Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Re-usable Launch and Payload Delivery System MDDP Group 2
James Dobberson Robert Taylor Matthew Chapman Timothy West Mukudzei Muchengeti William Wou
Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
1. Contents 1. Contents ...... i 2. Executive Summary ...... ii 3. Introduction ...... 1 4. Down Selection and Integration Methodology ...... 2 5. Presentation of System Concept and Operations ...... 5 6. System Investment Plan ...... 20 7. Numerical Analysis and Statement of Feasibility ...... 23 8. Conclusions and Future Work ...... 29 9. Launch Philosophy ...... 31 10. Propulsion ...... 42 11. Structures and Fuel Systems ...... 51 12. Materials ...... 62 13. In Orbit Operations ...... 69 14. Electronics ...... 75 15. Re-Entry ...... 91 16. Landing ...... 103 17. Payload and Markets ...... 110 18. Component Mass Estimation ...... 128 19. Infrastructure ...... 129 20. Finance ...... 141 21. Sensitivity Analysis ...... 148 22. References ...... i 23. Appendices ...... A1 A. Launch Philosophy Appendix ...... A1 B. Propulsion Appendix ...... B1 C. Structures and Materials Appendix ...... C1 D. In Orbit Operations and Electronics ...... D1 E. Re-entry Appendix ...... E1 F. Landing Appendix ...... F1 G. Payloads and Markets Appendix ...... G1 H. Mass Estimating Relationships ...... H1 I. Project Management ...... I1
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2. Executive Summary This report examines the feasibility of producing and operating a new reusable launch system to deliver payloads to a Low Earth Orbit by consideration of the technical and financial viability of such a system over a typical service life. Since the Space Shuttle was retired by NASA there has not been a reusable launch system available to the commercial market. Sincethe1960’swhenthe early design work for the shuttle program began, many in the space industry have believed that a highly reusable payload delivery system could be more financially viable than a traditional expendable rocket; however such a system has yet to be realised. For this feasibility study a conceptual design for a reusable payload delivery system was developed; this was done by comparing all the possible options for each subsystem to find the most suitable solution to the design criteria. The comparison was undertaken using a down selection process which defined a set of criteria that were used compare the subsystem solutions. The criteria used for the down selection were technical viability, financial feasibility, development program risk, concept integration, environmental impact and system reusability. From these criteria it was possible to find the most suitable collection of subsystems that offered the most viable design approach. Once the most suitable solution had been proposed for each subsystem, these solutions were integrated into a conceptual design to prove that a viable payload delivery system could be created. The integration phase allowed the solutions from the down selection process to be formed into a conceptual design which could then be financially analysed. The key design areas of a reusable space vehicle were examined to develop the system concept and concept of operations. Analysis of Launch Philosophy and Propulsion lead to the selection of a traditional vertical launch rocket over other concepts such as a Single Stage to Orbit space plane or an Air Launched expendable rocket system. To best satisfy the markets outlined in the Inception Report the rocket was optimised during the integration phase to a two rocket system family. The A-variant able to lift payloads of 25,000kg using two stages and the B- Variant able to lift payloads of 40,000kg using three stages. Both variants have an extremely high degree of commonality using common engines and a common first stage. The rocket engines’ cores are common to all of the rocket stages in both variants, although the first stage uses thrust augmentation nozzles to increase the efficiency of the rocket engines when they are operating within earth’s atmosphere. The A-Variant would be capable of carrying a human module containing 10 passengers and a safety system and would be evolved into a 25 passenger module later in the operational life. Both A and B variants would be able to carry unmanned payloads, such as satellites. The optimal structural solution was found to be a semi-monocoque structure that incorporates stringers and long runs to increase the stiffness and reduce the likelihood of failure through buckling. The fuel within the fuels tanks would need to be controlled and stored correctly to make sure that the fuel travels through the fuel lines to the engines and to ensure that
Section 2 - Executive Summary Page ii Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 it does not slosh, altering the vehicles centre of mass during operations. The spacecraft electronics are responsible for controlling and monitoring the system. The electronics subsystems are linked using a SpaceWire communication bus. Each subsystem would contain a high level of redundancy, including at least 3 field programmable gate arrays to ensure that the subsystems can be controlled even if failures do occur. The system monitoring subsystem would allow data to be collected on the integrity of the spacecraft, increasing reliability and reusability. The system would take off vertically from the launch pad at a site close to the equator and travel east as this would provide the most efficient method of achieving a 200km Orbit. Once the spacecraft has achieved orbit it can complete an orbit transfer to the higher orbit as required by individual missions. Once the mission has been completed some payloads (such as humans) may need to be returned to earth. All modules that re-enter theearth’satmosphere, including reusable stages, would be subjected to a ballistic re-entry profile, the aerodynamic drag causing the heat shield surface to increase in temperature as the crafts velocity is reduced. Once the craft has decelerated, thrusters would be used to land the craft at the launch site using the same thrusters form the launch emergency escape system. Once the vehicle has returned to the launch site it would be taken for overhaul during which the subsystems would be examined. The data from the on-board monitoring system would be used to make sure that the system is safe for its next mission. Once maintenance is complete the modules would be reassembled and transported to the launch site by rail. Financial analysis was completed to compare the reusable launch system with an equivalent expendable rocket. It was found that the cost of a reusable launch system is highly dependent on the reoccurrence of turnaround costs and the overall length of maintenance between missions. As the maintenance time increases the number of systems that are required to satisfy the market demand increases proportionally. It was noted that, for the market used in this analysis, during the first 4 years before the system enters full service the costs of the reusable launch system are much higher than an expendable use rocket due to the small number of launches and the higher development and test costs for a reusable launch system. From this point onwards the reusable system becomes highly profitable and the system becomes less dependent on the market trends reducing project risk as compared to a expendable system. Research was completed so that a realistic predication of the volume of space traffic during the systems 30 year life span could be made. It was found that it is feasible for a reusable payload delivery system to be designed built and operated. With the predicted increase in growth in space travel over the next 40 years and advances in technology, the system is likely to be highly financially viable when compared with expendable rockets. Therefore it is recommended that the required investments be made to allow the development of a new reusable payload delivery system, using the concepts outlined in the report.
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3. Introduction This report is a feasibility study for a Re-usable Launch and Payload Delivery System. The aim was to conduct a study into the optimal sub-systems and operations for such a system, generate an investment plan for the optimal subsystems so as to make a statement of the feasibility of a reusable launch system. The design brief outlined an initial launch date of approximately 2022 and for this report entry into service was predicted to be approximately 2024. The investment plans and feasibility were generated in accordance with these dates. There have been several reusable vehicles proposed to date, of which only one has been deemed to be successful; the STS Space Shuttle operated by NASA. However, the costs of operating the Space Shuttle were still in excess of the operation costs of a typical expendable rocket, which meant that it was not financially viable. Since the shuttle was retired there has not been a reusable payload delivery vehicle available to the commercial market and although the US military has been developing the sub-scale X-37 for its own uses, it is unlikely this technology would be expanded into the commercial market. This report examines the feasibility of using the latest technology to create a financially viable reusable rocket system. In this report many systems were examined to see if the technology can be improved and developed using new techniques to allow a new generation of reusable launch systems to be developed. Launch philosophy and propulsion were examined to find the most efficient method of reaching orbit. This research investigated a variety of different methods including the traditional vertical rocket take off and a horizontal take-off such as that proposed by Skylon. Orbital operations were examined to find the most effective and efficient methods of accurately manoeuvring in space, this would reduce the amount of fuel that required to be lifted into orbit, increasing the payload fraction. Once the spacecraft has completed its mission the spacecraft would need to re-enter the Earth’s atmosphere before landing. This section of operations was examined in detail as it has many technical challenges and is often considered the most dangerous part of space travel. The way the spacecraft operates was examined to ensure that maintenance and launch preparation could be done efficiently. The structure and the electronics of the vehicle were examined to make sure that they could operate safely over multiple launches. Improvements intechnologywere anticipated toallowthe craft’s mass andcosttobe reduced considerably compared to current generation expendable rockets. All of the possible solutions for each subsystem were listed in the concept register (Appendix I.3) from this the solutions were down selected to find the best option for each subsystem. As concepts are introduced during the report a code is included were possible to link it to the concept register, for example solar panels has a concept register code of [EE06-01]. These solutions for each subsystem were then used in the integration phase to create a conceptual design.
