256,000 -‐USD 204,384,000 -‐USD 324,612,000 -‐USD 501,040,003 -‐USD 1,546,145,503 -‐USD 1,710,129,807 -‐USD 1,873,154,111 -‐USD 2,052,138,415 -‐USD 2,215,162,719 -‐USD 2,420,107,023 -‐USD 2,370,589,431 -‐USD 2,041,609,943 -‐USD 1,711,670,455 -‐USD 1,382,690,967 -‐USD 1,037,751,479
-‐USD 163,984,304
Net Revenue
-‐USD 92,128,000 -‐USD 120,228,000 -‐USD 176,428,003 -‐USD 1,045,105,500
CF. Net Revenue
Assets
Current Assets
Fixed Assets
Current Ratio
Quick Ratio
Cash Ratio
Liabilities
Assets
Equity
Year
Cash
Account Receivable
Invertory (N/A planned)
Total
Net Fixed Assets
Total Assets
1
-‐20,128,000.00
0.00
0.00
-‐20,128,000.00
0.00
-‐20,128,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
2
-‐92,128,000.00
0.00
0.00
-‐92,128,000.00
0.00
-‐92,128,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
3
4
-‐92,128,000.00 -‐120,228,000.00
0.00
0.00
0.00
0.00
-‐92,128,000.00 -‐120,228,000.00
0.00
0.00
-‐92,128,000.00 -‐120,228,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
7
8
9
10
11
-‐163,984,304.00 -‐163,024,304.00 -‐178,984,304.00 -‐163,024,304.00 -‐204,944,304.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
-‐163,984,304.00 -‐163,024,304.00 -‐178,984,304.00 -‐163,024,304.00 -‐204,944,304.00
0.00
0.00
0.00
0.00 16,425,000.00
-‐163,984,304.00 -‐163,024,304.00 -‐178,984,304.00 -‐163,024,304.00 -‐188,519,304.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
0.00
0.00
0.00
0.00
0.00
-‐222,299,144.00
-‐222,299,144.00
0.00
0.00
0.00
0.00
0.00
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
49,517,592.00 328,979,488.00 329,939,488.00 328,979,488.00 344,939,488.00 328,979,488.00 329,939,488.00 328,979,488.00 344,939,488.00 344,939,488.00 173,347,552.00 75,477,592.00 90,477,592.00 90,477,592.00 91,437,592.00 -‐255,149,144.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
49,517,592.00 328,979,488.00 329,939,488.00 328,979,488.00 344,939,488.00 328,979,488.00 329,939,488.00 328,979,488.00 344,939,488.00 344,939,488.00 173,347,552.00 75,477,592.00 90,477,592.00 90,477,592.00 91,437,592.00 -‐255,149,144.00
32,850,000.00 49,275,000.00 49,275,000.00 49,275,000.00 49,275,000.00 49,275,000.00 49,275,000.00 49,275,000.00 49,275,000.00 49,275,000.00 49,275,000.00 32,850,000.00 32,850,000.00 32,850,000.00 32,850,000.00 32,850,000.00
82,367,592.00 378,254,488.00 379,214,488.00 378,254,488.00 394,214,488.00 378,254,488.00 379,214,488.00 378,254,488.00 394,214,488.00 394,214,488.00 222,622,552.00 108,327,592.00 123,327,592.00 123,327,592.00 124,287,592.00 -‐222,299,144.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
10,220,000.00
87%
103%
10,220,000.00
33.75
33.75
33.75
-‐1714%
-‐245%
-‐1959%
-‐51%
0%
-‐51%
-‐1908%
-‐1959%
88%
103%
10,220,000.00
32.19
32.19
32.19
-‐1634%
-‐245%
-‐1879%
-‐51%
0%
-‐51%
-‐1828%
-‐1879%
87%
103%
10,220,000.00
32.28
32.28
32.28
-‐1639%
-‐245%
-‐1884%
-‐51%
0%
-‐51%
-‐1833%
-‐1884%
87%
103%
10,220,000.00
32.19
32.19
32.19
-‐1634%
-‐245%
-‐1879%
-‐51%
0%
-‐51%
-‐1828%
-‐1879%
87%
103%
10,220,000.00
33.75
33.75
33.75
-‐1714%
-‐245%
-‐1959%
-‐51%
0%
-‐51%
-‐1908%
-‐1959%
88%
103%
10,220,000.00
33.75
33.75
33.75
-‐1714%
-‐245%
-‐1959%
-‐51%
0%
-‐51%
-‐1908%
-‐1959%
88%
103%
10,220,000.00
-‐861%
-‐245%
-‐1106%
0%
0%
0%
-‐1106%
-‐1106%
78%
100%
10,220,000.00
-‐375%
-‐163%
-‐538%
0%
0%
0%
-‐538%
-‐538%
70%
100%
0.00
-‐450%
-‐163%
-‐613%
0%
0%
0%
-‐613%
-‐613%
73%
100%
0.00
-‐450%
-‐163%
-‐613%
-‐51%
0%
-‐51%
-‐562%
-‐613%
73%
109%
0.00
24.00
25.00
90,477,592.00 90,477,592.00
47,850,000.00 32,850,000.00
15,000,000.00 -‐10,220,000.00
27,627,592.00 67,847,592.00
-‐454%
-‐163%
-‐617%
0%
0%
0%
-‐617%
-‐617%
74%
100%
10,220,000.00
26.00
91,437,592.00
33,810,000.00
11,180,000.00
46,447,592.00
1268%
-‐163%
1104%
0%
0%
0%
1104%
1104%
115%
100%
0.00
27.00
-‐255,149,144.00
-‐313,736,736.00
-‐346,586,736.00
405,174,328.00
-‐255,149,144.00
10,220,000.00
87%
103%
32.19
32.19
32.19
-‐1634%
-‐245%
-‐1879%
-‐51%
0%
-‐51%
-‐1828%
-‐1879%
10,220,000.00
87%
103%
32.28
32.28
32.28
-‐1639%
-‐245%
-‐1884%
-‐51%
0%
-‐51%
-‐1833%
-‐1884%
222,622,552.00 108,327,592.00 123,327,592.00 113,107,592.00 124,287,592.00
222,622,552.00 108,327,592.00 123,327,592.00 123,327,592.00 124,287,592.00
75,477,592.00 90,477,592.00 80,257,592.00 91,437,592.00
10,220,000.00
60%
114%
32.19
32.19
32.19
-‐1634%
-‐245%
-‐1879%
-‐51%
0%
-‐51%
-‐1828%
-‐1879%
173,347,552.00
109%
95%
4.85
4.85
4.85
-‐246%
-‐163%
-‐409%
-‐51%
0%
-‐51%
-‐358%
-‐409%
10,220,000.00
1018%
-‐82%
937%
-‐51%
0%
-‐51%
987%
937%
383,994,488.00 383,994,488.00
394,214,488.00 394,214,488.00
334,719,488.00 334,719,488.00
100%
94%
-‐20.05
-‐20.05
-‐20.05
10,220,000.00
810%
0%
810%
-‐51%
0%
-‐51%
861%
810%
368,994,488.00 368,034,488.00
379,214,488.00 378,254,488.00
319,719,488.00 318,759,488.00
100%
95%
-‐15.95
-‐15.95
-‐15.95
10,220,000.00
889%
0%
889%
-‐51%
0%
-‐51%
940%
889%
383,994,488.00 368,034,488.00
394,214,488.00 378,254,488.00
334,719,488.00 318,759,488.00
100%
94%
-‐17.51
-‐17.51
-‐17.51
10,220,000.00
810%
0%
810%
-‐51%
0%
-‐51%
861%
810%
368,994,488.00 368,034,488.00
379,214,488.00 378,254,488.00
319,719,488.00 318,759,488.00
100%
94%
-‐15.95
-‐15.95
-‐15.95
10,220,000.00
815%
0%
815%
-‐51%
0%
-‐51%
865%
815%
72,147,592.00 368,034,488.00
82,367,592.00 378,254,488.00
39,297,592.00 318,759,488.00
100%
99%
-‐16.05
-‐16.05
-‐16.05
Balance Sheet -‐ Vehicles Launching at year 7, 12 and 13 with 13 Human Modules each year
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
5
6
-‐176,428,003.00 -‐1,045,105,500.00
0.00
0.00
0.00
0.00
-‐176,428,003.00 -‐1,045,105,500.00
0.00
0.00
-‐176,428,003.00 -‐1,045,105,500.00
10,220,000.00
0.00
10,220,000.00
0.00
10,220,000.00
-‐198,739,304.00
-‐188,519,304.00
10,220,000.00
5192%
0%
5192%
-‐51%
0%
-‐51%
5243%
5192%
-‐215,164,304.00
100%
95%
-‐102.26
-‐102.26
-‐102.26
10,220,000.00
3.00
4.00
-‐92,128,000.00 -‐120,228,000.00
0.00 -‐28,100,000.00
0.00 -‐28,100,000.00
-‐92,128,000.00 -‐64,028,000.00
877%
0%
877%
-‐51%
0%
-‐51%
927%
877%
-‐189,204,304.00 -‐173,244,304.00
-‐178,984,304.00 -‐163,024,304.00
-‐189,204,304.00 -‐173,244,304.00
10,220,000.00
100%
92%
-‐17.26
-‐17.26
-‐17.26
1.00
-‐20,128,000.00
10,220,000.00
2.00
-‐92,128,000.00
-‐72,000,000.00
-‐72,000,000.00
51,872,000.00
100%
90%
597%
0%
597%
-‐51%
0%
-‐51%
648%
597%
-‐30,348,000.00 -‐102,348,000.00 -‐102,348,000.00 -‐130,448,000.00 -‐186,648,003.00 -‐1,055,325,500.00 -‐174,204,304.00 -‐173,244,304.00
-‐20,128,000.00 -‐92,128,000.00 -‐92,128,000.00 -‐120,228,000.00 -‐176,428,003.00 -‐1,045,105,500.00 -‐163,984,304.00 -‐163,024,304.00
-‐30,348,000.00 -‐102,348,000.00 -‐102,348,000.00 -‐130,448,000.00 -‐186,648,003.00 -‐1,055,325,500.00 -‐174,204,304.00 -‐173,244,304.00
Year
Cash flow from operations
Net capital spending
Change in Net Working Captial
Cash flow from assets
100%
90%
-‐11.76
-‐11.76
-‐11.76
Cash flowto creditors
100%
66%
-‐9.01
-‐9.01
-‐9.01
458%
0%
458%
-‐51%
0%
-‐51%
508%
458%
0.12
0.14
1.14
1.63
-‐1.58
3.69
0.27
0.03
0.03
1.03
27.37
22.55
0.19
0.71
0.73
3.69
0.27
0.03
0.03
1.03
27.46
22.64
0.19
0.71
0.73
3.69
0.27
0.03
0.03
1.03
27.37
22.55
0.20
0.72
0.74
3.55
0.28
0.03
0.03
1.03
28.93
24.11
0.19
0.71
0.73
3.69
0.27
0.03
0.03
1.03
27.37
22.55
0.19
0.71
0.73
3.69
0.27
0.03
0.03
1.03
27.46
22.64
0.19
0.71
0.73
3.69
0.27
0.03
0.03
1.03
27.37
22.55
0.20
0.72
0.74
3.55
0.28
0.03
0.03
1.03
28.93
24.11
0.20
0.72
0.74
3.55
0.28
0.03
0.03
1.03
28.93
24.11
0.11
0.56
5.20
0.19
0.04
0.39
9.30
0.11
0.06
0.47
8.17
0.12
0.05
0.38
0.42
8.17
0.12
0.06
0.47
8.11
0.12
-‐0.53
1.30
-‐2.43
-‐0.41
5.00
6.00
7.00
8.00
9.00
10.00
11.00
12.00
13.00
14.00
15.00
16.00
17.00
18.00
19.00
20.00
21.00
22.00
23.00
-‐176,428,003.00 -‐1,045,105,500.00 -‐163,984,304.00 -‐163,024,304.00 -‐178,984,304.00 -‐163,024,304.00 -‐204,944,304.00 49,517,592.00 328,979,488.00 329,939,488.00 328,979,488.00 344,939,488.00 328,979,488.00 329,939,488.00 328,979,488.00 344,939,488.00 344,939,488.00 173,347,552.00 75,477,592.00
-‐56,200,003.00 -‐868,677,497.00
881,121,196.00
960,000.00 -‐15,960,000.00 15,960,000.00 -‐25,495,000.00 287,311,896.00 328,736,896.00 50,235,000.00 48,315,000.00 65,235,000.00 33,315,000.00 50,235,000.00 48,315,000.00 65,235,000.00 49,275,000.00 -‐122,316,936.00 -‐65,019,960.00
-‐56,200,003.00 -‐868,677,497.00
881,121,196.00
960,000.00 -‐15,960,000.00 15,960,000.00 -‐25,495,000.00 270,886,896.00 295,886,896.00
960,000.00
-‐960,000.00 15,960,000.00 -‐15,960,000.00
960,000.00
-‐960,000.00 15,960,000.00
0.00 -‐161,371,936.00 -‐114,294,960.00
-‐64,027,997.00
692,249,494.00 -‐1,926,226,696.00 -‐164,944,304.00 -‐147,064,304.00 -‐194,944,304.00 -‐153,954,304.00 -‐508,681,200.00 -‐295,644,304.00 278,744,488.00 281,624,488.00 263,744,488.00 311,624,488.00 278,744,488.00 281,624,488.00 263,744,488.00 295,664,488.00 457,036,424.00 254,792,512.00
Ratio
Cash / Total Assets
Total Assets/Equity
-‐9.01
-‐9.01
-‐9.01
458%
0%
458%
-‐51%
0%
-‐51%
508%
458%
12.23
0.08
0.19
0.71
0.73
8.85
8.85
8.85
-‐1.97
-‐1.97
-‐1.97
-‐3.28
-‐0.31
0.01
0.08
0.09
0.08
0.09
1.09
5.64
2.42
-‐3.79
-‐0.26
-‐0.38
1.23
1.17
-‐0.05
-‐0.05
0.95
-‐21.66
-‐23.27
-‐3.45
-‐0.29
-‐0.28
1.06
-‐0.06
-‐0.06
0.94
-‐15.95
-‐15.95
-‐3.79
-‐0.26
-‐0.31
1.06
-‐0.06
-‐0.05
0.95
-‐17.51
-‐17.51
-‐3.77
-‐0.27
-‐0.28
1.06
-‐0.06
-‐0.06
0.94
-‐15.95
-‐15.95
-‐0.28
1.06
-‐0.06
-‐0.06
0.94
-‐16.05
-‐16.05
1.01
-‐0.01
-‐0.01
0.99
-‐102.26
-‐102.26
1.06
-‐0.06
-‐0.05
0.95
-‐17.26
-‐17.26
1.09
-‐0.09
-‐0.08
0.92
-‐11.76
-‐11.76
1.11
-‐0.11
-‐0.10
0.90
-‐9.01
-‐9.01
100%
0%
100%
-‐51%
0%
-‐51%
151%
100%
1.11
-‐0.11
-‐0.10
0.90
-‐9.01
-‐9.01
Current Assets
Net Fixed Assets
Total Assts
Current Liabilities
Long Term Debt
Total Liabilities
Total stockholder Equity
1.51
-‐0.51
-‐0.34
0.66
-‐1.97
-‐1.97
Assets / Liabilities
(Current Assets -‐ Invertory / Current Liablities
Cash / Current Liabilities
Working Capital
Total Liabilities and Equity
Total Liabilities and Equity Current Liabilities Accounts Patable
Notes Payable (N/A)
Total
Long-‐term debt debt
Total Liabilities
Liquid
Total Asset TurnoverSales / Total Assets
Capital Intenstion Total Assets / Sales
Leverage / Solvency Ratios Total Debt Ratio (Total Assets -‐ Total Equity) / Total Assets
can pay debt?