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To examine the viability of such a reusable launch system the possible markets for such a system were examined, this included new markets that were predicted to grow significantly during the 30 year operational life of the launch vehicle concept. These included; space tourism, space station supply and the construction of orbital infrastructure. Space construction projects are likely to form a significant part of the future of space exploration. The analysis of the payloads and markets allowed the business case for the project to be assessed, which in turn allowed for a conclusion on the financial viability of a reusable payload delivery system to be made. For a reusable system to be viable the majority of the vehicle needs to be reusable with minimal maintenance. The system also needs to operate with a high mission frequency across a large number of mission types compared to current expendable rockets. With growing markets and new technologies bringing down the cost of operations, it is likely that the feasibility of a reusable launch system would be high and that it represents a way forward for the space industry. 4. Down Selection and Integration Methodology 4.1. Down Selection methodology To conduct this feasibility study a number of concepts in a range of areas were assessed to find the optimal concept system design. These concepts were then assembled into a reusable rocket concept to demonstrate that they were a fully integrated solution. To effectively evaluate these concepts a systematic process was required that could be applied to all the concepts allowing the optimum solutions to be selected. In general, the concepts went through a two stage process. The initial concept elimination was performed where the concepts that failed to achievea‘threshold value’wereeliminatedfromthedownselection process. The elimination phase was then followed with a down-selection where concepts were compared against each other and the optimal concept in each sub-system area was selected. This phase was then repeated until the optimum solution to a particular sub-systemareawasfound.The‘thresholdvalue’systemwasimplementedtoallow more exotic concepts to be assessed and eliminated rather than down selected, where it may have caused the unnecessary elimination of a more conservative or viable solution. A key element of the down selection process was to only carry out sufficient calculation or research to successfully down select a concept. This was a key differentiator from a more traditional design project and allowed this study to cover a greater range of concept areas (these are listed in Appendix I.3) than a preliminary design review. The areas chosen for evaluation are shown below. For each down selection, concepts were analysed and a numerical ranking was applied for each of the areas. This process is presented throughout the report in each of the sub-system areas. Technical Viability: Anassessmentoftheconcept’stechnicalperformance.Thiscriterionwas used to quantify the performance gain associated with a particular concept. The Technical Viability of a concept was assessed independently from the financial feasibility so as to allow more exotic solutions that may have been difficult to realise to be given sufficient analysis.
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Financial Feasibility: An estimate of the development, manufacture and operational costs of a particular concept. This area assessed the financial feasibility as compared to similar systems to the one being analysed. Development program risk was considered separately as a suitable investment plan may have been able to de-risk certain areas without significantly affecting the financial feasibility.
Development Program Risk: An assessment of the likelihood that significant, unplanned, project costs or delays may be incurred during the development program of a particular concept. An exotic technology would have typically scored poorly in this area and it was often found to be inversely proportional to the technical viability. There were some situations where a lower performance solution was selected because of its lower development risk. Integration: To ensure that the concepts integrated correctly and that the selection of one concept did not adversely affect another sub-system area, an integration criterion was adopted. This, combined with regular group meetings meant that when the final solution was presented it was integrated and accurate. This is discussed in more detail below.
Environmental: For any new engineering project the environmental impact must be considered, for each sub-system the environmental was quantified and ranked. This area did not have a significant impact on the core concepts used in the final system design but was often key in the elimination of some exotic concepts. Reusability: The aim of this project was to specify a reusable payload delivery system. As a result each concept was assessed on its reusability to ensure that it can be maximised wherever possible. 4.2. Integration Integration was a key part of this project and therefore all decisions were made with integration in mind. In the initial stages of the project, problems were assessed and ideas were generated for possible solutions, this was followed by a down selection, as outlined above. During the down selection stage possible solutions were compared until the ideal solution for sub-system was found. During this phase all down selections were made with the rest of the team present, allowing Integration issues to be factored into the decision process. Once the down selection phase was completed the integration phase of the project began. During the Integration phase the majority of the solutions fitted together to form a conceptual design which was to be expected as Integration was considered from the start of the project. The first week of the integration stage contained a meeting in which the first conceptual designs were established. During the following week any integration issues were solved, so that in subsequent weekly meetings the final conceptual design could be realised. The third planned week of integration was found to not be required, this shows that including integration into the earlier decisions was worthwhile, as the
Section 4 - Down Selection and Integration Methodology Page 3 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 concepts fitted together in the final integration stage, and that the subsystem solutions presented in this report are integrated with each other. During the first section of the integration phase, the conceptual design for the rocket stages were developed for the 40,000kg variant of the rocket as it was found to be marginally more suited to the payloads outlined in the Inception Report. Once this had been completed it was decided that a 25,000kg variant could be designed that used the same first stage and a set of upper stages with a high degree of commonality. This added commonality between the designs reduced the development cost and risk associated with designing two systems. The majority of payloads would be able to be mounted and deployed using standard mountings. Once the payload has been mounted on top of the rocket stages it would be covered in a custom faring to reduce aerodynamic drag. The main exception to this is the human payload module; this was designed in the next stage of the Integration phase. It was decided that this human module would only be fitted to the smaller 25,000kg variant of the rocket and that on entry into service an escape system would be required to be compatible with current Human Launch regulations. The human module would be fitted directly to the top of the rocket and would not require a fairing. 4.3. Project Management Summary To make sure that this project was well managed the team members were assigned roles of responsibility. Tim was made responsible for Launch philosophy and propulsion, Matt was made responsible for structures and materials, with support from William. William was made responsible for finances and ground operations, with support from Matt. James was responsible for re-entry and landing. Robert was responsible for in orbit operations and also for the electronics required by the spacecraft. Mukudzei was responsible for investigating possible payloads, markets and human support systems. It was also decided to allocate areas of project responsibility, Tim was the Project Manager, Robert and Mukudzei were System Integration managers, Matt and James were System Architecture Managers and William managed Finance and Risk. This allowed every aspect of the project to be monitored throughout the project and as a result the project remained on schedule throughout. To help manage the project three management tools were used; A Gantt chart to plan the project which can be seen in the appendix I.1 an action list that allowed the group to keep track of current tasks which can be seen in Appendix I.2 and the concept register that allowed the group to follow which ideas were been taken forward and which had been discounted at down selections, this can be found in Appendix I.3.
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5. Presentation of System Concept and Operations 5.1. Presentation of system concept From the analysis conducted in this feasibility study the optimal system concept is presented below. Although the main body of research has been the selection of the most suitable subsystems for a generic reusable launch vehicle, the sub-systems are presented as an integrated reusable rocket system concept. 5.1.1. Launch Philosophy and Propulsion From the analysis of over 20 Launch methodologies and propulsion types the optimum payload delivery system was specified. The conclusion of the analysis was that two variants of reusable rocket would be used. An A-Variant with a 25,000kg payload for the delivery of human payloads and smaller supply missions and a larger B-Variant which with a 40,000kg payload. A number of innovative propulsive solutions and flight profiles were combined to create a reusable launch vehicle that could be very cost effective to operate. For the B-Variant (40,000kg payload) the optimum reusable rocket was found to have three stages and for the A-Variant (25,000kg payload) a two stage rocket was selected. The two variants have an extremely high degree of commonality. The first stage is common for both variants and although sized differently the design of the upper stages are extremely similar, varying only in height and engine number. To maximise reusability each of the stages on both the rocket variants are recovered back to the launch site where minimal overhaul is required before the next launch. The upper stages both incorporate a heat shield for re- entry from high altitudes and all stages are specified to conduct a controlled landing at the launch facility. A more detailed summary of this process is outlined in the concept of operations below. The unique method of recovering the three stages allowed reusability and system up-time to be maximised. The first stage of both variants utilised a pressure compensating engine to maximise atmospheric performance by reducing over-expansion losses. This was achieved through the use of innovative thrust augmentation nozzles that create a variable exhaust geometry through the manipulation of pressure regions within the nozzle via a set of small secondary combustors. This allowed the engine to match the jet expansion to atmospheric conditions during the initial launch phase. This significantly reduced fuel consumptions and take-off mass which improved the payload fraction of the rocket considerably. The reusable rocket used the same engine core on all the stages and this inherent modularity and reuse of engine technology was expected to reduce development costs considerably. The thrust augmentation nozzles are fitted to the first stage and were not required for the upper stages which operate in a vacuum. Apart from the difference in nozzles all the engines are identical in design. Development risk was also minimised by specifying the rocket could be initially operated with traditional engines that lacked thrust augmentation. Although this would have a detrimental effect on the payload capabilities it would de-risk some elements of the
Section 5 - Presentation of System Concept and Operations Page 5 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 engine development. The rocket system has been specified to use Liquid Hydrogen and Liquid Oxygen throughout all the stages. This allowed the fuel system design to be scaled and applied across all the stages, reducing cost as well as using a fuel that can be generated from renewable sources. The launch system used 12 or 13 engines for the A and B Variants respectively. The use of a higher number of smaller engines allowed a good degree of engine out performance for the cargo focussed B-Variant and a very high safety margin for the Human carrying A-Variant. 5.1.2. Materials, Structure and Fuel Systems The spacecraft should utilise structures and materials that are designed for superior mechanical properties and prolonged life expectancy. The fuselage of the reusable parts of the spacecraft was selected to be made from a semi-monocoque structure that incorporates stringers and longerons to increase stiffness and reduce the likelihood of failure through buckling. A large percentage of the craft's volume will be taken up by fuel tanks with fluid that is prone to sloshing. It is therefore important to be able to control the fluid position in order to ensure that the fuel is always entering the fuel lines and to control the crafts centre of gravity. This would allow efficient manoeuvring and positioning of the spacecraft during launch and in orbit. The optimum fuel fluid management solution is the use of a positive expulsion tank that utilises a rubber diaphragm; the rubber diaphragm has an unlimited cycle life and will not degrade with the chosen fuel and oxidiser. Fuel tanks will also only use an alloy thick enough to resist hoop stress in a sphere; any cylindrical tanks should have an organic composite overwrap as the stress will be double. The stringers in the fuselage should have an ‘I’ cross section to provide the most lightweight solution to be stiff enough for parallel bending in two directions, any reinforcing beams and structures should incorporate hollow circular section as this provides the optimal lightweight stiffness solution for bending being applied in all directions. The hollow circular section would also be the optimal solution for any truss structures within the spacecraft, such as those to support the fuel tank, as it would provide the best resistance to buckling of any cross section geometry. To greatly increase the stiffness of structures with only a small mass penalty, sandwich panels should be implemented. The fuselage should incorporate foam core panels and its stringers should use a honeycomb core. Honeycomb cores should also be used for any structures that are sensitive to vibration or require stability due to its vibration damping properties. Table 1 recommends what materials to use for the manufacturing of structures.