Debt/Equity Ratio Total Liability / Total Equity
Equity Multiplier Total Assets / Total Equity
Times Interest Earned
Earnings before interest and tax / Interest
Cash Coverage Ratio(Earnings before interest and tax + Depreciation) / Interest
Assets Utilization Ratio
Profit Margin
Return on Assets
Return on Equity
Net Income / Net Sales
Net Income / Total Assets
Net Income / Total Equity
Profitability Ratios
Tax Rate
Annual interest rate
Ratio
Net Income/Net Sales
Common sized Financial Statement
Net Sales
Cost of Good Sold
2 Depreciation Expense
1 Earnings before interest and tax
Interest paid
Taxes
Net Income
Income Statement
sale volume Turnover
Cost of Good Sold
Depreciation Expense
Earnings before interest and tax
Interest paid
Taxable income
Taxes
Net Income
$0.63 NI / $1 NS
0%
3.5%
1
0
20,128,000
0
-‐20,128,000
10220000
-‐30348000
0
-‐30348000
2
0
92,128,000
0
-‐92128000
10220000
-‐102348000
0
-‐102348000
3
4
5
0
0
0
92,128,000 120,228,000 176,428,003
0
0
0
-‐92128000 -‐120228000 -‐176428003
10220000
10220000
10220000
-‐102348000 -‐130448000 -‐186648003
0
0
0
-‐102348000 -‐130448000 -‐186648003
Income Statement -‐ Vehicles Launching at year 7, 12 and 13 with 13 Human Modules each year
100%
127%
0%
-‐27%
2%
-‐28%
0%
-‐28%
100%
126%
0%
-‐26%
2%
-‐28%
0%
-‐31%
100%
129%
0%
-‐29%
2%
-‐31%
0%
-‐28%
100%
126%
0%
-‐26%
2%
-‐28%
0%
-‐38%
100%
133%
3%
-‐36%
2%
-‐38%
0%
1%
100%
95%
3%
2%
1%
1%
0%
19%
100%
76%
4%
20%
1%
19%
0%
19%
100%
76%
4%
20%
1%
19%
0%
19%
100%
76%
4%
20%
1%
19%
0%
20%
100%
75%
4%
21%
1%
20%
0%
19%
100%
76%
4%
20%
1%
19%
0%
19%
100%
76%
4%
20%
1%
19%
0%
19%
100%
76%
4%
20%
1%
19%
0%
20%
100%
75%
4%
21%
1%
20%
0%
20%
100%
75%
4%
21%
1%
20%
0%
11%
100%
85%
4%
11%
0%
11%
0%
4%
100%
93%
3%
4%
0%
4%
0%
6%
100%
91%
3%
6%
0%
6%
0%
5%
100%
91%
3%
6%
1%
5%
0%
6%
100%
91%
3%
6%
0%
6%
0%
-‐53%
100%
147%
6%
-‐53%
0%
-‐53%
0%
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
0 617,500,000 617,500,000 617,500,000 617,500,000 617,500,000 1,007,500,000 1,397,500,000 1,397,500,000 1,397,500,000 1,397,500,000 1,397,500,000 1,397,500,000 1,397,500,000 1,397,500,000 1,397,500,000 1,157,500,000 1,007,500,000 1,007,500,000 1,007,500,000 1,007,500,000 540,000,000
########### 781,484,304 780,524,304 796,484,304 780,524,304 822,444,304 957,982,408 1,068,520,512 1,067,560,512 1,068,520,512 1,052,560,512 1,068,520,512 1,067,560,512 1,068,520,512 1,052,560,512 1,052,560,512
984,152,448
932,022,408
917,022,408
917,022,408
916,062,408 795,149,144
0
0
0
0
0
16425000
32850000
49275000
49275000
49275000
49275000
49275000
49275000
49275000
49275000
49275000
49275000
32850000
32850000
32850000
32850000
32850000
-‐1045105500 -‐163984304 -‐163024304 -‐178984304 -‐163024304 -‐221369304
16667592
279704488
280664488
279704488
295664488
279704488
280664488
279704488
295664488
295664488
124072552
42627592
57627592
57627592
58587592 -‐287999144
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
10220000
0
0
0
10220000
0
0
-‐1055325500 -‐174204304 -‐173244304 -‐189204304 -‐173244304 -‐231589304
6447592
269484488
270444488
269484488
285444488
269484488
270444488
269484488
285444488
285444488
124072552
42627592
57627592
47407592
58587592 -‐287999144
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
-‐1055325500 -‐174204304 -‐173244304 -‐189204304 -‐173244304 -‐231589304
6447592
269484488
270444488
269484488
285444488
269484488
270444488
269484488
285444488
285444488
124072552
42627592
57627592
47407592
58587592 -‐287999144
-‐28%
12.7.5.
APPENDIX
30 TONNES MECHANICAL PAYLOAD
Revenue against time for 30 tonnes payload with 2
vehicles
USD 500,000,000
Revenue
USD 0
-USD 500,000,000
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27
Net
Revenue
-USD 1,000,000,000
-USD 1,500,000,000
Year of Project
Figure 89 – Vehicles delivering with mechanical payload only
servicing at year 7 and 12 with price of $1000/kg with fairing
approximate net revenue of - $661m
vehicle
USD 1,000,000,000
USD 500,000,000
Revenue
USD 0
-USD 500,000,000
1
2
3
4
5
6
7
8
9 10 11 12 13 14 15 16 17 18 19 20 21 22
Net
Revenue
-USD 1,000,000,000
Year of Project
servicing at year 7 with price of $1500/kg with fairing
approximate net revenue of $325m
vehicles
USD 4,000,000,000
Revenue
USD 3,000,000,000
USD 2,000,000,000
Net
Revenue
USD 1,000,000,000
USD 0
-USD 1,000,000,000
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28
-USD 2,000,000,000
Year of Project
servicing at year 7, 12 and 13 with price of $1000/kg with fairing
approximate net revenue of $2.80b
| Page 160 of 262
12.7.6.
APPENDIX
HUMAN MODULES
Revenue against time for 20 tonnes payload and 7
tonnes by Human Modules for with 1 vehicle
USD 1,000,000,000
Revenue
USD 0
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22
Net
Revenue
-USD 1,000,000,000
-USD 2,000,000,000
Year of Project
Figure 92 – Vehicles delivering with the aid human capsule of 7 tonnes
servicing at year 7 with price of $1000/kg with fairing and $7000/kg with capsule
approximate net revenue is $733m
Revenue against time for 20 tonnes payload and 7
tonnes by Human Modules for with 3 vehicles
1.6E+09
1.4E+09
1.2E+09
1E+09
800000000
600000000
400000000
200000000
0
Revenue
Net
Revenue
1 2 3 4 5 6 7 8 9 10 11 12 Year
13 14of15Project
16 17 18 19 20 21 22 23 24 25 26 27
Figure 93 – Vehicles delivering with the aid human capsule of 7 tonnes
servicing at year 7, 12 and 13 with price of $1000/kg with fairing and $3500/kg with capsule
approximate net revenue is $642m
| Page 161 of 262
12.7.7.
APPENDIX
SENSITIVITY ANALYSIS
Project Sensitivity to Reusability of
engine
USD 2,000,000,000
Net Revenue
USD 0
-USD 2,000,000,000
0
10
20
30
40
50
60
-USD 4,000,000,000
-USD 6,000,000,000
-USD 8,000,000,000
-USD 10,000,000,000
Reusability of engine
Figure 94 - Impact to the net revenue of engine
approximate $642m with 50 time of reuse
approximate - $8.53b with new engine every launch
Net Revenue
Project Sensitivity to Reusability of fuel
tank
USD 650,000,000
USD 640,000,000
USD 630,000,000
USD 620,000,000
USD 610,000,000
USD 600,000,000
USD 590,000,000
USD 580,000,000
USD 570,000,000
0%
20%
40%
60%
80%
Percentage reusability of fuel tank
100%
120%
Figure 95 - Impact to the net revenue of fuel tank
| Page 162 of 262
APPENDIX
Net Revenue
Project Sensitivity to Re-entry
systems
USD 645,000,000
USD 640,000,000
USD 635,000,000
USD 630,000,000
USD 625,000,000
USD 620,000,000
USD 615,000,000
USD 610,000,000
USD 605,000,000
0%
20%
40%
60%
80%
100%
Percentage Reuability of re-entry systems
120%
Figure 96 - Impact to the net revenue of re-entry systems
Project Sensitivity to fuel price
USD 900,000,000
USD 800,000,000
Net Revenue
USD 700,000,000
USD 600,000,000
USD 500,000,000
USD 400,000,000
USD 300,000,000
USD 200,000,000
USD 100,000,000
USD 0
-100% -80% -60% -40% -20% 0% 20% 40% 60% 80% 100% 120%
Percentage change of fuel price
Figure 97 - Impact to the net revenue from price of fuel
approximate $642m with price base from NASA
approximate $787m with 75% less
approximate $450m with 100%
| Page 163 of 262
APPENDIX
Net Revenue
Project Sensitivity to the cost of
Infrastructure
$ 1,600,000,000
$ 1,400,000,000
$ 1,200,000,000
$ 1,000,000,000
$ 800,000,000
$ 600,000,000
$ 400,000,000
$ 200,000,000
$0
0%
20%
40%
60%
80%
100%
Infrstrructure construction cost percentage change
120%
Figure 98 – Impact to the net revenue from infrastructure construction cost
approximate $642m with price base from NASA
approximate $1.38b with 50% less
Project Sensitivity to Human
module
USD 700,000,000
Net Revenue
USD 600,000,000
USD 500,000,000
USD 400,000,000
USD 300,000,000
USD 200,000,000
USD 100,000,000
USD 0
0%
20%
40%
60%
80%
100%
Percentage reusability of human module
120%
Figure 99 Impact to the net revenue from human module
approximate $642m with 13 launches
approximate $192m manufacture every launch
| Page 164 of 262
APPENDIX
Project Sensitivity to Discount Rate
Net Present Value (NPV)
$ 100,000,000
$0
-$ 100,000,0000.0%
2.0%
4.0%
6.0%
8.0%
10.0% 12.0% 14.0%
-$ 200,000,000
-$ 300,000,000
-$ 400,000,000
-$ 500,000,000
-$ 600,000,000
-$ 700,000,000
-$ 800,000,000
-$ 900,000,000
Discount Rate
Figure 100 – Impact to the NPV by the discount rate
approximate - $776m with discount rate of 12.5%
approximate $10.6m with discount rate of 3%
| Page 165 of 262
12.8.
APPENDIX
RISK ASSESSMENT
12.8.1.