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Table 1 – Summary of Materials Selection Structure Material Fuselage skin Titanium Ti-6Al-6V-2SN Annealed Standard Carbon fibre in epoxy matrix (fabric) Fuselage Stringers Sandwich panelled with HRH10 Nomex Aluminium 2090-T83 with a Standard Carbon fibre in Cryogenic tanks epoxy matrix (uni directional) overwrap Support trusses Standard boron fibres in epoxy matrix (uni directional) Reinforcing Openings Titanium Ti-6Al-6V-2SN Annealed Stable/vibration sensitive Graphite Al GA 7-230 components Sandwiched panelled with HRH10 Nomex Non re-usable structures Aluminium Al7075-T6
5.1.3. Structure Sizing The 2 variants have been evaluated for size, based on individual component and substructure geometry. The diameter of all the stages is 8 m, based on the size and number of engines. Table 2 shows both variants length.
Table 2 - Length Estimations Liquid Liquid Oxygen Heat Shield Total (m) Engine 40 T Variant hydrogen Tank Tank Length and/or Height (m) Length (m) (m) Interstage (m) Stage 1 3 19.4 9.7 3.2 35.3 Stage 2 3 12.2 7.2 4 26.4 Stage 3 3 6.5 4.9 4 18.4 Payload (20% of - - - - 16 stages height) 25 T Variant 3 Stage 1 3 15.2 7.9 3.2 29.3 Stage 2 3 7.8 5.8 4 20.6 Payload (20% of - - - - 10 stages height) Adding up the totals estimates that the 40,00kg B-Variant would be 96.1 m in height and the 25,000kg A-Variant would be 65.9 m high. Due to the diameter, the first stage tanks would all be cylindrical, the second stage for the B-Variant would have a cylindrical fuel tank and a spherical oxidiser tank while the A-Variant shall have all spherical tanks in the second stage. The third stage tanks for the B-Variant would all be spherical. Table 3 shows estimates of the tank masses.
Table 3 - Estimated Tank Mass Oxidiser Tank Stage Fuel Volume (m3) Oxidiser Volume (m3) Fuel tank mass (kg) mass (kg) 40 T Stage 1 843 351 4909 2117 40 T Stage 2 477 200 2832 846 40 T Stage 3 144 60 612 254 25 T Stage 1 629 262 369 1613 25 T Stage 2 251 104 1068 442 5.1.4. Re-entry The benefits of a lifting body re-entry vehicle were found to not justify the additional mass and integration issues. The craft will re-enter in the classic ballistic re-entry profile. This would be
Section 5 - Presentation of System Concept and Operations Page 7 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 initiated by a short reverse burn at a specified point in the orbit. This would aim the landing footprint away from the base of operations, towards a large flat expanse near the launch site that would also be used to recover the stages. The craft would enter the atmosphere at an angle of 5 degrees and begin to experience aerodynamic drag, which would quickly heat up the craft, decrease the craft speed and decay the orbit. The heat shield would be a deployable device fitted to the outside of the craft as described in US Patent 7837154B2 (U. Trabandt, 2010). The device would be an open topped cone with an inner radius of 2.5 m (the craft radius), an angle of 70 degrees with the craft and an outer radius of 3.5 metres (1 metre protrusion). The nose cone of the craft would be a dome whose sides match the angle of the heat shield cone, forming a seamless joint. The underlying structure would be sufficiently strong to support the panels as they withstand the forces generated from a 10 degree re-entry. The device fits tightly around the craft when deployed.
The material used for the nose cone would be a single piece of Reinforced Carbon Carbon (RCC), at a weight of 905 Kg. The panel weight depends on the area of panel required (6.86 m2), the thickness of the panel (5 cm) and the density of the material used (SiC 1900Kg/m3). This gave a panel weight of 652 kg. The supporting frame structure would also have a weight proportional to the panel weight to prevent collapse. This was estimated at 40% of the panel weight which is 261 kg. The rest of the capsule was assumed to be covered in Flexible Temperature Reusable Surface Insulation (FRSI) which was used on the space shuttle (NASA, n.d.). This had a weight per area of 1.6 kg/m2, giving the remaining capsule protection a weight of 126 kg. This gives a total Thermal Protection System (TPS) weight of 1.9 Tonnes. Table 4 gives the breakdown of the TPS weight. This is higher than the 1.4 Tonnes of a traditional surface ceramic TPS. However, the added benefits of the deployable heat shield justify the choice. These include increased aerodynamic stability, modularity and vastly increased deceleration. The heat shield patent requires investment, with the traditional ceramic heat shield being the backup choice if the development of the deployable shield was delayed or found to be unfeasible.
Once the craft has decelerated to below Mach 3, drogue parachutes would be deployed to provide more deceleration and an increase in craft stability. The deployable heat shield would then either re-fold or be ejected.. Once the craft reaches less than Mach 1, pilot parachutes would be deployed, almost immediately followed by the 3 main parachutes. The 3 main parachutes would be released in stages, gradually opening to their full capacity. At full capacity, the craft would then decelerate to 10m/s. The weight of the 3 main parachutes is 309 kg. The weight of the pilot and drogue chutes would be small. The chutes would be constructed of Nylon with Kevlar tethers.
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Temperatures Insulation Total weight of Craft area TPS Surface area (m2) expected (K) selected insulation (kg) 20 RCC Nose Cone Common <1650 905 (See section A.2) (44.7 kg/m2) Capsule 35 HRSI Bottom Surface 650 - 1370 322 (8π/9X2.5X5) (9.2 kg/m2) (160°) Capsule Sides 22 LRSI Surface 315 - 650 88 (2x 50°) (5π/9X2.5X5) (4 kg/m2) Capsule Top 22 FRSI Surface Surface < 315 35 (5π/9X2.5X5) (1.6 kg/m2) (100°) 6.86 SIC Panels Deployable >1650 652 (See section A.2) (1900 kg/m3) Support 261 Deployable - < 315 Steel alloy structure (Estimated) 78.5 FRSI Capsule Deployable < 315 126 (2πX2.5X5) (1.6 kg/m2) Total 1,350/1,943 Table 4 - Summary of two TPS systems, with deployable being the primary choice. 5.1.5. Landing
The final landing method selected was to use thrusters. These thrusters would be the same used by the emergency escape system. They would operate when the craft is 51 metres above the ground, as the main chutes are cut away. The thrusters would operate with a combined thrust of 99.1 kN, decelerating the craft by 1 m/s2. This would bring the craft to rest 1m above the ground. The craft can then be lowered onto the ground in a controlled fashion, using the 364 kg of spare fuel. The craft would deploy small legs from behind the heat shield to act as contact points with the ground.
The craft would land in a large open plain away from the base of operations, where it would then be retrieved by ground operations. However, once enough landings have proven the reliability of the system, the craft would then aim to land in a designated zone in the base of operations. This would vastly reduce craft turnaround time, as the craft would no longer need to be retrieved. 5.1.6. Electronics summary All of the key electronics subsystems were examined in this project to find the best way of building the required electronics for the reusable launch system. The system should be based around a communication bus such as SpaceWire. This has been developed for space by The European Space Agency (ESA) and allows high speed Low Voltage Differential Signals (LVDS) communications. As can be seen in Figure 1, there are two SpaceWire buses for the system - one is coloured red and the other is coloured blue. There are two buses to increase the amount of data that can be sent over the system and to add redundancy. The SpaceWire buses connect all the other subsystem modules. Each of these modules will contain Triple Modular Redundancy (TMR), meaning that there are at least 3 of every system.
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Therefore under normal operations, the answers from each of the three systems can be compared and the majority answer taken, reducing errors. If one of the TMR modules suffers a failure the remaining 2 modules can continue to complete the task allowing the system to remain operational. The only exception to this is the On Board Data Handling (OBDH) which will have 5 times redundancy, as this is the critical system that controls the other subsystems and is responsible for controlling the power to other subsystems, turning the power on or off as required and issuing commands over the buses for tasks to be completed. As all modules contain Field Programmable Gate Arrays (FPGA), if there was a complete failure of a critical system, another FPGA in a non critical system could potentially be reprogrammed to take over this task allowing the vehicle to remain semi operational.
Figure 1: Diagram of proposed electronics system
The Attitude Determination and Control System (ADCS) module is responsible for controlling the spacecraft’sorbitandattitude.Todothisit has a series of sensors including an Infra-red horizon sensor to track the Earth’shorizon by tracking the edge of the atmosphere, a GPS unit to receive information about its position and star cameras that allow the spacecraft to calculate its position and attitude using the stars as a reference point. These 3 systems combined allow the spacecraft to constantly know its position and attitude in space. During launch, the system uses an inertial measurement unit, as it is difficult to take accurate measurements from outside the spacecraft during launch. Once the ADCS has used this information to calculate its current position and attitude it can use the actuators to maintain its orbit by correcting for astrodynamic disturbances.
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The actuators can also be used to change the orbit and attitude of the spacecraft if necessary. To make changes to attitude the spacecraft has control moment gyros and small Pulsed Plasma Thrusters (PPT). The thrusters can be used for either large attitude changes or small orbit manoeuvres (that may be required during docking). For large changes in orbital projection, the kick motor is used to change the spacecrafts angular velocity around the Earth. The ADCS system will also control any parachutes and thrusters used during the re-entry and landing process. The power system regulates the power from the batteries or hydrogen fuel cells depending on the size of the module (hydrogen fuels cells will be used in the human module, whereas batteries will be used in the rocket stages). This regulation allows a constant bus voltage to be maintained. The power is then regulated at each subsystem with a current limiting and under voltage protection circuit to protect each subsystem. The spacecraft would contain a VHF band telemetry link with the Earth, which would be used to send critical telemetry and telecommand data, which is required to control the spacecraft. This is designed to be very robust and reliable using a bipolar antenna that does not require pointing to transmit the signal to the Earth. Therefore, the system would have a relatively low data rate and the data would contain redundancy bits to allow for forward error correction to take place. The human module and other payloads that may require a high data rate would include an S-band communication system that would use a phased array antenna to communicate. The phased array removes the need for the system to have a mechanical pointing mechanism, as the communication signal can be pointed electronically. This is done by having a small time delay between the times at which the signal is sent from the apertures in the phased array antenna. The S-band system would be able to communicate directly with the ground or send data via the TDRSS communication relay satellite set up by NASA. The spacecraft would contain a system monitoring system; this would gather data from all areas of the spacecraft, measuring many parameters including the temperature of the heat shields, the stress and strain on the structure, the level of cosmic radiation passing through the spacecraft and air quality in the human module. These additional parameters and the standard telemetry data would be stored on the spacecraft, with the most critical pieces of data being sent back to the Earth over the VHF telemetry link. If operators are interested in a particular piece of data when the spacecraft is in flight they can request it from the spacecraft. When the system returns to Earth all of this data can be downloaded providing information about the areas of the spacecraft that would require maintenance before the next flight. All stages would include landing gear controllers, as it is planned that all modules would return to the space centre and land. All modules apart from the first stage (that does not leave the atmosphere) would require a heat shield for the re-entry process; these would require a control system to deploy the heat shields. The human module would require an active temperature control system to maintain the temperature of the spacecraft, payload and equipment within the desired
Section 5 - Presentation of System Concept and Operations Page 11 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 operating range. To do this the spacecraft will use a combination of pumped fluid loops, heaters and radiators; this would allow the excess heat energy to be irradiated into space and cold areas of the spacecraft to be heated. The human module and the rocket stages would also use passive temperature control techniques such as paints and coatings to control the absorptivity and emissivity of the spacecraft, helping to regulate the temperature of the spacecraft.
All of these systems combined should allow the electronics to be highly reliable, allowing the spacecraft to be monitored and controlled over many missions. The system monitoring should reduce the time required to maintain the spacecraft reducing the time between missions. 5.1.7. Ground Operations The maintenance for the launch vehicle is the main activity during ground operations. As a result the launch vehicle was specified with maintainability in mind using aircraft design techniques. This would allow ground staff to access all of the vehicles systems with relative ease during maintenance. Furthermore, due to the vehicle's modularity, ground crew would be able to quickly swap and remove components if they are deemed unserviceable in the timeframe before the next mission. The maintenance operations must be designed for smooth and efficient processing of the launch system. If a stage or engine takes longer than expected, the sub-system should be taken off the maintenance line for heavy maintenance and a new sub-system would take its place. To aid the ground staff, several system monitoring systems were embedded into the vehicle. These included an integrated vehicle system management system. From this system, the launch system would be able to deliver reports specifying areas in the vehicle that require increased maintenance once the vehicle lands. These would reduce inspection time and increase the turnover rate as maintenance cycles would be reduced and the reoccurring operational costs that plague reusable launch systems would be minimised. 5.1.8. Infrastructure The proposed infrastructure is to be designed and organised around the reusable launch system so as to maximise its capability by reducing the reoccurring operational cost. These costs arise due to the recovery and refurbishment costs after each mission for both of the system variants. The infrastructure must be able to handle and efficiently inspect, maintain and, overhaul all of the vehicle's systems to a safe standard in the minimal amount of time. The infrastructure was specified to follow methods used by the airline and airport industry as they have fully developed the necessary technology and activities to maximise their vehicle's profitability. This is because the launch system generates its revenue by launching customer's payloads into space. When the system is on the ground undergoing maintenance, the increasing refurbishment cost diminishes the systems' capability to generate revenue. To minimise these refurbishment costs the infrastructure is to be designed to turnaround the system through a manufacturing line technique. The maintenance buildings have been specified to have maintenance lines where stages and engines are broken down into common pools where ground staff would prepare sub-systems for
Section 5 - Presentation of System Concept and Operations Page 12 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 servicing. Each of the stages and engines would be partitioned down the maintenance building where they would serviced at stations on a line before moving on to the next station. The launch vehicle is to be assembled and maintained in a horizontal orientation where stages are to be moved on specially designed stage stands that are capable of being moved around on the ground floor. This would allow the launch system to be moved and assembled efficiently by ground staff in preparation for the next mission. 5.1.9. Payload Market Analysis The launch rates per year of different space launch vehicles were collected and analysed for the markets outlined in the Inception Report. These include; Human Cargo to ISS, Human Cargo to a future Space Hotel, automated Space Station supply, Geostationary, LEO and MEO Satellite Deployment, and the construction of Space Infrastructure. The launch values from this analysis were summed and averaged to give the overall launch rate per year for each market. The factors that would affect the number of launches per year were considered. Predictions of when different markets would require a launch system and payload delivery were made. A growth rate of 2% of markets was also considered. The resulting 30 year forecast of the total number of launches per year (total of number of launches made by both A-variant and B-variant) is illustrated in Figure 2
70 60 50 40 30 20 10 0
Totalnumber of missions 2024 2026 2028 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 Years in Service
Figure 2 - Estimated number of missions per year of operation
To allow approximate payload geometries to be calculated the volume and mass capabilities of different launch system were considered. The average volume per unit mass for each system was found to be 0.012676 . This value was very useful in specifying the dimensions of a human payload module. 5.1.10. Human Cargo Capability The Human rated A-variant system has the capability of carrying 25 people to a LEO orbit. However because the system would need to comply with the NASA safety requirements of having a launch escape system it was anticipated that on entry into service the system would only be carrying a maximum of 10 people. It is assumed that once the safety of the system can be
Section 5 - Presentation of System Concept and Operations Page 13 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 demonstrated through repeated launches this requirement could be lifted as human exploration of space is severely restricted by the launch escape system requirement. It was necessary to know the mass fractions of the constituents of the human module. To calculate these, mass fractions of an existing similar space vehicle, DELTA- Advanced Reusable Transport (DART), were considered. For the system with the capability of carrying 10 people, the payload fractions of the subsystems are as follows: Propulsion 31%, ECLSS (Environmental Control and Life Support System) hardware 14%, supplies1 7.3%, heat shield 10%, structures 7.76%, escape system 6.24%, crew 3.2% and cargo 8.5%. The payload fractions include cargo to use up the remaining space. For the system with the capability of carrying 25 people, the payload fractions of the subsystems are as follows: Structures fraction 17%, propulsion 31%, ECLSS hardware 14%, Supplies 18%, Electronics 12% and human cargo 8%. For the 10 passenger Human Cargo launches passengers would occupy the nose cone section of the rocket. The module was specified to have an upper deck and a lower deck. Each deck would accommodate 5 passengers. For the 25 passenger module the seating would resemble the isle and rows of an airliner.