QUANTIFY RISKS
Table 28 – Typical risk severity categories
Score
Technical
1
2
3
4
5
Health & Safety
Impact on operational
performance
negligible
Minor impact on
operational
performance
Noticeable impact on
operational
performance
Substantial operation
failure
Catastrophic failure
Environmental
Minor injury/ inconvenience
Short term local
damage
Minor injury
Medium term local
damage Short term
regional damage
Reportable injury
Long term local damage
Regional damage
Major injury / illness Long
term effects
Fatalities
Long term widespread
damage
Widespread permanent
damage
Table 29 – Typical risk likelihood categories
Score
Descriptor
Description
1 Improbable
About 1 in 1000
2 Remote
About 1 in 100
3 Occasional
About 1 in 10
4 Probable
Likely to happen
5 Frequent
Expect to
happen
| Page 166 of 262
APPENDIX
Table 30 – Risk rating and associated action
Likelihood Score
Severity
Score
1
2
3
4
5
1
Low
Low
Low
Medium
Medium
2
Low
Low
Medium
Medium
High
3
Low
Medium
Medium
High
High
4
Medium
Medium
High
High
High
5
Medium
High
High
High
High
Risk Level
Low
Medium
High
Action Required
Check that risks cannot be reduced further or eliminated
Consider alternative approach or list residual hazards and specify
precautions
Seek alternative solutions and list residual hazards and
precautions.
| Page 167 of 262
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12.1.
APPENDIX
MASS CONSIDERATIONS/DIMENSIONS AND CALCULATIONS
12.1.1.
SATELLITE DIMENSIONS AND MASSES
Table 31: Name and specifications of some common Satellites
Name
Payload
Type
EROS
Earth
Observation
0.7
1.2 (cylindrical
head)
Anik F1
Broadcasting
2.1
3.4
4
4710 35786
Astra 1N
Broadcasting
2.8
6.5
3.2
5300 35786
CBERS
Earth
Observation
1.8
2
2.2
1450
778
ChandrayaanResearch
1
1.5
1.5
1.5
1380
Moon
Orbit
Chang'e 1
Research
2
1.72
2.2
2350
Moon
Orbit
Cube Sat
Science and
exploration
0.1
0.3
0.1
1.3
DirectTV
Broadcasting
2.8
3.3
3.8
1727 35786
EnviSat
Earth
Observation
4
4
10
8211
Galileo
Navigation
1.58
3.02
2.7
IKONOS
Earth
Observation
1.83
1.57
3.04
IntelSat
Broadcasting
3.3
3.2
6.9
Jason-2
Earth
Observation
1
1.3
3.7
Landsat 7
Earth
Observation
2.8
4.3
OSO-1
Research
Width/Diameter Length
(m)
(m)
2.18 cylinder
Height Weight Orbit
(m)
(kg)
(km)
2.3
3
250
480
788
700 23222
720
681
6000 35786
525
1336
2200
705
1042
555
| Page 169 of 262
APPENDIX
QuickBird
Earth
Observation
1.83
1.57
3.04
950
482
RapidEye
Earth
Observation
0.78
1.17
0.938
175
630
SAC-D
Earth
Observation
2.7
2.7
5
657
1350
SELENE
Research
2.1
2.1
4.8
2914
Moon
Orbit
SMART-1
Research
1
1
1
367
Moon
Orbit
Spacebus
4000C4
Broadcasting
2.2
2
5.5
5900
SPOT 6
Earth
Observation
1.55
1.75
2.7
712
694
Terra
Earth
Observation
2.4
3.3
6.8
4864
712
TerraSAR-X
Earth
Observation
2.4 cylinder
5
1230
514
12.2.
HUMAN PAYLOAD
12.2.1.
THIRTY PERSON HUMAN ADAPTATION OF SHUTTLE CARGO
For a cylindrical vessel of the general dimensions of 20metre length and 6 metre diameter
𝐶𝑎𝑟𝑏𝑜𝑛 − 𝐴𝑙𝑢𝑚𝑖𝑛𝑖𝑢𝑚 𝑐𝑜𝑚𝑝𝑜𝑠𝑖𝑡𝑒 𝑌𝑜𝑢𝑛𝑔 𝑠 𝑀𝑜𝑑𝑢𝑙𝑢𝑠 𝐸 = 230𝐺P𝑎 𝑈𝑙𝑡𝑖𝑚𝑎𝑡𝑒 𝑆𝑡𝑟𝑒𝑛𝑔𝑡ℎ, 𝜎 = 0.9𝐺𝑃𝑎 𝐷𝑒𝑛𝑠𝑖𝑡𝑦, 𝜌 =
𝐹𝑜𝑟 𝑡ℎ𝑖𝑛𝑛𝑒𝑑 𝑤𝑎𝑙𝑙𝑒𝑑 𝑝𝑟𝑒𝑠𝑠𝑢𝑟𝑒 𝑣𝑒𝑠𝑠𝑒𝑙𝑠
𝑓𝑜𝑟 𝑠𝑎𝑓𝑒𝑡𝑦 𝑡𝑎𝑘𝑒 𝜎 =
2250𝑘𝑔
𝑚
𝑃𝑟
= 𝜎 𝑤ℎ𝑒𝑟𝑒 𝑃 = 101𝑘𝑃𝑎 (𝑎𝑡𝑚𝑜𝑠𝑝ℎ𝑒𝑟𝑖𝑐)
𝑡
𝜎
(2𝑃𝑟)
∴ 𝑟𝑒𝑎𝑟𝑟𝑎n𝑔𝑖𝑛𝑔 𝑓𝑜𝑟 𝑡 =
2
𝜎
| Page 170 of 262
𝑡=
APPENDIX
(2 × 101 × 10 × 3)
= 0.67𝑚𝑚 0.9 × 10
∴ 𝑚𝑎𝑠𝑠 𝑚𝑎𝑦 𝑏𝑒 𝑓𝑜𝑢𝑛𝑑 𝑓𝑟𝑜𝑚 𝑡ℎ𝑒 𝑑𝑒𝑠𝑖𝑡𝑦 𝑚𝑎𝑠𝑠 = 𝜌 × 𝑣 𝑣 = 𝜋𝑟
− 𝜋𝑟
ℎ 𝑣 = (𝜋3 − 𝜋2.99 ) × 20 = 3.76𝑚 ∴ 𝑚𝑎𝑠𝑠 = 2250 × 3.76 = 8468𝑘𝑔
12.2.2.
CAPSULE MASS AND CALCULATIONS
Figure 101 Approximate shape of Capsule (truncated cone) (CASIO COMPUTER CO., LTD, 2013)
𝑇𝑜 𝑒𝑠𝑡𝑖𝑚𝑎𝑡𝑒 𝑡ℎ𝑒 𝑝𝑜𝑡𝑒𝑛𝑡𝑖𝑎𝑙 𝑚𝑎𝑠𝑠 𝑜𝑓 𝑡ℎ𝑒 𝑐𝑎𝑝𝑠𝑢𝑙𝑒 𝑢𝑠𝑖𝑛𝑔 𝑡ℎ𝑒 𝑣𝑜𝑙𝑢𝑚𝑒 𝑒𝑞𝑢𝑎𝑡𝑖𝑜𝑛 𝑓𝑜𝑟 𝑎 𝑡𝑟𝑢𝑛𝑐𝑎𝑡𝑒𝑑 𝑐𝑜𝑛𝑒 1
𝑉 = 𝜋(𝑟
3
+ 𝑟 𝑟 + 𝑟 )ℎ 𝑡ℎ𝑒𝑛 𝑢𝑠𝑖𝑛𝑔 ℎ = 6𝑚, 𝑟 = 2𝑚 𝑎𝑛𝑑 𝑟 = 1𝑚 𝑉
1
= 𝜋 × (1 + (1 × 2) + 2 ) × 6 = 43.98𝑚
3
𝑇𝑜 o𝑏𝑡𝑎𝑖𝑛 𝑎 ℎ𝑜𝑙𝑙𝑜𝑤 𝑠𝑡𝑟𝑢𝑐𝑡𝑢𝑟𝑒 𝑎𝑠𝑠𝑢𝑚𝑒 𝑡ℎ𝑎𝑡 𝑒𝑙𝑖𝑚𝑖𝑛𝑎𝑡𝑖𝑛𝑔 𝑡ℎ𝑒 𝑖𝑛𝑛𝑒𝑟 𝑐𝑦𝑐𝑙𝑖𝑛𝑑𝑒𝑟 𝑐𝑟𝑒𝑎𝑡𝑒𝑑 𝑏𝑦 𝑡ℎ𝑒 𝑟 𝑑𝑖𝑚𝑒𝑛𝑠𝑖𝑜𝑛
𝑤𝑜𝑢𝑙𝑑 𝑠𝑢𝑓𝑓𝑖𝑐𝑒 ∴ 𝑣𝑜𝑙𝑢𝑚𝑒 𝑜𝑓 𝑐𝑦𝑙𝑖𝑛𝑑𝑒𝑟 = 𝜋𝑟 ℎ = 𝜋 × 1 × 6
= 18.85𝑚 𝑟𝑒𝑠𝑢𝑙𝑡𝑖𝑛𝑔 𝑖𝑛 𝑓𝑖𝑛𝑎𝑙 𝑜𝑣𝑒𝑟𝑎𝑙𝑙 𝑣𝑜𝑙𝑢𝑚𝑒 𝑜𝑓 𝑉
= 25.1𝑚 𝑇𝑜 𝑒𝑠𝑡𝑖𝑚𝑎𝑡𝑒 𝑡ℎ𝑒 𝑚𝑎𝑠𝑠 𝑜𝑓 𝑡ℎ𝑒 𝑠𝑡𝑟𝑢𝑐𝑡𝑢𝑟𝑒 𝑚𝑎𝑠𝑠 = 𝜌 × 𝑣 𝑤ℎ𝑒𝑟𝑒 𝜌 = 2700
𝑘𝑔
𝑚
𝑚𝑎𝑠𝑠 𝑖𝑓 𝑚𝑎𝑑𝑒 𝑤𝑖𝑡ℎ 𝑠𝑜𝑙𝑖𝑑 𝑎𝑙𝑢𝑚𝑖𝑛𝑖𝑢𝑚 𝑚 = 2700 × 25.1 = 67770𝑘𝑔 𝐻𝑜𝑤𝑒𝑣𝑒𝑟 𝑖𝑛 𝑝𝑟𝑎𝑐𝑡𝑖𝑐𝑒 𝑜𝑛𝑙𝑦 𝑡ℎ𝑒 𝑝𝑟𝑒𝑠𝑠𝑢𝑟𝑒 𝑜𝑢𝑡𝑡𝑒𝑟 𝑖𝑛 𝑖𝑛𝑛𝑒𝑟 𝑙𝑎𝑦𝑒𝑟 𝑤𝑖𝑙𝑙 𝑏𝑒 𝑚𝑎𝑑𝑒 𝑜𝑓 𝑡ℎ𝑒 𝑝𝑢𝑟𝑒 𝑚𝑎𝑡𝑒𝑟𝑖𝑎𝑙 𝑎𝑛𝑑 𝑡ℎ𝑒
| Page 171 of 262
APPENDIX
𝑖𝑛𝑡𝑒𝑟𝑖𝑜𝑟 𝑤𝑜𝑢𝑙𝑑 𝑏𝑒 𝑚𝑎𝑑𝑒 𝑜𝑓 ℎ𝑜𝑛𝑒𝑦 𝑐𝑜𝑚𝑏 𝑠𝑡𝑟𝑢𝑐𝑡𝑢𝑟𝑒 𝑤𝑖𝑡ℎ 𝑑𝑒𝑛𝑠𝑖𝑡𝑦 𝜌
= 83
𝑘𝑔
𝑎𝑛𝑑 𝑡ℎ𝑢𝑠 𝑡ℎ𝑒 𝑚𝑎𝑠𝑠 𝑤𝑜𝑢𝑙𝑑 𝑏𝑒 𝑐𝑙𝑜𝑠𝑒𝑟 𝑡𝑜 𝑚 = 83 × 25.1
𝑚
= 2083.3𝑘𝑔 (𝑚𝑢𝑙𝑡𝑖𝑝𝑙𝑦 𝑡ℎ𝑖𝑠 𝑏𝑦 𝑡𝑤𝑜 𝑎𝑠 𝑖𝑡 𝑖𝑠 𝑜𝑛𝑙𝑦 5𝑐𝑚 𝑡ℎ𝑖𝑐𝑘 𝑎𝑛𝑑
∴ 𝑛𝑒𝑒𝑑 𝑡𝑤𝑜 𝑙𝑎y𝑒𝑟𝑠)(4167𝑘𝑔) + (1200) = 5367𝑘𝑔
( 𝐶𝑎𝑛 𝑎𝑝𝑝𝑟𝑜𝑥𝑖𝑚𝑎𝑡𝑒 𝑎 𝑠ℎ𝑒𝑙𝑙 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑜𝑓 𝑎 𝑓𝑒𝑤 𝑚𝑖𝑙𝑙𝑖𝑚𝑒𝑡𝑟𝑒𝑠 𝑀𝑎𝑠𝑠 𝑎𝑙𝑙𝑜𝑐𝑎𝑡𝑒𝑑 𝑡𝑜 𝑜𝑢𝑡𝑒𝑟 𝑤𝑎𝑙𝑙 𝑠ℎ𝑒𝑙𝑙 𝑠𝑡𝑟𝑢𝑐𝑡𝑢𝑟𝑒 𝑤𝑖𝑙𝑙 𝑏𝑒 𝑖𝑛 𝑡ℎ𝑒 𝑟𝑎𝑛𝑔𝑒 𝑜𝑓 𝑎 𝑓𝑒𝑤 𝑡ℎ𝑜𝑢𝑠𝑎𝑛𝑑 𝑘𝑔 Figure 102: Slant height for a truncated cone. (GORDILLO, J.C.F, 2010)
𝑤ℎ𝑒𝑟𝑒 ℎ = 6𝑚 𝑅(𝑟 ) = 2𝑎𝑛𝑑 𝑟(𝑟 ) = 1
𝑠=
6 + 1 = 6.1𝑚
| Page 172 of 262
12.2.1.