An important requirement for carrying a Human Cargo is an ECLSS. A non-regenerative ECLSS was considered to be the most suitable for the relatively short flight defined in the mission requirements. The optimal technologies for the respective subsystems were found to be the following: The atmosphere control would be provided by a mass spectrometer. The atmosphere revitalisation would be provided by the Lithium hydroxide filters. The temperature and humidity control would be provided by condensing heat exchangers and the fire detection will be provided by the combination of Photoelectric Smoke detectors, Triple Infra Red Flame Detectors and human senses. The fire suppression would be provided by CO2 extinguishers. Water would be stored in tanks and food would be stored in food packages. Hygiene washing would not be needed for missions lasting less than 4 days, but if customers prefer to bath, a sponge in a water cocoon could be used. The shower would utilise approximately 2kg of water. 5.2. Proposed Concepts of Operations This section outlines the concept of operations for the optimum reusable rocket and includes a summary of the key technical features and innovative areas. The decision was made to define two rocket variants with a high degree of commonality. The A-Variant lifting 25,000kg (including Human) payloads to a 200km LEO with two stages and the B-Variant lifting 40,000kg unmanned, payloads. Both rockets share a common first stage and have a very high degree of reusability in the engines and other subsystems.
1 Supplies are consumables that sustain human cargo during the mission duration. These include: oxygen, water and food.
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5.2.1. Launch to a 200km Low Earth Orbit 5.2.1.1. Launch, First Stage Operations and Recovery The rocket would have been assembled and fuelled on the launch pad. The launch command is given and the 9 main engines ignite and reach the correct thrust levels. During this process a whole system electronic health check is conducted and the launch can be aborted on the pad, post- engine ignition. Once the health check is completed the rocket is released from the pad and the launch begins. At launch the rocket provides the highest thrust level at any point throughout the flight but can still tolerate a single engine failure. For the 25,000kg A-Variant up to two engines can be shut down at any point in the assent and not compromise the launch. This increased engine margin is a by-product of the commonality between the two rocket variants and allows a significant margin of safety. As the first stage flight continues the engines are throttled back for the transition through Mach 1 before throttling up to maximise performance. As the end of the firststageburnapproachestheenginesthrottlebackagaintominimise‘g’loadingonthecraft, this is especially important for Human payloadsthatrequirea‘g’levelless than4gtoremain comfortable. At the end of the first stage burn at an altitude of 69km for the A-Variant or 38km for the B-Variant the 1st Stage separates from the upper stages and the 2nd stage ignites its engines. Once a good separation distance from the upper stages has been achieved, the 1st stage executes an 180o turn and reignites its engines, using hypergolic products in the combustion chamber, and executes a deceleration burn. Once this burn is completed the rocket would be in a falling trajectory approximately the same as the ascent profile. Once it has rotated again, the rocket would continue to use its engines during the descent to minimise the increase in speed, and therefore the aerodynamic heating effects on the rocket body. The fuel for this manoeuvre had to be allocated into the rocket mass calculations but represents only a small fraction of the launch mass as the thrust required to manoeuvre as the empty rocket mass is very low compared to the launch conditions. Then as the rocket approaches an area close to the launch site a set of small landing legs are deployed and the stage lands on its tail end ready for preparation for the next flight. 5.2.1.2. Second Stage Operations and Recovery Once separated from the first stage the second stage continues to fly its launch profile. It uses the same engines as the first stage although it lacks the thrust augmentation nozzle as this is not required for operation in a vacuum. Like the first stage, the performance of the engines is varied to prevent over stressing the craft. The flight profile from this point onwards differs depending on the rocket variant. For the A-Variant the burn continues until orbital velocity has been obtained (8.03km/s), from here the second stage and payload coasts until orbit altitude is reached (200km). At this altitude the second stage reignites and corrects the ascent profile into a circular orbit. Once this
Section 5 - Presentation of System Concept and Operations Page 15 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 manoeuvre is completed, separation occurs and the 2nd stage begins its descent. The descent profile is very similar to the B-Variant third stage and is outlined below. For the B-Variant the second stage separates at 4.8km/s and 82km altitude. After separation the stage coasts to allow the third stage to achieve a separation distance. During this time the trajectory of the stage begins to decay and once sufficient difference is achieved between the trajectory of the third stage the second stage re-ignites and performs a partial orbit which allows it to return to the launch site. The second stage of both variants is fitted with an aerodynamic heat shield to resist the heading effects of re-entry. Once re-entry is completed, similar to the first stage, the craft would rotate to an engine down position and reignite, regaining control and allowing a safe landing to be made near the launch site. 5.2.1.3. Third Stage and Payload Delivery The B-Variant utilises a third stage to achieve orbit for a 40,000kg payload and uses the same recovery profile as the A-Variant second stage. Once the payload has been delivered the third stage will be in a 200km orbit. To recover from this position at the appropriate point in one of the following orbits, the engine is reignited and the stage decelerates onto a re-entry trajectory that brings it back to the launch site. Like the A-Variant it is fitted with a heat shield which allows the stage to re-enter the atmosphere. Once the third stage exits the region of significant aerodynamic drag the engine is reignited and the craft is stabilised and landed safely. This process allows the rocket to be efficiently launched and recovered with minimal expense and logisticalimpact.UnliketheSpaceShuttle’sSolidRocketBoosterswhichwererecoveredfrom the sea, these stages recover to an area near the launch site where re-processing can begin almost immediately. 5.2.2. In Orbit Operations Once the 3 stages of the rocket have delivered the payload section to 200km altitude the kick motor and ADCS system in the human module or the payload support module can be used to manoeuvre the spacecraft in space. The kick motor allows the spacecraft to change between orbital altitudes. To make a change to the crafts orbital altitude the angular velocity of the spacecraft’s orbit needs to be changed. To increase the orbit altitude two increases in orbital velocity are required.
The first increase in the spacecraft’s velocity stops the spacecraft orbiting at 200km and starts it travelling along an elliptical transfer orbit. As the spacecraft travels along the elliptical transfer orbit it would be travelling away from the Earth. During this process the Earth’s gravitational field reduces the velocity of the spacecraft. Therefore when the spacecraft reaches apogee (the furthest point from the Earth of an elliptical orbit) it would normally continue along its elliptical orbit and increase its velocity as travels back down to perigee (200km). The spacecraft would stay in this elliptical orbit unless another change in velocity is made. The second change in velocity is
Section 5 - Presentation of System Concept and Operations Page 16 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 normally made when the spacecraft is at apogee (the furthest point from the Earth) which would be the new desired orbital altitude. At this point the second change in velocity ( ) is used to circularise the orbit of the spacecraft allowing it to stay at the new higher orbit altitude. Once the vehicle has reached the higher altitude the spacecraft may deploy the satellite or other payloads that the mission is carrying, completing the delivery mission. At this point the kick motor would be transferred into a grave yard orbit (an orbit that does not have a useful application and is therefore allocated for space junk). Once the payload has completed its mission it may also be placed into a grave yard orbit. Grave yard orbits are required as once spacecraft have travelled to Medium Earth Orbit (MEO) and Geostationary Earth Orbit (GEO) orbit it is impractical for the spacecraft to carry the additional fuel required to bring the payload back to Earth. Once the human spacecraft module has reached its higher orbital altitude it can carry out its mission task, such as docking with the International Space Station. Once this section of the mission is completed the human module would need to return to the Earth. Before the vehicle re- enters it would need to transfer back down to an orbit altitude of 200km. This would require two more changes in the spacecraft velocity. The first slows the spacecraft down, causing it fall into an elliptical orbit. As the spacecraft travels along its elliptical orbit it would lose altitude and gain velocity. When the spacecraft reaches 200km another engine burn (second change in velocity) is required to reduce the spacecraft velocity, circularising the orbit at 200km. Once at 200km the spacecraft can perform a de-orbit burn that reduces the spacecraft velocity, causing it to travel back into the Earth atmosphere using a standard re-entry profile. This would require more fuel. During all of these manoeuvres the spacecraft would be stabilised by the ADCS system. The ADCS controller would take data from the attitude control sensors to calculate the position and orientation of the spacecraft and then uses the actuators to control the attitude of the spacecraft. The orientation of the spacecraft is very important to maintain communication links. More information can be found about the ADCS in the orbital operation and electronics sections of the report. 5.2.3. Human Module Re-entry and Landing At the end of the orbital mission, the craft will begin re-entry. The craft would orient itself with its main engines pointing against the direction of orbit. They would then burn for a specified amount of time, to decelerate the craft to a specified speed. The returning portion of the craft would then separate from the rest of the craft, including the main engines. The rest of the craft would perform further deceleration, to pull away from the returning section and enter a steep re-entry profile to burn up in the atmosphere. The deployable heat shield on the returning portion would unfold, locking in place. The craft would orient itself with its heat shield pointing in the direction of travel. The craft would then begin a ballistic re-entry at an angle of 5° (for early missions). It is essential that the craft re-
Section 5 - Presentation of System Concept and Operations Page 17 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 enters within the re-entry corridor. This is mainly controlled before hitting the atmosphere. The craft would begin to experience extreme drag forces, causing a superheated plasma field to form around the outside of the vehicle, severely interfering with all communications. A detached shockwave would form around the bow and heat shield. This shockwave would trail behind the vehicle, with the entire vehicle staying inside it. The craft would decelerate from near Mach 40 down to Mach 5 using this drag force. The craft would lose most of its altitude, gain downwards velocity and lose horizontal velocity. Once these lower speeds have been reached, communications would come back online, and the craft would begin to cool. The ground tracking station can estimate a landing location. The capsule would then need to begin subsonic deceleration. The drogue parachutes would deploy, decelerating the vehicle, and making the flight trajectory more vertical. Roughly 30 seconds later, the drogue parachutes would be released, and the pilot parachutes would be deployed briefly. The main parachutes would then follow, only being released a small amount. A period of time later, the parachutes would be released slightly more. Another period later, the parachutes would be fully released, providing the maximum deceleration. Once the craft reaches a velocity of 10m/s at 51 m above the ground, the deployable heat shield would either fold back if possible or be ejected. The thrusters which provide the emergency escape system would then fire. The main parachutes would be released, and would lift away from the craft. The thrusters would bring the craft into a controlled descent towards the ground at a rate of 1m/s. Limited guidance would also be provided away from obstacles. When the ground is very near (1m), the thrusters would hold the craft in place briefly, before very gently lowering the craft onto the ground. The thrusters would then cut out. In later missions, a gradually steeper angle of re-entry would be used, up to an angle of 10°. This would give greater landing accuracy, and once the parachutes are deployed, the craft would begin to aim towards landing inside the base of operations. Once landed, the crew must wait inside the capsule until a predefined period has passed. This is to allow the craft to properly cool and the toxic hydrazine, used for manoeuvres in orbit, to disperse. They are then free to either leave the capsule, or wait for the recovery crew to arrive. The recovery crew would pick up the crew and return them to the base if necessary. The craft recovery crew would arrive later to pick up the capsule and return it to the base to be refurbished if required. 5.2.4. Ground Operations Once the stages have been verified to be safe and reliable enough to land near the maintenance hangar, the stages would automatically land on a recovery pad via its on-board flight computers. Once the landed stages have vented all toxic gases, a ground crew would move to each stage where an automated recovery vehicle would lift the stage and move it outside the maintenance
Section 5 - Presentation of System Concept and Operations Page 18 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 building. The recovery vehicle would position the stage onto a stage stand which holds the stage in a horizontal orientation for the rest of the maintenance cycle. From there, the ground crew would purge any remaining gases or fluids remaining in the tanks before it receives its main overhaul. Each stage stand is a cradle structure that holds the stage horizontally allowing the ground crew to manoeuvre each stage around the maintenance floor under electrically motorized drive units. Initially, every stage that comes into the maintenance building would have their engines removed and placed onto individual engine stands when they are taken to the engine maintenance line. The maintenance building is to be divided into four sections; the first three being the three stages from variant A and B and the final section would house all the engines from both variants due to the high commonality between components. The four sub-systems of the vehicle would be overhauled down a maintenance line in their individual sections, where they would be mated with equipment and ground crew at stations. The ground crew would perform the necessary activities to overhaul the components for the next launch. Once the station has finished the set work, the sub-system would be moved down the maintenance line to the next station. Floors and equipment would allow the ground crew to have access around the entire sub-system while the integrated vehicle system monitoring system would report faults which require special attention. Due to the reusability aspect of the vehicle, the components would have a limited life before they reach a certain safe tolerance. With the embedded system monitoring system and scheduling for each stage, the components in each sub-system would be replaced for newly produced components before reaching their maximum tolerances. Organisation of these replacement components for each system would be an efficient task, allowing for a continuous maintenance flow.
Once each sub-system has been validated and verified to a launch worth state, the assembly of variants occurs by mating all the components together in the assembly building via their respective stands. Due to the modularity of the system, stages from other vehicles would be available to connect with each other, allowing for a complete variant to be built if unforeseen circumstances arise. During assembly, the payload would be integrated once the stages have been assembled. If it is a crewed launch, the crew modules would be attached here. Once fully assembled, the ground crew would conduct the relevant launch safety tests before the assembled system is taken to the launch pad via rail. The launch pad would contain the equipment needed to lift the launch vehicle into its final horizontal position for launch by powerful jacks. From here, crew would enter by the launch tower for manned missions. The necessary pre-launch activities would be executed before launch, such as fuelling and testing of launch systems.
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6. System Investment Plan In accordance with the design brief requirement, the investment plan presented here is a strategy for developing the essential technologies and facilities to meet operational goals for the proposed reusable launch system. These technologies have been identified as key development activities that enable enhanced capabilities to the launch system. This would allow the launch system to increase in efficiency, reduce in weight, reduce reoccurring operational cost and overall provide a safe system for customers to operate. The following investments are described below showing their intended use and cost of investment: 6.1.1.1. Spaceport Investment The new spaceport where the launch system would be launched, recovered, refurbished and assembled must be able to maximise the reusable aspect of the system through cost effective processes and facilities. The spaceport would be designed using lessons learned from the airport and aircraft manufacturing industry to reduce turnaround times. Specialised maintenance and assembly buildings would be built to efficiently service used stages and engines before the next mission. The refurbishment processes is an efficient automated system where ground staff would have full access around each stage and engine for delicate procedures. The spaceport would have all the necessary facilities and building requirements to fully serve and support the launch system over its intended project lifetime. The investment would design and build this spaceport before operational service. Tests and validation of the launch vehicle's system before service would be conducted at existing infrastructural sites like Cape Canaveral, FL so the Spaceport could be constructed in parallel to the development program.
Investment Cost: $5,450,000,000 6.1.1.2. Ground Operations Investment For the operations in the ground facilities to increase their refurbishment turnaround cycles, two main automated vehicle monitoring systems called Integrated Vehicle Health Management (IVHM) and Informed Maintenance (IM) are to be implemented into the infrastructure and ground maintenance procedures. These monitoring systems allow the ground crew to receive diagnostic information from the launch vehicle about the vehicle's component state. This allows efficient organisation of equipment and ground staff to be ready in the maintenance building before the vehicle touches down. Furthermore, the vehicle is able to inspect itself allowing a reduced inspection time by ground staff causing turnaround times to reduce significantly.
Investment Cost: $200,000,000 6.1.1.3. Electronics Systems Investment The S band phased array antenna would need to be developed to remove the need for a mechanical antenna mount. A variety of electronics components would need to be developed to keep the system up to date electronically, including thedevelopmentofradiationhardFPGA’s.
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For the ADCS system, the pulse plasma thrusters would need to be developed to allow the spacecraft to make efficient small changes to its attitude. The design also includes a system monitoring system, which would be built into every part of the reusable spacecraft and therefore this needs to be developed with every system separately to make sure that it meets requirements and monitors all the critical data points. The system monitoring system would increase the cost of every section of the spacecraft. Development would also be required to design mountings for the electronics that would securely hold the components through the life of the rocket, as they would be exposed to at least 50 launches before they need to be replaced. Another area that would need large development steps is battery technology and therefore this would require a large investment. Batteries need to be improved to reduce the size and weight so that the charge stored per unit volume is increased (improved charge density). However, other industrial sectors (such as the car industry) are investing in battery technology and therefore the investment that is required from this project could be significantly reduced.
Investment Cost: $210,000,000 6.1.1.4. Engine Development Investment Thrust augmentation nozzles offer higher optimised designed engines allowing for high thrust- to-weight ratios at sea level with the benefits of modularity in the design of the engines. The invested cost would develop commonality between components from existing engines and the bespoke pressure compensating nozzles. The completion of this engine development would be completed before the launch system is operational allowing the vehicle to benefit from this development.
Investment Cost: $2,500,000,000 6.1.1.5. Material for Structure The materials of the launch system must be able to survive in very harsh environments from launch to space. A material that is capable of enduring these conditions is continuously reinforced metal matrix composites that provide excellent mechanical properties including low coefficient of thermal expansion and high thermal conductivity. The material offers far superior capabilities over traditional alloys, allowing the launch system's structural weight to be reduced. By reducing the structure's weight, greater payloads can be lifted once this material is implemented. The investment cost would allow development of this material, providing a more reliable, affordable and maintainable composite.