APPENDIX
MASS OF HUMAN MODULE AND LIFE-SUPPORT
Table 32Table 4 Mass of 30 person system
Component
mass/kg
Passengers
2370
Passenger seats
6390
98.6
Lithium Hydroxide Canisters
45
0.006 (0.09)
Trace contaminant Control Subsystem
78.2
Fans/ air circulation sys
9.6
Oxygen
122.25
Individual oxygen canister
TBA
Nitrogen
366.75
Waste management commode
50
High
efficiency
particle
atmosphere
volume/m^3
1.00E-02
filter
(HEPA)
2
Sensors
negligible
Smoke Detector
1.5
Portable Fire extinguisher PFE
15.1
Water Storage unit (Fuel cells)
300
Food lockers (full)
54.9
0.112
storage lockers
378
8.494
Food heating system
3475
Individual Monitors
100
Defibrillator system
3.2
medications and bandage kit
negligible
emergency medical kit
negligible
Space suits
1200
0.01
5.54E-03
| Page 173 of 262
APPENDIX
Vessel mass
Aluminium(with rad shielding)
6768(8468)
251.3
21729.5
Total mass
(23429)
if CFRP used for vessel
(or cheaper light
material)~
15722.64
Table 33 Mass breakdown for 6 passengers (8day mission)
component
mass/kg
Passengers
480
Passenger seats
120
98.6
Lithium Hydroxide Canisters
72
0.006 (0.09)
Trace contaminant Control Subsystem
78.2
Fans/ air circulation sys
4.8
Oxygen
195.6
Individual oxygen canister
TBA
Nitrogen
~586.8
Waste management commode
50
High efficiency particle atmosphere filter (HEPA)
2
Sensors
Negligible
Smoke Detector
1.5
Portable Fire extinguisher PFE
15.1
Water Storage unit (Fuel cells)
300
Food lockers (full)
512.4
0.112
storage lockers
2052
8.494
Food heating system
3475
Defibrillator system
3.2
medications and bandage kit
Negligible
volume/m^3
1.00E-02
0.01
5.54E-03
| Page 174 of 262
APPENDIX
emergency medical kit
Negligible
Space suits
60
Vessel mass
Aluminium (with Radiation shielding)
5367
Telemetry
TBA
Total mass
8774.8 (12788)
if CFRP used for vessel
material)~
251.3
(or cheaper light
8182.94(12197)
Table 34Mass breakdown for six person mission using regenerative Life support systems on ISS (SALOTTI,
Jean Marc, 2006)
Mass/ Volume/
Average
System/item
kg
m3
Power/kW
Air Revitalisation System(ARS)
Carbon Dioxide Removal Assembly (CDRA)
201
0.39
0.86
Trace contaminant Control System (TCCS)
78.2
0.175
Major Constituent Analyzer (MCA)
54.7
0.44
0.0088
Oxygen Generation Assembly (OGA)
113
0.14
1.47
Temperature and Humidity Control System(THCS)
Common Cabin Air Assembly (CCAA)
112
0.4
0.468
Avionics Air Assemby (AAA)
12.4
0.03
0.083
Intermodule Ventilation (IMV) Fan
4.8
0.01
0.055
Intermodule Ventilation (IMV) Valve
5.1
0.01
0.006
High Efficiency Particle Atmosphere (HEPA) Filter
2
0.01
Fire Detection and Suppression
Smoke Detector
1.5
0.002
Portable Fire Extinguisher (PFE)
15.1
0.04
Crew Cabin
Volume: 50 m³/person
300
Crew Cabin
Passengers (6 @average mass of 80kg)
480
Water Recovery and Management (WRM) and Waste
Management (WM)
Water Processor (WP)
476
10.39
0.3
Process Control Water Quality Monitor (PCWQM)
38
0.51
0.03
| Page 175 of 262
Urine Processor (UP)
Fuel Cell Water Storage
Condensate Storage
Commode / Urinal
Miscellaneous
Food
Food heating system
Defibrulator system
APPENDIX
128
21
21
50
Spacesuits
Storage lockers
Refrigurated storage
Total Mass (life support only)
0.37
0.1
0.1
0.091
0.072
40
3475
3.2
5.54E-03
320
2052
4.2475
7704
In order to find the volume of Oxygen this corresponded to the following equation was used:
Oxygen gas obeys the gas law
PV=nRT 𝑃𝑉 = 𝑛𝑅𝑇, 𝑤ℎ𝑒𝑟𝑒 𝑃 𝑖𝑠 𝑝𝑟𝑒𝑠𝑠𝑢𝑟𝑒 𝑃 = 101𝑘𝑃𝑎 , 𝑉 𝑖𝑠 𝑡ℎ𝑒 𝑣𝑜𝑙𝑢𝑚𝑒 𝑜𝑓 𝑔𝑎𝑠 𝑢𝑛𝑘𝑛𝑜𝑤𝑛, 𝑛 𝑖𝑠 𝑡ℎ𝑒 𝑛𝑢𝑚𝑏𝑒𝑟 𝑜𝑓 𝑚𝑜𝑙𝑒𝑠 𝑜𝑓 𝑔𝑎𝑠, 𝑅 𝑖𝑠 𝑡ℎ𝑒 𝑚𝑜𝑙𝑎𝑟 𝑔𝑎𝑠 𝑐𝑜𝑛𝑠𝑡𝑎𝑛𝑡8.3145 (𝐽
/𝑚𝑜𝑙𝑒 𝑥 𝐾) 𝑎𝑛𝑑 𝑇 𝑖𝑠 𝑡ℎ𝑒 𝑟𝑜𝑜𝑚 𝑡𝑒𝑚𝑝𝑒𝑟𝑎𝑡𝑢𝑟𝑒 𝑇 = 294𝐾. 𝑇ℎ𝑢𝑠 𝑓𝑜𝑟 40.32𝑘𝑔 𝑡ℎ𝑒𝑟𝑒 𝑎𝑟𝑒
(40.32 × 10 )
𝑚𝑜𝑙𝑒𝑠
32
= 1260𝑚𝑜𝑙𝑒𝑠 𝑇ℎ𝑒𝑛 𝑟𝑒𝑎𝑟𝑟𝑎𝑛𝑔𝑖𝑛𝑔 1 𝑔𝑖𝑣𝑒𝑠 𝑉 =
=
𝑛𝑅𝑇
𝑠𝑢𝑏𝑠𝑡𝑖𝑡𝑢𝑡𝑖𝑛𝑔 𝑉
𝑃
(1260 × 8.3145 × 294)
= 30.5𝑚
(101 × 10 )
| Page 176 of 262
12.3.
12.3.1.
APPENDIX
FURTHER PROPELLANT INFORMATION.
MONOPROPELLANTS
Monopropellants derive their heat, typically, by catalytic chemical decomposition. These fuels
have historically been toxic and chemically unstable, making them dangerous both in handling
and in operation. These properties are abundantly displayed by examples such as hydrazine and
methyl nitrate. ISP values between 200 and 220 seconds are achieved, making these barely
efficient enough to be used in a launcher of sizeable dry-mass fraction.
A
safer
and
higher-performance
family
of
monopropellants,
known
as
HAN
(Hydroxylammonium Nitrate), have been proposed with theoretical ISPs of 270s (JANKOVSKY,
Robet S, 1996, p.3). However, even with peak efficiency, this specific impulse is not great enough
for a launcher requiring structural safety factors for re-usability.
Another monopropellant family could include high explosives. This method of propulsion was
tested as a part of project Orion (1958). However, high explosives provide very low specific
impulses, in the order of 2s-2.5s (KESHAVARZ, Mohammad Hossein, 2008, p.363) meaning that
they would deliver very little delta-V for a given mass.
Monopropellants do offer advantages – most notably, monopropellant systems are simple and
light-weight. This factor makes them appealing for on-orbit reaction control systems. However,
they are not widely available in most cases and must be produced by dedicated chemical plants,
driving up their costs.
On this basis, monopropellants will receive no further consideration for the launcher. This is
primarily due to technical infeasibility, environmental impact, safety hazards and cost.
12.3.2.
LIQUID BI-PROPELLANTS: OXIDISERS
Oxidisers containing Fluorine typically produce the highest ISP, in the region of typically 420s on
average. However, Fluorine is highly toxic and flammable. The products of combustion with
Fluorine are also toxic, corrosive and damaging to the environment. Fluorine in its own right is
corrosive, but containers can be constructed from conventional materials, such as aluminium or
steel. However, fluorine boils at 84K, meaning it must be stored cryogenically and allowed to boil
off, making the design more challenging.
The next highest ISP is achieved by liquid oxygen, in the range of 380s. Oxygen has the
advantages over Fluorine that it is non-toxic and non flammable. The products it forms are also in
general non toxic. Liquid oxygen must also be stored cryogenically at less than 90K, leading to
frosting problems and adversely affecting the fracture properties of materials in contact with it.
Typically, oxygen is benign to handle and store, having good corrosion properties.
Author: James Roper
| Page 177 of 262
APPENDIX
Other oxidisers include chlorine tri-fluoride, tetra-fluoric-hydrazine and other Halogen
compounds. Also, oxides of nitrogen and strong acids are used, such as red-fuming nitric acid. All
of these are very highly toxic, some are mutagenic or carcinogenic. Many are also hypergolic,
meaning they ignite on contact, which is less than desirable if they are spilled. Due to their
corrosive and toxic nature, they are expensive to produce, handle and transport. In addition to this,
these toxic oxidisers offer reduced specific impulses averaging 320s (ideal), making them less
attractive options.
Of the oxidisers mentioned above, it seems likely that liquid oxygen is the most attractive. It is
generally benign, widely available and relatively inexpensive. A NASA document on the Space
Shuttle fuels indicated the price to be $0.155/kg in 2001 (NASA, 2001, p.1). Oxygen is also dense
enough (1142kg/m^3) that the tanks required will not be excessively large.
The main drawback of oxygen is that it is cryogenic, meaning it cannot be stored for long periods
due to boil-off and it can cause frosting. The reduced ISP compared to Fluorine and Oxygen Difluoride can be tolerated by selecting an appropriate fuel, possibly with additives.
12.3.3.
LIQUID BI-PROPELLANTS: EXPANSION ON RP-1, LH2 AND LNG
Numerous fuels are available. As with oxidisers, the major factors that will distinguish between
them are specific impulse, toxicity and storability. Typically, it will be necessary that the reaction
products have low molecular weights and produce high temperatures. Thus, the fuel itself must
contain low atomic mass elements and produce strongly-bonded products (to release high
temperatures).The following fuels are considered on the basis that liquid oxygen is the oxidiser.
Hydrocarbons such as methane or other petroleum gasses, RP-1, etc are known to produce high
combustion temperatures (~3350C on average). However, their specific impulses are limited due
to the mass of their combustion products (<370s) (HUZEL, Dieter K and Huang, David H, 1992,
p.20). Depending on the constituent, hydrocarbons may have densities ranging from 400820kg/m^3 and can be stored at room temperature and pressure, making them attractive for design
simplicity.
Hydrocarbons are widely available and inexpensive. The combustion products are steam and
carbon dioxide when oxidised using liquid oxygen, both of which are not toxic but do contribute
to the green-house effect. Handling of hydrocarbons presents dangers, as they are typically
flammable, but these are well understood. The current environmental and economic trend
indicates that hydrocarbons will be replaced in the coming decades, as they become scarcer.
It is possible that hydrocarbons could be replaced by naturally produced environmentally neutral
gasses. LNG (liquefied natural gas) is chiefly methane, offering a specific impulse (ideal, in
vacuum) of 369s. It must be stored cryogenically at similar temperatures and pressures to oxygen
Author: James Roper
| Page 178 of 262
APPENDIX
– thus, if liquid oxygen is used, there will be a design and production saving as both fuel vessels
could be similar (or identical).
On this basis, hydrocarbons will be further considered due to their low cost and design
advantages.
Hydrogen has obvious attractions due to its low molecular mass and consequently offers the
highest specific impulses (455s). The major drawbacks of hydrogen are its low density, only
67.8kg/m^3, and deep cryogenic requirements (<33K). As a result, hydrogen will incur higher
developmental costs than hydrocarbons in order to design for reusability. It has however been
used with success in the past – the space shuttle external main tank and engines being the obvious
example.
One benefit of hydrogen is that the low molecular mass of its combustion products reduces the
temperature required to produce a high exhaust velocity – as a result, the high thermal stresses
within the engine are reduced. The temperature is reduced by approximately 12% relative to
hydrocarbons, for a specific impulse increase of 27% (HUZEL, Dieter K and Huang, David H,
1992, p.20).
A notable drawback to hydrogen is hydrogen embrittlement. This occurs when hydrogen becomes
dissolved in the metal containers that encase it. The dissolved hydrogen re-combines within the
metal, causing porosity that may result in fracture. Designing against this has been previously
achieved by NASA for the external tank (ET) of the space shuttle, by the use of heavy transition
metal coatings (HARRIS, Yolanda, 2010, p.210).
12.3.4.
USE OF METALS AS FUELS
Another fuel type to be considered includes metal powders. Aluminium powder is typically added
to solid rocket fuels to increase the temperature of combustion. A fluidised powder could be used
as a fuel in its own right, or in conjunction with a flammable carrier gas. Studies using metal
powders and carbon dioxide oxidiser have shown a specific impulse of 280s may be achieved
(FOOTE, J.P. & Litchford, R.J., 2007, p.16). If oxygen were used, this could be higher.