Investment Cost: $579,500,000 6.1.1.6. Re-entry and Landing During re-entry, the vehicle experiences immense heat that must be contained away from the vehicle itself. The traditional heat shield uses ceramic tiles that absorb this intense heat, which
Section 6 - System Investment Plan Page 21 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 was used by the Space Shuttle. However, the time to inspect and replace these tiles was significantly high. A deployable heat shield is to be developed from a US patent 7,837,154 in which an “umbrella” unfurls itself protecting the launch vehicle from the heat. It further acts as a deceleration structure during the re-entry manoeuvre by creating drag from its body area. During maintenance, a new heat shield would attach onto the vehicle allowing the refurbishment time and cost to be significantly reduced. The investment cost would not only be development cost, but also a significant cost attributed to testing the system through test flights and validation in ground facilities to verify its integrity to protect the launch system.
Investment Cost: $500,000,000 6.1.1.7. Human Cargo - Life Support Equipment As it is predicted that human-rated missions would rapidly increase over the 30 years in service, it is necessary to make improvements to current technologies that make up the optimum ECLSS assembly. The developments that would be done are mentioned below for each subsystem. In the atmosphere control subsystem, the current mass spectrometer has short falls in distinguishing between compounds with the same molecular mass. Lithium hydroxide filters in the atmosphere revitalisation subsystem have a very short life cycle. For a crew of 7 people the filter has to be changed every 11 hours. It would be advantageous to improve the life cycle of the filters to reduce man-hours the crew has to put in replacing the filter. It would save money during space tourism missions as no trained crew members would be needed to change the filters. For the temperature and humidity control subsystem, the coating of current condensing heat exchangers experience degradation due to exposure to contaminants. As a result the failure of the coating causes carryover of liquid instead of total condensation. Improvements need to be made to enhancethecharacteristicsofthecoatingsothatitdoesn’tfailduringthemission. For the fire detection subsystem, the Photoelectric Smoke detectors are susceptible to false alarm due to dust accumulation. Improvements will be made so that smoke detectors will be able to more accurately distinguish between smoke and dust particles.
Investment Cost: $15,000,000
The total investment cost for the technologies and facilities above amounts to $9,424,800,000. This cost can be broken down in Figure 3. This investment cost would be spread over the 10 year development cycle as key technologies are produced enabling the launch system to be optimised for operational use.
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Infrastrucuture Operations Structures Electronics Re-entry Mechanism Launch Engines Medical Equipment
Figure 3 showing the breakdown in investment costs for each sub-system 7. Numerical Analysis and Statement of Feasibility The financial feasibility study was conducted under two methods. The first method was to compare the proposed Reusable Launch System (RLS) against an equivalent Expendable Launch System (ELS). This would validate whether the reusable mechanism of the launch system is a required operational tool by reusing the vehicle repeatedly. The second method was to determine whether the proposed reusable launch system is competitive to the leading launch systems in today's market. Both of these methods were achieved by simulating the launch systems under an operational model set by the market forecast over the course of the project. The launch system does not necessarily have to make a high profit to be deemed feasible, but must show a loss that is not substantially high that would be undesirable to operate.
The development and production cost for both launch systems used cost-estimating relationships (CERs) which were derived using historical data from existing launch systems to predict future launch systems. The total development and total production cost for each variant are shown in Table 5. The proposed reusable and equivalent launch systems are shown in Table 6.
Variant-A Variant-B Reusable vehicle's development Cost $35,860,000,000 Reusable vehicle's production Cost $371,000,000 $509,000,000 Expendable vehicle's development Cost $14,760,000,000 Expendable vehicle's production Cost $182,000,000 $249,500,000 Table 5 showing the total development and production for the reusable and equivalent expendable launch system for both variants
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Stage 1 Stage 2 Stage 3 Variant-A Variant-B Variant-A Variant-B Variant-B Reusable vehicle's $14,600,000,000 $9,400,000,000 $6,360,000,000 $5,500,000,000 development Cost Reusable vehicle's $282,600,000 $89,100,000 $154,000,000 $72,400,000 production Cost Expendable vehicle's $6,140,000,000 $3,870,000,000 $2,550,000,000 $2,200,000,000 development Cost Expendable vehicle's $138,500,000 $43,600,000 $75,500,000 $35,500,000 production Cost Table 6 showing the development and production cost for the reusable and equivalent expendable launch systems
Both launch systems had reoccurring costs for each launch attributed to them. The RLS had constant recovery and refurbishment costs while the ELS had constant production cost for each launch. The RLS is heavily dependent on the reoccurrence of the turnaround cost and overall length of maintenance to its operation. The turnaround cost and time not only dictates the feasibility of the system but determines the number of vehicles needed to satisfy the market demand. A longer maintenance cycle causes an increased fleet size as launch vehicles are effectively grounded when being refurbished, until they are deemed launch worthy. Due to this reliance on the turnaround cost and time, modelling of the turnaround for the RLS was calculated using equations that used statistical data to estimate the turnaround cost and time. This was used to calculate the operational launch cost for each system to determine total cost per year using the forecasted market model. The launch cost for each system was determined through the equation below, which allows the cost to broken down into individually defined costs.
Where the ( ) and ) are the reoccurring costs for the ELS and RLS respectively.
Variant-A with manned missions was separated from the unmanned missions as higher factors of safety were used as an inbuilt safety system for the crew. Figure 4 shows the RLS with equivalent ELS operating these manned missions. The numbers of launches per year for these missions are highlighted in purple with its corresponding secondary axis on the right. From Figure 4, the ELS's reoccurring production cost causes the launch cost to increase with the increase number of missions per year. The RLS's launch cost per year decreases with the higher market demand as the reoccurring operational costs are significantly lower than the equivalent ELS's production cost.
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Crewed missions for a reusable and equivalent expendable launch system 30
6,000
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0 0 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054
Launch Launch (US Cost Dollars (Millions)) Year Expendable Reusable Missions per year Figure 4 showing the launch cost for the manned missions for the reusable and expendable launch systems It is to be noted that during the first four years of operational service, the RLS has a relatively higher launch cost due to the higher pre-launch cost of assembly and integration for the RLS. The RLS benefits from operating the system at higher frequencies as it reduces significantly in launch cost during the year 2034 due to the beginning of space tourism. By taking the launch price as the profits gained to the launch system, the cash flow for both systems can be shown below. Figure 5 shows that both systems show losses during the first 4 years due to the low launch rate. Once space tourism begins, both systems generate significant profits with the RLS producing the highest profit margins. As mentioned above, the RLS has a relatively higher pre-launch cost due to its assembly and integration. This is shown by the higher losses exhibited by the RLS during those 4 years of operational service.
Cash flow for manned missions for the reusable and expendable system 16000
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2,038 2,051 2,030 2,031 2,032 2,033 2,034 2,035 2,036 2,037 2,039 2,040 2,041 2,042 2,043 2,044 2,045 2,046 2,047 2,048 2,049 2,050 2,052 2,053 2,054 -5000 Years Reusable Crewed System Expendable Crewed System Launch (US Cost Dollars (Millions)) Figure 5 showing the cash flow for the reusable and expendable launch system
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The market forecasts were generated using constant launch prices that were used as payment to the launch operator. Over the project lifetime, both systems reduce in their overall launch cost with a 0.85 learning factor due to the increase in efficiency from repeated launch activities. By taking the launch cost with a 20% profit margin, the new cost per seat for the manned mission is shown in Figure 6 with other leading launch system's cost per seat.
Comparing Cost Per Seat for manned launch systems
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Cost/Kg Cost/Kg (USDollars (Millions)) 0 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 years Expendable Reusable Soyuz Space Shuttle Space X Figure 6 showing the cost per seat for the reusable and expendable launch systems compared to other leading competitors This shows that both launch systems (ELS and RLS) are able to meet the competitors cost per seat by the year 2025. It is to be noted that these estimates for the competitors are taken from when this report was created. Once the RLS is in operation, these estimates would be relatively lower than as shown above. However, the graph shows how the cost per seat reduces for the proposed RLS over the project allowing the launch vehicle to meet its competitors as predicted. In terms of cost per kilogram, the proposed RLS is compared against the leading launch systems in Figure 7. It can be seen that when compared to the leading launch systems of today, Variant-A is able to meet its competitors.
15,000 Cost/Kg for various rocket systems
10,000
5,000 Cost/Kg Cost/Kg (USDollars/Kg) 0 Atlas V Ariane 5 Space Shuttle Launch Systems Expendable Falcon 9 Reusable Variant A-Crewed Falcon Heavy
Figure 7 showing the cost per kilogram for the reusable launch system and other launch systems
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Variants A and B for unmanned missions show similar trends for the RLS, however the ELS's cannot gain profit since the launch prices set in the market forecast are too low as payments due to the high production costs for each mission shown by Figure 8 and Figure 9. This verifies that the reusable aspect of the launch system is necessary for the system to be feasible.