Significant advantages in fuel density and safety would thus be achieved, as metals are very dense
and solid. However, the machinery used to pump the powdered fuel would be quickly eroded due
to the abrasive nature of the powder and solid slag produced by combustion would block injectors
and accumulate in the engine nozzle, making reliability and reusability very poor.
Beryllium is another metal that is proposed for use as a fuel, as an additive to hydrogen
(increasing specific impulse by 85s, or 18.5%). However, beryllium is both expensive and toxic,
making this an unattractive option. It should be clarified that cost is a major project driver, so
Author: James Roper
| Page 179 of 262
APPENDIX
technical performance sufficient for a sensible payload fraction is adequate in comparison to high
costs associated with more advanced fuels.
Compounds of Boron with hydrogen (Boranes) are also proposed, however these are toxic and
hypergolic with moist air, making them extremely dangerous and as such unattractive (HUZEL,
Dieter K and Huang, David H, 1992, p.20). They exist due to the increased specific impulses in
comparison with hydrocarbons (410s) and their storable nature.
12.3.5.
SOLID AND HYBRID PROPELLANTS
All modern solid propellants use explosives, often as ‘energetic plasticizers’. This makes their manufacture and proving extremely complex and hazardous. In addition, oxidiser agents that may
be used, such as ammonium perchlorate, produce toxic and corrosive combustion products (14%
hydrochloric acid is quoted (SUTTON, George P and Biblarz, Oscar, 2001, p.493)). These two
considerations show that the safety of such propellants is low and high expense would be incurred
during their manufacture.
Solid propellants have the advantage of particularly long ‘shelf-lives’, meaning they can be stored for a long time (typically 5-25 years (SUTTON, George P and Biblarz, Oscar, 2001, p.489)).
However, this is of no use to a launcher that is used on a regular basis. It is more typically applied
to silo-based nuclear missiles.
A typical solid propellant mixture for high specific impulse would include liquid high explosives,
such as nitro-glycerine, absorbed in to a stabilising agent. Handling of such explosives is
extremely hazardous. In addition, poor manufacturing can lead to accumulations of the explosive
material in a single location, resulting in a detonation that may damage or destroy the motor
casing.
The highest quoted specific impulse of a solid propellant is 270s, and uses nitroamine high
explosive HMX. This is relatively low when compared with liquid fuels. Another technical aspect
relating to these fuels is their uncontrollable burn characteristics – once ignited, they cannot be
shut-down or throttled. As a result, any malfunction could not be dealt with except by jettisoning
the motor, which would follow an unpredictable and dangerous trajectory.
Hybrid propellants circumvent the unattractive use of explosives in solid propellants. In a hybrid
system, a liquid oxidiser such as liquid oxygen is typically used, with a stable solid fuel such as
butadiene, dramatically improving safety. In normal conditions, the fuel and oxidiser will not
combust due to insufficient activation heat. In addition, the solid propellant is easy to handle and
load in to the casing.
Author: James Roper
| Page 180 of 262
APPENDIX
The use of a liquid oxidiser allows throttling and engine stop or re-start in flight. The specific
impulse of polymer/liquid-oxidiser systems are comparable to those for liquid hydrocarbon
systems.
At present, hybrid engines are not employed by any agency. The solid motors suffer from
combustion instability and uneven pressure characteristics at larger scales (SUTTON, George P
and Biblarz, Oscar, 2001, p.599).
Common to both solid and hybrid systems, the motor structures are placed under extreme
operating conditions over a wider surface area than in fully liquid systems. This entails that the
entire casing must be a high-pressure vessel, leading to poor fuel mass to structural mass ratios
(typically 70-90%) (SUTTON, George P and Biblarz, Oscar, 2001, p.542), excluding the engine
components such as nozzles (and pumps in the case of the hybrid).
In addition, carbon dioxide is an inevitable product of combusting a polymer, which is
environmentally unattractive. Polymers are based on hydrocarbons and as a result will follow the
trends for increased cost as oil becomes scarcer, making their potential as a ‘future’ fuel less credible. An opportunity may exist to base the fuel on purely re-cycled plastic for some time,
shredded and melted to form fuel cores. This is an attractive proposal, but would need to be a
short-term, short-duration solution accompanied by a longer-term launcher project.
Solid-fuelled motors can be made re-usable. NASA’s Space Shuttle solid rocket booster (SRB) is
one example of this. According to NASA’s SRB overhaul practices documentation (JPL, pp.1-6)
this is achieved with washing, non-destructive testing (X-Ray and ultrasound) and re-painting.
The SRB is disassembled in to three casing stages and the nozzle stage, and re-assembled with
mechanical fasteners. A functional test at 1.125 times the maximum operating pressure is also
employed. This system has 10 flight-reusability of parachutes, 20 flight-reusability of the nozzle,
thrust vectoring and electrical systems and 40 flight-reusability for main structural elements,
including the casing stages.
In summary, the dangerous nature of pure solid propellants is highly unattractive, as is their lack
of controllability. Hybrid propellants, with their higher performance and safety, do show
promising and attractive features. They are at present unable to take advantage of the high specific
impulse (and therefore payload fraction) of hydrogen due to the increased molecular weight of
their products.
12.3.6.
INERT PROPELLANTS
Inert propellants offer the benefit of handling safety. They are also widely available – substances
such as water or nitrogen could be used – which ensures that the propellants themselves would be
Author: James Roper
| Page 181 of 262
APPENDIX
cheap. In addition, the inert nature of these propellants makes them easy to design for (at least,
from the vehicle perspective).
However, the provision of energy externally is a major challenge. Calculations based on existing
engines and ideal conversion of heat energy in to velocity result in Gigawatt-range power
requirements (to achieve lift-off). Nuclear power sources would be the optimum solution as they
have very high power to mass ratios. However, radioactive debris may result even during normal
operation and would be almost certain if an accident occurred, unless highly massive
reinforcement is employed.
Calculations have been performed to show that the most energy-dense Li-ion batteries cannot be
made light-weight enough to supply power to the propellant. Carrying the power supply aboard a
launch vehicle would be impossible, due to its own mass exceeding that of the propellant.
Power could be provided externally. Laser beams at frequencies not absorbed by atmospheric
gasses are one possibility. Lasers can provide the required power in pulses lasting a fraction of a
second, but cannot provide it continuously (due to power generation limitations). Such lasers
would be huge and extremely expensive, as would their power supply.
In addition, there is a need to consider the reliability of such a system. Riding an externally
provided beam of laser light would require extremely accurate flight profiles. The laser light
would need to be re-directed to compensate for the vehicle passing over the Earth’s horizon.
Reflected laser light would be extremely dangerous to any aircraft and the surrounding terrain.
Due to the extensive technical issues involved in providing power externally and the enormous
mass of the power supply, this method of propulsion is clearly expensive and unlikely to succeed
with foreseeable technologies. As a result, propellants that provide their own energy, as is
commonly practiced today, will be the only option to discuss.
Author: James Roper
| Page 182 of 262
12.4.
APPENDIX
ROCKET SIZING SPREADSHEET
A spreadsheet was constructed in order to calculate the sizing features of the rocket – most
notably, the fuel mass requirements and the launch mass. In addition, the spreadsheet was used to
determine the value of increasing the number of stages, in terms of the vehicle launch mass.
In principle, this was achieved as follows. A top-down approach was taken, whereby the last stage
mass was calculated first, then used as the effective payload (MP_eff) for the preceding stage. The
table is laid out such that vehicle configurations of 1, 2 or 3 stages are compared in adjacent
columns, and individual stage masses are calculated in rows. See Table 35.

The velocity requirement of each stage (dV/stage) was calculated as an equal share of the
total velocity requirement (V_BO). This reflects the optimum solution, since equal
velocity requirements drive equal fuel:dry mass fractions (as explained in “ENABLING TECHNOLOGY: STAGING” in the main report).

The fuel:dry mass ratio was calculated from the rocket equation, 𝑒𝑥𝑝
−1=
𝑚 /𝑀.

Using the payload (or the effective payload if this stage carries another above it), the fuel
mass was calculated. This took in to consideration the dry mass fraction of the stage
“SIG” (which may include either the engine and the tank mass, or the tank alone if a
single engine is used). The formula to include the tank mass was determined to be
𝑚 =

×
×
.
The tank mass “mT” is also calculated, as the tank:fuel mass ratio “SIG” multiplied by
the fuel mass.

The cumulative mass of this stage and any stages or payloads above it is then calculated.
This will act as the effective payload for the next-lowest stage.
The sum of this data is then used to determine the launch mass of the vehicle. From this, the
engine mass requirement is predicted using a typical engine thrust to weight ratio of ~80. This
engine mass is compared to the predicted engine mass, and further iterations are carried out as
necessary to match the engine mass to the T/W prediction. This is applicable only in the case
where the engine is considered as being with the launch vehicle through all stages, rather than
using different engines on each stage.
Additionally, the propellant masses are used to calculate the fuel and oxidiser masses, by merit of
their stoichiometric mixture ratios. This is then used to determine the required volume of each
fuel, and the resultant tank dimensions, using the density of the propellant. The payload fraction is
also calculated.
Author: James Roper
| Page 183 of 262
APPENDIX
9100
9.807
380
1
9100
2
4550
3
3033.333333
Stage mass ratio
1
2
3
10.49
0.00
0.00
2.39
2.39
0.00
1.26
1.26
1.26
1,275,206.32
76,512.38
45,000.00
1,396,718.70
125,575.83
7,534.55
45,000.00
178,110.38
61,169.44
3,670.17
45,000.00
109,839.61
497,030.18
29,821.81
178,110.38
704,962.37
149,307.28
8,958.44
109,839.61
268,105.32
Calculation of the
fuel:dry mass ratio
V_BO
g
Isp
stages
dV/stage
Sizing stage 3
15000
30000
45000
0.06
Stage 3
mf (kg)
mT (kg)
mP_eff (kg)
stage 3 mass
Sizing stage 2
M(Engine)
M(Payload)
M
SIG
Stage 2
mf (kg)
mT (kg)
mP_eff (kg)
stage 2+3 mass
Sizing stage 1
INPUT VARIABLES
Table 35: An example of the spreadsheet used to calculate vehicle sizing and compare stage configurations. This
example is for LH2 fuel.
Stage 1
mf (kg)
mT (kg)
mP_eff (kg)
stage 1+2+3 mass
stages
Total Launch Mass
Payload (mass%)
Author: James Roper
364,441.17
21,866.47
268,105.32
654,412.96
1
1,396,718.70
2.15%
2
704,962.37
4.26%
3
654,412.96
4.58%
| Page 184 of 262
12.5.
APPENDIX
MATLAB PROGRAM FOR LAUNCH TRAJECTORY.
A MATLAB program was produced in order to simulate the launch profile. This was designed to
ensure that factors previously applied to estimate the effects of drag and gravity were sufficient
and that a correct attitude, velocity and altitude would be obtained. It was also used to ensure that
the crew would not suffer excessive accelerations during the flight (no more than 4G).
The simplifying assumption of a “flat Earth” was applied, in order to prevent difficulties associated with the MATLAB trigonometry functions (namely, that the inverse sine or cosine of a
given amount results in 2 possible angles, over a circle). In addition, Earth’s rotation has been ignored. Principally, the program calculates the forces on the rocket body and the present mass
(which decreases linearly due to fuel burn). In general, the process is summarized in Figure 103
Figure 103: Process map for MATLAB launch simulation
The initial masses of each stage, including fuel mass and tank mass (and those masses of the
stages ‘above’ this stage) were obtained from the excel spreadsheet calculation detailed in Appendix JIM 2.
The forces on the body are its thrust and its drag. Thrust is estimated based upon the desired
initial acceleration, using the launch mass multiplied by the desired acceleration and accounting
for gravity. This is converted to a fuel mass flow rate by dividing by exhaust velocity.
𝑇ℎ𝑟𝑢𝑠𝑡,
𝑇=
(𝑑𝑒𝑠𝑖𝑟𝑒𝑑 𝑎𝑐𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛 + 𝑎𝑐𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑑𝑢𝑒 𝑡𝑜 𝑔𝑟𝑎𝑣𝑖𝑡𝑦)
𝑙𝑎𝑢𝑛𝑐ℎ 𝑚𝑎𝑠𝑠
𝑒𝑥ℎ𝑎𝑢𝑠𝑡 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦,
Author: James Roper
𝑣 =𝑔 𝐼
| Page 185 of 262
APPENDIX
𝑚 = ̇ 𝑇/𝑣
𝑓𝑢𝑒𝑙 𝑚𝑎𝑠𝑠 𝑓𝑙𝑜𝑤 𝑟𝑎𝑡𝑒,
Drag is estimated based upon a coefficient of drag for a slender body of 0.05 and the International
Standard atmosphere at the present altitude (with density extrapolated up to 170000km, then
assumed to be zero).
𝑑𝑒𝑛𝑠𝑖𝑡𝑦,
𝜌(ℎ < 170𝑘𝑚) = 𝑎𝑡𝑚𝑜𝑠𝑐𝑜𝑒𝑠𝑎,
1
𝑞 = 𝜌𝑣
2
𝑑𝑦𝑛𝑎𝑚𝑖𝑐 𝑝𝑟𝑒𝑠𝑠𝑢𝑟𝑒,
𝐷𝑟𝑎𝑔,
𝜌(ℎ > 170𝑘𝑚) = 0
𝐷 = 𝑞𝐶 𝐴
These forces are divided by the mass to expresses them as accelerations, resolved to the launchsite x and y (down-range and altitude) directions. Additionally, the acceleration due to gravity is
included in the y acceleration.
The flight angle of the vehicle is assumed to be obtained by a combination of thrust vectoring and
gravity-turning. This is modelled, with the flight angle relative to the ground-horizontal. The
minimum flight angle is calculated – this is the angle at which thrust exactly balances weight,
giving a zero vertical acceleration. In order to give the vehicle some upward acceleration, this
angle is averaged with a factor (facta, factb, factc), which is determined by trial and error to
achieve the desired flight conditions at the end of each burn.
𝑚𝑖𝑛𝑖𝑚𝑢𝑚 𝑓𝑙𝑖𝑔ℎ𝑡 𝑎𝑛𝑔𝑙𝑒,
𝑎𝑝𝑝𝑙𝑖𝑒𝑑 𝑓𝑙𝑖𝑔ℎ𝑡 𝑎𝑛𝑔𝑙𝑒,
𝑤ℎ𝑒𝑟𝑒 𝜃
𝜃
= 𝑎𝑠𝑖𝑛
𝜃=
𝜃
𝑣𝑒ℎ𝑖𝑐𝑙𝑒 𝑤𝑒𝑖𝑔ℎ𝑡
𝑇ℎ𝑟𝑢𝑠𝑡 − 𝐷𝑟𝑎𝑔
+ 𝜃
2
𝑖𝑠 𝑎 𝑤𝑒𝑖𝑔ℎ𝑡𝑖𝑛𝑔 d𝑒𝑠𝑖𝑔𝑛𝑒𝑑 𝑡𝑜 𝑎𝑐ℎ𝑒𝑖𝑣𝑒 𝑠𝑜𝑚𝑒 𝑣𝑒𝑟𝑡𝑖𝑐𝑎𝑙 𝑎𝑐𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛
The resolved accelerations and the time step are then used to predict the velocity and the position.
Staging is accounted for by repeating the process above, having subtracted any jettisoned masses
(i.e. empty tanks). At the end of each of these loops, the cut-off velocity, altitude and down-range
distance are stored in arrays (vco, hco, rangeco, where *co indicates stage cut-off properties).
The velocities, positions, accelerations and flight angles are plotted relative to flight time in
Figure 104.
Author: James Roper
| Page 186 of 262
APPENDIX
Figure 104: Flight data for the three stage vehicle, versus flight time. This demonstrates that orbital velocity can
be achieved. A final coasting and orbital-manoeuvring stage have not been shown
Author: James Roper
| Page 187 of 262
12.5.1.
APPENDIX
CODE FOR THE MATLAB PROGRAM.
clear
clc
%flight angle weighting factors (for averaging with minimum angle
facta = 80
factb = 20
factc = 0
%frontal vehicle area, for drag determination.
area = 113;
%payload (stage 1, includes mass of payload, stage 2 and stage 3)
mpa = 268105.3197;
% structure and fuel
msa = 21866.47008;
mfa = 364441.168;
%initial accn
aini = 9.81*2;
%exhaust vel
ve = 3727.8;
%thrust
T = aini* (mpa+mfa+msa);
%fuel mass flowrate
mdotf = T/ve;
%burn time
bt = round(mfa/mdotf);
%drag coef for streamlined body
coefd = 0.05;
%initial conditions.
n=1;
time(n) = 0;
Author: James Roper
| Page 188 of 262
APPENDIX
accnx(n) = 0;
vx(n) = 0;
sx(n) = 0;
accny(n) = 0;
vy(n) = 0;
v(n)=0
sy(n) = 0;
nmax = 1000;
q(n)=0;
%total mass
mt(1) = mpa+mfa+msa;
%launch angle
theta(n) = (asin((mt*9.81)/(T*0.9)));
Drag(1)=0;
btt=0;
btt= btt+bt;
dt = bt/nmax;
z=0;
% FIRST STAGE FIRING LOOP
while time<btt
n=n+1;
z=z+1;
time(n)=time(n-1)+dt;
% Determine atmospheric density at the indicated altitude sy.
if sy(n-1)<170000
[Temp sound press rho(n)] = atmoscoesa(sy(n-1));
else
rho(n)=0;
end
Author: James Roper
| Page 189 of 262
APPENDIX
q(n) = 0.5*rho(n)*(v(n-1))^2;
mt(n) = mpa+msa+mfa - mdotf*dt*z;
Drag(n) = q(n)*area*coefd;
% Calculate flight angle from minimum averaged with weighting factor
theta(n) = deg2rad((facta+rad2deg(asin(
Drag(n)) )))/2);
(mt(n)*9.81)
/
(T*0.9-
accnx(n) = cos(theta(n))*(T-Drag(n))/(mt(n));
accny(n) = sin(theta(n))*(T-Drag(n))/(mt(n))-9.81;
vx(n) = vx(n-1) + (dt)*accnx(n);
vy(n) = vy(n-1) + (dt)*accny(n);
sx(n) = sx(n-1) + (dt)*vx(n);
sy(n) = sy(n-1) + (dt)*vy(n);
v(n) = (vx(n)^2+vy(n)^2)^(1/2);
accn(n) = (accnx(n)^2+accny(n)^2)^(1/2);
end
vco(1)=v(n);
hco(1)=sy(n);
rangeco(1)=sx(n);
%payload
mpa = 109839.6076;
msa = 8958.436539;
mfa = 149307.2756;
%initial accn
aini = 9.81*1.8;
%exhaust vel
Author: James Roper
| Page 190 of 262
APPENDIX
ve = 3727.8;
%thrust
%fuel mass flowrate
mdotf = T/ve;
%burn time
btt= btt+bt;
dt = bt/nmax;
z=0;
% SECOND STAGE FIRING LOOP
while time<btt
n=n+1;
z=z+1;
if sy(n-1)<170000
else
rho(n)=0;
end
q(n) = 0.5*rho(n)*(v(n-1))^2;
theta(n) = deg2rad((factb+rad2deg(asin(
Drag(n)) )))/2);
(mt(n)*9.81) /
(T*0.9-
Author: James Roper
| Page 191 of 262
APPENDIX
sx(n) = sx(n-1) + (dt)*vx(n);
sy(n) = sy(n-1) + (dt)*vy(n);
v(n) = (vx(n)^2+vy(n)^2)^(1/2);
end
vco(2)=v(n);
hco(2)=sy(n);
rangeco(2)=sx(n);
%payload
mpa = 45000;
msa = 3670.166465;
mfa = 61169.44109;
%initial accn
aini = 9.81*1.8;
%exhaust vel
ve = 3727.8;
%thrust
%fuel mass flowrate
mdotf = T/ve;
%burn time
btt= btt+bt;
dt = bt/nmax;
z=0;
% THIRD STAGE FIRING LOOP
Author: James Roper
| Page 192 of 262
APPENDIX
while time<btt
n=n+1;
z=z+1;
if sy(n-1)<170000
else
rho(n)=0;
end
q(n) = 0.5*rho(n)*(v(n-1))^2;
theta(n) = deg2rad((factc+rad2deg(asin(
Drag(n)) )))/2);
(mt(n)*9.81) /
(T*0.9-
sx(n) = sx(n-1) + (dt)*vx(n);
sy(n) = sy(n-1) + (dt)*vy(n);
v(n) = (vx(n)^2+vy(n)^2)^(1/2);
end
vco(3)=v(n);
hco(3)=sy(n);
rangeco(3)=sx(n);
vco
Author: James Roper
| Page 193 of 262
APPENDIX
hco
rangeco
subplot(2,5,1), plot(time,accnx/9.81), title('Gs in x')
subplot(2,5,2), plot(time,accny/9.81), title('Gs in y')
subplot(2,5,3), plot(time,vx), title('V in x, m/s')
subplot(2,5,4), plot(time,vy), title('V in y, m/s')
subplot(2,5,5), plot (time,sx/1000), title('km x')
subplot(2,5,6), plot(time,sy/1000), title('km y')
subplot(2,5,7),
horizontal)')
plot(time,rad2deg(theta)),
title('Theta
(to
subplot(2,5,8), plot(time,accn/9.81), title('total G')
subplot(2,5,9), plot(time,v), title('total V')
Author: James Roper
| Page 194 of 262
12.6.
APPENDIX
TANK DESIGN
Table 36 Initial Tank Sizing Calculations
Al-Li 2195
Maximum allowable stress (Pa)
p critical
π
3.141592654
2685 kg/m3
Density
Height (m)
16
20
20
5.00E+05 bar
6.09E+08 Pa
Diameter (m) Radius (m)
3.22
1.61
3.22
1.61
3.69
1.845
X1
X2
X4
Wall thickness, t (m)
Cylinder Spherical end
1.32E-03
6.61E-04
1.32E-03
6.61E-04
1.52E-03
7.58E-04
Weight (kg)
Cylinder Spherical end
574.8
28.92
718.5
28.92
943.6
43.52
Single tank (kg)
633
776
1031
Weight (kg)
633
X2
1553
X4
4123
TOTAL
6308
30% extra
OVERALL
1892
8200
Table 37 Final tank sizing for 31 m long tank
Thickness (m)
1.26E-02
2.50E-03
2.60E-03
2.70E-03
2.00E-03
Area (m2)
1.19E-01
2.36E-02
2.45E-02
2.54E-02
1.88E-02
σcr (Pa)
1.19E+08
2.37E+07
2.46E+07
2.56E+07
1.89E+07
Pcr (N)
14207651
557633
603136
650423
356885
MS
24.909
0.017
0.100
0.186
-0.349
Mass (kg)
9899
1961
2040
2118
1569
Table 38 Final tank sizing for 25 m long tank
Thickness (m)
6.62E-03
2.50E-03
2.60E-03
2.70E-03
2.00E-03
Area (m2)
6.24E-02
2.36E-02
2.45E-02
2.54E-02
1.88E-02
σcr (Pa)
6.27E+07
2.37E+07
2.46E+07
2.56E+07
1.89E+07
Pcr (N)
3908338
557633
603136
650423
356885
MS
7.679
0.238
0.339
0.444
-0.207
Mass (kg)
4187
1582
1645
1708
1265
Table 39 Final tank sizing for 21 long tank
Thickness (m)
3.92E-03
2.50E-03
2.60E-03
2.70E-03
2.00E-03
Area (m2)
3.70E-02
2.36E-02
2.45E-02
2.54E-02
1.88E-02
σcr (Pa)
3.71E+07
2.37E+07
2.46E+07
2.56E+07
1.89E+07
Pcr (N)
1372992
557633
603136
650423
356885
MS
2.567
0.449
0.567
0.690
-0.073
Mass (kg)
2085
1329
1382
1435
1063
| Page 195 of 262
12.7.
APPENDIX
RE-ENTRY SIMULATION CALCULATIONS
The following list details the variables used in the calculations:
Atmospheric density (kg/m3)
ρa,n
Atmospheric pressure (Pa)
Pn
Atmospheric temperature (˚C)
Ta,n
Deceleration chute drag coefficient
cD,d
Descent distance between deceleration chute
Δzd,n
deployments (m)
Descent distance
deployments (m)
between
landing
chute
Δzl,n
Equivalent ground distance (m)
ln
Final altitude (m)
hf,n
Final angle (degrees)
θf,n
Final horizontal distance (m)
df,n
Final horizontal velocity (m/s)
vh,n
Final running time (s)
tf,n
Final surface temperature (K)
Ts,f,n
Final under-surface temperature (K)
Tu,f,n
Final vertical velocity (m/s)
vv,n
Heat tile density (kg/m3)
ρh
Heat tile emissivity
ε
Heat tile specific heat capacity (J/kgK)
C
Heat tile thermal conductivity (W/mK)
k
Heat tile thermal diffusivity (mm2/s)
α
Heat tile thickness (mm)
x
Heat tiles weight (kg)
w
Heating rate (W/m2)
𝑞̇ n
Horizontal acceleration (m/s2)
ah,n
Initial altitude (m)
hi,n
Initial angle (degrees)
θi,n
| Page 196 of 262
APPENDIX
Initial deceleration chute deployment altitude
zd
(m)
Initial horizontal distance (m)
di,n
Initial horizontal velocity (m/s)
uh,n
Initial landing chute deployment altitude (m)
zl
Initial re-entry velocity (m/s)
ur
Initial running time (s)
ti,n
Initial surface temperature (K)
Ts,i,n
Initial under-surface temperature (K)
Tu,i,n
Initial vertical velocity (m/s)
uv,n
Landing chute drag coefficient
cD,l
Number of deceleration chutes
m
Number of landing chutes
n
Parachute cross-sectional area (m2)
Ap,n
Radius of deceleration chutes (m)
rd
Radius of landing chutes (m)
rl
Re-entry angle (degrees)
θr
Time step (s)
Δtn
Total descent time (s)
tT
Total range (m)
lT
Vehicle base radius (m)
rv
Vehicle cross-sectional area (m2)
Av
Vehicle drag coefficient
cD,v
Vehicle mass (kg)
M
Vertical acceleration (m/s2)
av,n
Vertical acceleration in G’s (G)
Gn
| Page 197 of 262
,
Column 2: 𝑣
,
=
.
,
,
+ 𝑎
×
,
.
⁄
⎝
⎛
× 0.1 + ⎜
⎝
⎛
𝐼𝐹 ⎜ℎ
⎜
Column 5:
∆𝑡 =
Column 4: 𝐺 =
⎠
⎞
0.1 ⎟ > 0,
⎝
,
= 𝐼𝐹 𝑢
= 𝑢 × 𝑠𝑖𝑛
⎛
0, 𝐼𝐹 ⎜ℎ , + 𝑢
⎝
⎛
𝐼𝐹 ⎜∆𝑡 >
Column 3:
𝑎 , =
,
Column 1: 𝑢
FIRST ROW
,
×
. ×
,
.
⁄
,
×
× 0.1 ≤ 0, 𝑢
,
,
×
,
+ 𝑎
,
×
.
×
. ×
,
×
×
× ∆𝑡 , 0
The following calculations were used to create the simulations:
,
, ×
,
×
.
,
×
×
,
, ×
,
| Page 198 of 262
. ×
×
×
. ×
APPENDIX
,
,
×
,
.
,
,
.
,
×
×
,
×
,
, ×
,
.
,
,
.
∆
×
,
×
⎠
⎠
⎞ ⎞
⎟ , 0⎟
,
, ×
×
,
.
⁄
,
×
× 0.1 +
. ×
,
×
,
,
,
= ∆𝑡 + 𝑡 ,
×
.
×
×
,
+ 𝑎
,
⎠
⎞
× ∆𝑡 , 0⎟
⁄
×
. ×
,
×
, ×
,
,
. ×
×
,
.
×
×
,
×
∆
,
,0
≤ 𝑛, 𝐴𝐵𝑆 𝑅𝑂𝑈𝑁𝐷𝑈𝑃
,
,
,0
∆
, ×
×
.
,
,
×
.
×
,
,
×
,
×
, ×
,
≤ 𝑚, 𝐴𝐵𝑆 𝑅𝑂𝑈𝑁𝐷𝑈𝑃
. ×
,
,𝑛 ,0
,0
,
APPENDIX
| Page 199 of 262
∆
,
= 𝐼𝐹 ℎ , > 𝑧 , 𝐼𝐹 ℎ , > 𝑧 , 0, 𝐼𝐹 𝑚 > 0, 𝐼𝐹 𝐴𝐵𝑆 𝑅𝑂𝑈𝑁𝐷𝑈𝑃
× ∆𝑡
,
.
, 𝐼𝐹 𝑛 > 0, 𝐼𝐹 𝐴𝐵𝑆 𝑅𝑂𝑈𝑁𝐷𝑈𝑃
,
𝑟 ,0
Column 10: 𝐴
⎠
,
⎝
⎛
× 0.1 + ⎜
⎞
0.1 ⎟ > 0, ℎ , + 𝑢
⎝
⎛
𝐼𝐹 ⎜ℎ , + 𝑢
Column 9:
ℎ, =
Column 8: ℎ , = [𝑖𝑛𝑝𝑢𝑡 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒, 𝑠𝑒𝑡 𝑎𝑠 350000]
Column 7: 𝑡
Column 6: 𝑡 , = [𝑖𝑛𝑝𝑢𝑡 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑡𝑖𝑚𝑒, 𝑠𝑒𝑡 𝑎𝑠 0]
⎝
⎛
⎜
⎝
⎛
0,0, 𝐼𝐹 ⎜ℎ , + 𝑢
,
,
∆
.
,
×
,0
.
⎠
×
,
⎠⎠
⎞⎞
⎟⎟
⎟
, ×
,
,𝑚 × 𝜋 ×
,
⎞
× 0.1 ⎟ ≥ 0,0.1,
×
,
,
Column 13: 𝑢
Column 14: 𝑣
,
= 𝐼𝐹 𝑢
,
×
× ∆𝑡
> 0,
× ∆𝑡
, 𝐼𝐹 𝑣
. ×
,
,
×
,
+ 𝑎
,
,
,
=𝑑, + 𝑢
,
× ∆𝑡
× 𝑆𝑄𝑅𝑇 𝑢
,
+𝑢
,
,
× ∆𝑡
,
×
.
×
,
+ 0.5 × 𝑎
× 𝑆𝑄𝑅𝑇
× ∆𝑡
,
,
, ,
,
Column 23: 𝑇
Column 24: 𝜌
̇
× .
×
×
×
× ,,
×∆
,
×
× ∆𝑡
, ×
=
. ×
= 𝐼𝐹 𝑇
,
,,
> 𝑇 , , ,𝑇
,,
+
(
)×
,,
,,
×
×
∆
×
,
−
𝑢
. ×
,
(
,
,
,
)×
× ∆𝑡
×
,
,,
,
×
,
,,
×
×
+ 0.5 × 𝑎
×
APPENDIX
| Page 200 of 262
,𝑇
+
,,
− 𝐴𝑆𝐼𝑁
= [𝑖𝑛𝑝𝑢𝑡 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑢𝑛𝑑𝑒𝑟 − 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑡𝑒𝑚𝑝𝑒𝑟𝑎𝑡𝑢𝑟𝑒, 𝑠𝑒𝑡 𝑎𝑠 288]
=𝑇,, +
,,
,
Column 22: 𝑇
Column 21: 𝑇 ,
Column 20: 𝑇 , , = [𝑖𝑛𝑝𝑢𝑡 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑡𝑒𝑚𝑝𝑒𝑟𝑎𝑡𝑢𝑟𝑒, 𝑠𝑒𝑡 𝑎𝑠 116]
Column 19: 𝑞̇ = 1.83 × 10
𝑢
+ 0.5 × 𝑎
Column 18: 𝑙 = 6378100 × 𝐴𝑆𝐼𝑁 𝑆𝑄𝑅𝑇
Column 17: 𝑑
,𝑣
× ∆𝑡 , 0
= 0,0, 𝐴𝑇𝐴𝑁2 𝑣
≥ 0, 𝑢
,
Column 16: 𝑑 , = [𝑖𝑛𝑝𝑢𝑡 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 ℎ𝑜𝑟𝑖𝑧𝑜𝑛𝑡𝑎𝑙 𝑑𝑖𝑠𝑡𝑎𝑛𝑐𝑒, 𝑠𝑒𝑡 𝑎𝑠 0]
,
,
= 0,
+ 𝑎
= 𝑢 × 𝑐𝑜𝑠
= 𝐼𝐹 𝑣
𝐼𝐹 ∆𝑡 > 0, 𝐼𝐹 ℎ , − 𝑢
Column 15:
𝑎 , =
,
Column 12: 𝜃
Column 11: 𝜃 , = 𝜃
∆
,
,
,
×
× ∆𝑡
.
.
×
,
×
×
,
, ×
, ×
,
∆
,
,0
,
,
,
Column 2: 𝑣
Column 3: 𝑎
,
+𝑣
,
× 𝑠𝑖𝑛
= [𝑠𝑎𝑚𝑒 𝑓𝑜𝑟𝑚𝑢𝑙𝑎 𝑎𝑠 𝑓𝑖𝑟𝑠𝑡 𝑙𝑖𝑛𝑒]
,
,
,
,
,
Column 13: 𝑢
Column 14: 𝑣
Column 15: 𝑎
,
+𝑣
,
× 𝑐𝑜𝑠
,
Column 17: 𝑑
,
= 𝑆𝑄𝑅𝑇 𝑣
Column 16: 𝑑 , = 𝑑
,
Column 12: 𝜃
,
,
Column 11: 𝜃 , = 𝜃
Column 10: 𝐴
Column 9: ℎ
,
Column 8: ℎ , = ℎ
,
Column 6: 𝑡 , = 𝑡
Column 5: ∆𝑡 = [𝑠𝑎𝑚𝑒 𝑓𝑜𝑟𝑚𝑢𝑙𝑎 𝑎𝑠 𝑓𝑖𝑟𝑠𝑡 𝑙𝑖𝑛𝑒]
Column 7: 𝑡
.
= 𝑆𝑄𝑅𝑇 𝑣
Column 4: 𝐺 = [𝑠𝑎𝑚𝑒 𝑓𝑜𝑟𝑚𝑢𝑙𝑎 𝑎𝑠 𝑓𝑖𝑟𝑠𝑡 𝑙𝑖𝑛𝑒]
,
Column 1: 𝑢
,
.
,
,
×
×
.
.
×
APPENDIX
× ,
, 101.29 ×
| Page 201 of 262
× ℎ , , 𝐼𝐹 ℎ , > 11000, −56.46,15.04 − 6.49 × 10
, 𝐼𝐹 ℎ , > 11000,22.65 × 𝑒
= 𝐼𝐹 ℎ , > 25000, −131.21 + 2.99 × 10
SECOND ROW
Column 26: 𝑇
Column 25: 𝑃 = 𝐼𝐹 ℎ , > 25000,2.488 ×
×ℎ ,
.
,
.
+ 273.1
× 1000
, ,
,
Column 23: 𝑇
Column 24: 𝜌
, ,
=𝑇
,
,
,
𝑐
𝑐
= [𝑖𝑛𝑝𝑢𝑡 𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑐ℎ𝑢𝑡𝑒 𝑑𝑟𝑎𝑔 𝑐𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡, 𝑠𝑒𝑡 𝑎𝑠 0.78]
= [𝑖𝑛𝑝𝑢𝑡 𝑑𝑒𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑐ℎ𝑢𝑡𝑒 𝑑𝑟𝑎g 𝑐𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡, 𝑠𝑒𝑡 𝑎𝑠 0.9]
= [𝑖𝑛𝑝𝑢𝑡 𝑣𝑒ℎ𝑖𝑐𝑙𝑒 𝑑𝑟𝑎𝑔 𝑐𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡, 𝑠𝑒𝑡 𝑎𝑠 0.7]
𝑘
𝜌 ×∝× 10
𝜀 = [𝑖𝑛𝑝𝑢𝑡 ℎ𝑒𝑎𝑡 𝑡𝑖𝑙𝑒 𝑒𝑚𝑖𝑠𝑠𝑖𝑣𝑖𝑡𝑦, 𝑠𝑒𝑡 𝑎𝑠 0.9]
𝑥
𝑤=𝐴 ×
×𝜌
1000
𝐶=
𝑥 = [𝑖𝑛𝑝𝑢𝑡 ℎ𝑒𝑎𝑡 𝑡𝑖𝑙𝑒 𝑡ℎ𝑖𝑐𝑘𝑛𝑒𝑠𝑠, 𝑠𝑒𝑡 𝑎𝑠 150]
∝= [𝑖𝑛𝑝𝑢𝑡 ℎ𝑒𝑎𝑡 𝑡𝑖𝑙𝑒 𝑡ℎ𝑒𝑟𝑚𝑎𝑙 𝑑𝑖𝑓𝑓𝑢𝑠𝑖𝑣𝑖𝑡𝑦, 𝑠𝑒𝑡 𝑎𝑠 0.677]
𝑘 = [𝑖𝑛𝑝𝑢𝑡 ℎ𝑒𝑎𝑡 𝑡𝑖𝑙𝑒 𝑡ℎ𝑒𝑟𝑚𝑎𝑙 𝑐𝑜𝑛𝑑𝑢𝑐𝑡𝑖𝑣𝑖𝑡𝑦, 𝑠𝑒𝑡 𝑎𝑠 0.167]
𝜌 = [𝑖𝑛𝑝𝑢𝑡 ℎ𝑒𝑎𝑡 𝑡𝑖𝑙𝑒 𝑑𝑒𝑛𝑠𝑖𝑡𝑦, 𝑠𝑒𝑡 𝑎𝑠 280]
,
𝑐
INPUT AND OUTPUT VALUES
All subsequent lines are the same as the previous lines.
Column 26: 𝑇
Column 25: P = [𝑠𝑎𝑚𝑒 𝑓𝑜𝑟𝑚𝑢𝑙𝑎 𝑎𝑠 𝑓𝑖𝑟𝑠𝑡 𝑙𝑖𝑛𝑒]
,,
,
Column 22: 𝑇
Column 21: 𝑇 ,
Column 20: 𝑇 , , = 𝑇 , ,
Column 19: 𝑞̇ = [𝑠𝑎𝑚𝑒 𝑓𝑜𝑟𝑚𝑢l𝑎 𝑎𝑠 𝑓𝑖𝑟𝑠𝑡 𝑙𝑖𝑛𝑒]
Column 18: 𝑙 = [𝑠𝑎𝑚𝑒 𝑓𝑜𝑟𝑚𝑢𝑙𝑎 𝑎𝑠 𝑓𝑖𝑟𝑠𝑡 𝑙𝑖𝑛𝑒]
| Page 202 of 262
APPENDIX
9.81 × 6378100
6378100 + ℎ ,
…
)
…
)
APPENDIX
| Page 203 of 262
The maximum internal temperature is identified, and the altitude at which it occurs is also identified.
The maximum external temperature is identified, and the altitude at which it occurs is also identified.
(l1…n are all the equivalent ground distances from beginning to end of the simulation)
𝑙 = 𝑆𝑈𝑀(𝑙
(Δt1…n are all the time steps from beginning to end of the simulation)
𝑡 = 𝑆𝑈𝑀(∆𝑡
The total heat load is calculated.
The maximum heating rate is identified, and the altitude at which it occurs is also identified.
𝜃 = [𝑖𝑛𝑝𝑢𝑡 𝑟𝑒 − 𝑒𝑛𝑡𝑟𝑦 𝑎𝑛𝑔𝑙𝑒, 𝑠𝑒𝑡 𝑎𝑠 0]
(hi,1 is the first initial altitude, the only manual input altitude)
𝑢 = 𝑆𝑄𝑅𝑇
𝑟 = [𝑖𝑛𝑝𝑢𝑡 𝑣𝑒ℎ𝑖𝑐𝑙𝑒 𝑏𝑎𝑠𝑒 𝑟𝑎𝑑𝑖𝑢𝑠, 𝑠𝑒𝑡 𝑎𝑠 2.5]
∆𝑧 = [𝑖𝑛𝑝𝑢𝑡 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑑𝑖𝑠𝑡𝑎𝑛𝑐𝑒 𝑏𝑒𝑡𝑤𝑒𝑒𝑛 𝑑𝑒𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑐ℎ𝑢𝑡𝑒 𝑑𝑒𝑝𝑙𝑜𝑦𝑚𝑒𝑛𝑡𝑠, 𝑠𝑒𝑡 𝑎𝑠 0]
𝑧 = [𝑖𝑛𝑝𝑢𝑡 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑑𝑒𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑐ℎ𝑢𝑡𝑒 𝑑𝑒𝑝𝑙𝑜𝑦𝑚𝑒𝑛𝑡 𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒, 𝑠𝑒𝑡 𝑎𝑠 0]
𝑟 = [𝑖𝑛𝑝𝑢𝑡 𝑟𝑎𝑑𝑖𝑢𝑠 𝑜𝑓 𝑑𝑒𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑐ℎ𝑢𝑡𝑒𝑠, 𝑠𝑒𝑡 𝑎𝑠 0]
𝑚 = [𝑖𝑛𝑝𝑢𝑡 𝑛𝑢𝑚𝑏𝑒𝑟 𝑜𝑓 𝑑𝑒𝑐𝑒𝑙𝑒𝑟𝑎𝑡𝑖𝑜𝑛 𝑐ℎ𝑢𝑡𝑒s, 𝑠𝑒𝑡 𝑎𝑠 0]
∆𝑧 = [𝑖𝑛𝑝𝑢𝑡 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑑𝑖𝑠𝑡𝑎𝑛𝑐𝑒 𝑏𝑒𝑡𝑤𝑒𝑒𝑛 𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑐ℎ𝑢𝑡𝑒 𝑑𝑒𝑝𝑙𝑜𝑦𝑚𝑒𝑛𝑡𝑠, 𝑠𝑒𝑡 𝑎𝑠 1000]
𝑧 = [𝑖𝑛𝑝𝑢𝑡 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑐ℎ𝑢𝑡𝑒 𝑑𝑒𝑝𝑙𝑜𝑦𝑚𝑒𝑛𝑡 𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒, 𝑠𝑒𝑡 𝑎𝑠 9000]
𝑟 = [𝑖𝑛𝑝𝑢𝑡 𝑟𝑎𝑑𝑖𝑢𝑠 𝑜𝑓 𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑐ℎ𝑢𝑡𝑒𝑠, 𝑠𝑒𝑡 𝑎𝑠 10]
𝑛 = [𝑖𝑛𝑝𝑢𝑡 𝑛𝑢𝑚𝑏𝑒𝑟 𝑜𝑓 𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑐ℎ𝑢𝑡𝑒𝑠, 𝑠𝑒𝑡 𝑎𝑠 4]
𝐴 =𝜋×𝑟
30,000 30,000 9000 9000 1 2 3 4 184 17 105 12 Landing velocity (m/s) 235 611 298 775 Descent time (s) 340,724 340,741 340,725 340,734 Total range (m) 2144 2144 1348 1348 Maximum External Temperature (K) 289 298 290 298 Maximum Internal Temperature (K) Figure 1 Emissivity of virgin and charred standard PICA (Tran, et al., 1997) With regards to the following, Vehicle 1 is the 30 tonne capsule falling purely ballistically with no parachutes; Vehicle 2 is the 30 tonne capsule falling with 3 landing parachutes deployed, each with a radius of 15m, beginning at 10km, and with intervals of 2000m; Vehicle 3 is the 9 tonne capsule falling purely ballistically with no parachutes; Vehicle 4 is the key vehicle to note, as it is the 9 tonne capsule falling with 4 landing parachutes deployed, each with a radius of 10m, beginning at 9km, and with intervals of 1000m. Mass (kg) Vehicle Table 1 Table of the key values obtained from the MS Excel simulations 350000.00
300000.00
250000.00
Figure 2 Simulation results -‐ velocity against altitude 400000.00
Altitude (m)
200000.00
150000.00
100000.00
Velocity against Altitude
50000.00
0.00
-‐3,500.00
-‐3,000.00
-‐2,500.00
-‐2,000.00
-‐1,500.00
-‐1,000.00
-‐500.00
0.00
Velocity (m/s)
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
18000.00
16000.00
14000.00
12000.00
10000.00
Altitude (m)
Figure 3 Simulation results -‐ velocity against altitude -‐ focused 20000.00
8000.00
6000.00
4000.00
Velocity against Altitude
2000.00
0.00
-‐1,000.00
-‐900.00
-‐800.00
-‐700.00
-‐600.00
-‐500.00
-‐400.00
-‐300.00
-‐200.00
-‐100.00
0.00
Velocity (m/s)
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
0.00
0.000
50000.00
100000.00
150000.00
200000.00
250000.00
300000.00
350000.00
400000.00
100.000
200.000
Figure 4 Simulation results -‐ altitude against time Altitude (m)
300.000
500.000
Time (s)
400.000
600.000
Altitude against Time
700.000
800.000
900.000
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
0.00
175.000
5000.00
10000.00
15000.00
20000.00
25000.00
30000.00
225.000
275.000
Figure 5 Simulation results -‐ altitude against time -‐ focused Altitude (m)
Time (s)
325.000
375.000
Altitude against Time
425.000
475.000
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
350000.00
300000.00
250000.00
Figure 6 Simulation results -‐ heating rate against altitude 400000.00
Altitude (m)
200000.00
150000.00
100000.00
Heating Rate against Altitude
50000.00
0.00
0.000
1,000,000.000
2,000,000.000
3,000,000.000
4,000,000.000
5,000,000.000
6,000,000.000
7,000,000.000
8,000,000.000
9,000,000.000
Heating Rate (W/m2)
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
0
0.000
500
1,000
1,500
2,000
2,500
50.000
100.000
Time (s)
150.000
200.000
Temperature against Time
Figure 7 Simulation results -‐ vehicle 1 temperature against time Temperature (K)
250.000
Vehicle 1 Internal Temperature
Vehicle 1 External Temperature
0
0.000
500
1,000
1,500
2,000
2,500
100.000
200.000
300.000
Time (s)
400.000
500.000
600.000
700.000
0
0.000
200
400
600
800
1,000
1,200
1,400
1,600
50.000
100.000
150.000
Time (s)
200.000
250.000
300.000
350.000
0
0.000
200
400
600
800
1,000
1,200
1,400
1,600
100.000
200.000
300.000
500.000
Time (s)
400.000
600.000
700.000
800.000
900.000
-‐25
25
75
125
175
225
275
325
0
250
500
750
Figure 11 Simulation results -‐ g's against time Acceleration (g)
1000
1250
Time (s)
1500
1750
G's against Time
2000
2250
2500
2750
3000
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
-‐10
0
10
20
30
40
50
160
260
360
Figure 12 Simulation results -‐ g's against time -‐ focused Acceleration (g)
460
Time (s)
560
G's against Time
660
760
860
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
-‐10
0
10
20
30
40
50
160
180
200
220
Figure 13 Simulation results -‐ g's against time -‐ focused double Acceleration (g)
240
Time (s)
260
G's against Time
280
300
320
340
Vehicle 4
Vehicle 3
Vehicle 2
Vehicle 1
12.8.
PROJECT MANAGEMENT
12.8.1.
Author: Group
APPENDIX
INITIAL GANTT-CHART
| Page 217 of 262
12.8.2.
Author: Group
APPENDIX
FINAL GANTT-CHART
| Page 218 of 262
29
22
15
8
1
Informal Meetings
Group Meetings
Initial Group Meetings
Tuesday
October 2012
MDDP
Exam
Deadline
Monday
Informal Meetings
Informal Meetings
Informal Meetings
2
Wednesday
3
Thursday
Friday
M
T
F
7
1
S
9
2
S
8
1
M
9 10 11 12 13 14
2
3
T W
4
T
5
F
6
S
7
S
October 2012
8
September 2012
T W
6
M
5
8
1
T
9 10 11
2
F
3
S
4
S
November 2012
7
T W
6
12 13 14 15 16 17 18
5
19 20 21 22 23 24 25
4
15 16 17 18 19 20 21
26 27 28 29 30
3
10 11 12 13 14 15 16
22 23 24 25 26 27 28
7
29 30 31
Sunday
17 18 19 20 21 22 23
6
24 25 26 27 28 29 30
Saturday
14
5
13
4
12
21
Informal Meetings
20
28
11
19
27
10
26
4
9
25
3
18
Group Meetings
Informal Meetings
Group Meetings
Extended Group Meetings
Informal Meetings
Group Meetings
2
17
24
31
1
16
23
30
Informal Meetings
Page 1/5
26
19
12
5
29
Tuesday
November 2012
MDDP
Exam
Deadline
Monday
Informal Meetings
Law - Norman
Informal Meetings
Durability - Norman
Informal Meetings
Informal Meetings
Informal Meetings
30
6
13
20
27
Wednesday
31
7
14
21
28
Thursday
Informal Meetings
Group Meetings
Group Meetings
Informal Meetings
Informal Meetings
Group Meetings
Informal Meetings
Group Meetings
Informal Meetings
Group Meetings
M
2
3
T W
4
T
5
F
6
S
7
S
October 2012
1
M
8
1
T
9 10 11
2
F
3
S
4
S
November 2012
T W
7
M
3
T
6
F
7
8
1
S
9
2
S
December 2012
5
T W
4
10 11 12 13 14 15 16
6
17 18 19 20 21 22 23
5
12 13 14 15 16 17 18
24 25 26 27 28 29 30
9 10 11 12 13 14
15 16 17 18 19 20 21
19 20 21 22 23 24 25
8
22 23 24 25 26 27 28
26 27 28 29 30
31
29 30 31
4
11
Sunday
10
18
3
9
17
25
Saturday
8
16
24
2
2
15
23
1
Friday
22
30
1
29
Page 2/5
Tuesday
27
Wednesday
28
Thursday
Group Meetings
M
1
T
9 10 11
2
F
3
S
4
S
November 2012
T W
8
M
3
T
6
F
7
8
1
S
9
2
S
December 2012
5
T W
4
M
7
1
9 10 11 12 13
2
T W
3
T
4
F
5
S
6
S
January 2013
8
14 15 16 17 18 19 20
7
21 22 23 24 25 26 27
6
10 11 12 13 14 15 16
28 29 30 31
5
12 13 14 15 16 17 18
17 18 19 20 21 22 23
2
24 25 26 27 28 29 30
Sunday
19 20 21 22 23 24 25
31
Saturday
1
26 27 28 29 30
30
9
Friday
8
29
7
Informal Meetings
6
Group Meetings
16
5
15
4
14
23
Group Meetings
Informal Meetings
Informal Meetings
13
22
30
12
21
29
6
11
20
28
5
10
19
27
4
Advanced Manufacturing - Samuel
18
26
3
Business Srategy - Samuel
17
25
2
Fluid Dynamics- Aero
24
1
3
26
December 2012
MDDP
Exam
Deadline
Monday
Informal Meetings
Informal Meetings
Informal Meetings
Informal Meetings
31
Page 3/5
Wednesday
Thursday
3
Friday
M
T
F
7
8
1
S
9
2
S
December 2012
T W
6
M
7
1
9 10 11 12 13
2
T W
3
T
4
F
5
S
6
S
January 2013
8
M
4
T
7
8
1
F
9 10
2
S
3
S
February 2013
6
T W
5
11 12 13 14 15 16 17
5
18 19 20 21 22 23 24
4
14 15 16 17 18 19 20
3
10 11 12 13 14 15 16
25 26 27 28
6
21 22 23 24 25 26 27
Sunday
28 29 30 31
5
17 18 19 20 21 22 23
Saturday
24 25 26 27 28 29 30
31
4
13
2
12
1
11
Tuesday
10
Final Group Meetings
20
9
19
27
8
18
26
3
7
17
25
2
Poster Presentation
Informal Meetings
16
24
1
Research Skill - Emily
15
23
31
Strategic Managment - Samuel
14
22
30
31
January 2013
MDDP
Exam
Deadline
Monday
Durability - Norman
Informal Meetings
Final Report
Measurement - Norman
Advanced Manufacturing - Samuel
21
29
Oral Examinations
28
Page 4/5
M
2
T W
3
T
4
F
5
S
6
S
January 2013
1
M
T
7
8
1
F
9 10
2
S
3
S
February 2013
6
T W
5
M
4
T
7
8
1
F
9 10
2
S
3
S
March 2013
6
T W
5
11 12 13 14 15 16 17
4
18 19 20 21 22 23 24
9 10 11 12 13
11 12 13 14 15 16 17
25 26 27 28 29 30 31
8
18 19 20 21 22 23 24
7
14 15 16 17 18 19 20
25 26 27 28
3
21 22 23 24 25 26 27
Sunday
28 29 30 31
2
10
Saturday
9
17
1
8
16
24
Friday
30
7
15
23
3
Thursday
29
6
14
22
2
Wednesday
28
5
13
21
1
Tuesday
February 2013
MDDP
Exam
Deadline
Monday
4
12
20
28
Oral Examinations
11
19
27
31
18
26
Poster Presentation
25
Page 5/5

Reusable Launch Vehical - Group 1

Transcription

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