Cash flow for variant-A unmanned missions
2000 1500 1000 500 0 -500 -1000 -1500
-2000 Launch Launch (US Cost Dollars (Millions)) -2500 Reusable Expendable
Figure 8 showing the cash for the reusable and expendable launch system for unmanned missions
Cash flow for Variant-B unmanned missions
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-6000 2024 2026 2028 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 Launch Launch (US Cost Dollars (Millions)) Years Reusable Expendable Figure 9 showing the cash flow for the reusable and expendable launch system for variant-B performing unmanned missions
Looking at the reusable launch system's cash flow for all the variants over the project life as shown in the below figure, it can be seen that the first 10 operational years causes loses in the "valley of death" period of the project. However, after year 2035, the reusability aspect of the system is able to produce significant increases in revenue due to the increased launch frequency per year.
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It is to be noted that the cash flow in Figure 10 takes into account development, production, maintenance and component replacement cost which is spread over the operation life time of the project. The system is able to gain a 34% profit margin after the 30 years in service and provide adequate return value for the early investments made into the system.
Cash flow for the reusable launch system for all variants 18000
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US Dollars (Millions) Dollars US 2000
-2000 2024 2026 2028 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054
-6000 Years
Figure 10 showing the cash flow for the reusable launch system over its project lifetime
When the reusable launch system for all variants is presented against the leading competitors, the system is able to meet their launch cost by year 2030 and continue to decrease over the project lifetime as shown in Figure 11.
Launch Cost For all Variants Against Current Market 1200 1000
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(Millions)) 400
200 Launc Cost (US Dollars Dollars (US Cost Launc 0 2024 2027 2030 2033 2036 2039 2042 2045 2048 2051 2054 Years Group 2 Space X ULA
Figure 11 showing the launch cost for the reusable launch vehicle against leading competitors
The financial feasibility study concluded that the reusable mechanism is a necessary component in the launch system as it provides low reoccurring cost when compared to an equivalent expendable launch system. This was further emphasised by the calculation of the Internal Rate of Return
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(IRR) for both launch systems. The reusable launch system had an 11% IRR while the expendable launch system only allowed for 2%. This IRR value is the discount rate for a system that results in a zero net present value, meaning higher is better. This concludes that the reusable launch system is able to provide good economic return over its service life. Furthermore, the proposed reusable launch system is capable of surpassing the current launch system market in terms of launch cost once the market for launch systems becomes saturated. With the development and investment cost added together and spread over the 10-year development cycle, investors are able to recuperate their initial investment into the launch system after year 2035 as it starts to gain significant profit. Therefore, from the financial study posed above, the reusable launch system is a feasible system showing significant benefits against other types of launch systems. 8. Conclusions and Future Work 8.1. Reflection of Project Aims To draw conclusions on the outcome of the study, it is necessary to look back to the project aims outlined in the inception report.
– Two variants of rockets have been proposed with the aim of minimising wasted capacity for a given mission. Variant A is capable of lifting a 25,000 kg human payload to a 200 km Low Earth orbit. Variant B is capable of lifting 40,000 kg non-human payloads into a Low Earth Oribt. The system is presented in section 5. – Each of the markets outlined in the inception report are serviceable through either the A or B variant rocket. – The payload cost has been minimised by maximising the payload fraction, which lowers the cost per kg of payload per launch. The cost per Kg varies with the number of a launches in a year, meaning it is sensitive to turnaround time and demand. It stays lower than all other launch systems throughout most of the life cycle, meaning it is the most accessible launch platform. – The Variant A system is capable of lifting 25 people into a Low Earth Orbit. Initially, a launch escape system will be required for human payloads, limiting the system to 10 people. – Every major part of the system is re-used, with the exception of fairings, kick motors and other small components. The launch stages land vertically near to the launch site, and any manned missions will return to land near to the base of operations. This allows the maximum level of re-usability whilst still reducing mass. The effect of this is a higher structural mass for the stages.
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– The Concept of Operation in section 5.2 outlines the operation of the vehicle for launch, in-orbit operation, re-entry and landing and also ground operations. The infrastructure required is also outlined in section 5.1.8 – The Feasibility Study in section 7 compares the financial side of the project to that of other current and upcoming launch systems and concludes that the system proposed would be more profitable over its lifetime than any other system, based on the expected number of launches. – The optimum number of stages have been found to be 2 for the A-Variant and 3 for the B- Variant. Thrust augmenting nozzles are used on the first stages to maximise the efficiency of the engines as the craft passes through the atmosphere. – The Down Selection Methodology used to initially assess all concepts covered all aspects, including technical viability, financial feasibility, development issues, integration, environmental and degree of reusability. This ensured concepts were not chosen for their technical viability without considering the other factors. – Mass reduction has played a key role in concept development. A lower mass also allows a larger payload mass, making the system more profitable. – The Investment Plan is presented in section 6. All major technologies which would be required for the proposed system are included, with the length and cost of investment required included.
8.2. Outcome The proposed system would be a viable replacement for previous cancelled projects, such as the Space shuttle or the X-33. The A-Variant is capable of taking up to 25 humans into space at a time with a high frequency of launches. This means that it could facilitate a large future expansion in Space tourism. The B-Variant allows for much larger projects, such as construction of infrastructure and large interplanetary craft to be assembled in space. This presents the opportunity for the industrialisation of space and space exploration, with the frequent low-cost launches greatly reducing the cost of such projects. This is all only achievable due to the high re- usability of the system, and the inherent reduction in costs that it offers. 8.3. Future Work This feasibility study was based on a single market model, due to the time constraints of the project. A more detailed market model would be required to fully assess the sensitivity to market fluctuations. Any organisation intending to carry this project forward should perform this analysis, however a reusable launch system is recommended over an expandable one in all cases. This project also contains an investment plan of current and near future technologies. The development of these technologies is vital to realise the concept specified in this report.
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9. Launch Philosophy Vehicle Launch Philosophy is the profile and methodology used to attain an orbit where a payload can be delivered. This section presents the concepts that were selected for analysis, summarises the Down Selection process applied to each concept and presents conclusions on the optimal launch philosophy for a reusable launch system. This section makes references to the launch philosophy appendix and the Group 2 concept register. 9.1. Launch Philosophy Concept Analysis This section introduces the concepts that were selected for analysis and presents an assessment of the potential benefits or design challenges identified during preliminary research. 9.1.1. Vertical Launch Rockets – [LP01] A Vertical launch system was defined as a vehicle that supports its entire mass on the thrust of the engines at launch and continues to do so throughout the assent. All current expendable launch systems launch in this way. If a vertical launch system were selected, it would likely an optimal amalgamation of the concepts below. Vertical launch is a well understood concept and there was significant evidence to suggest that it will be the most efficient way to achieve orbit. 9.1.1.1. Staged Expendable (Benchmark Case) – [LP01-00] In order to quantify and compare the performance of different orbital launch philosophies a benchmark case was required. Typically a traditional expendable rocket optimises its launch performance by manipulating the Tsiolkovsky rocket equation (see Appendix Section A.1). To minimise excess mass, stages are separated when they are empty and for expendable rockets this means that the engines and fuel tank systems are discarded during each flight. The main aim of this report was to assess the feasibility of a reusable launch system and whether it could be a more cost effective system than an expendable rocket and so this research area was used for comparison data. 9.1.1.2. Staged Reusable Rocket – [LP01-01] The logical progression from a staged expandable system was to re-use the stages once they have separated during the launch. For a typical rocket launch the cost of the propellants is about 0.4% of the launch cost (Musk, 2011) and so reusing the flight hardware, which makes up a large portion of launch costs, appeared to be one of the most sensible ways of reducing costs. The Inception Report briefly outlined a SpaceX system that intends to recover the separated stages using a tail landing system (Malik, 2012). This system would carry a weight penalty versus a fully expendable system because some fuel and structural mass would be required to land the spent stages safely. However there may have still been a significant benefit in reusing stages, particularly the complex engine assemblies. A typical two stage Low Earth Orbit (LEO) launch system is comprised of a first stage, a second stage and a payload deployment system. It was likely to be feasible to reuse the first stage due to
Tim West Section 9 - Launch Philosophy Page 31 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 its low separation altitude and it may have been possible to recover the second stage as well. Because of the need to reduce space debris the payload delivery stage would have to be de-orbited and burnt up in the atmosphere or placed into a grave yard orbit if at high altitudes. 9.1.1.3. Staged Reusable with Returning Orbital Component – [LP01-02] The vertical launch concept would have been further optimised by fully reusing the launch system by recovering the orbital component which could renter the atmosphere once the mission is complete. The viability of this concept is based on the complexity of the payload delivery system and the requirement to carry a human cargo. A human cargo would have required a vehicle capable of returning through the atmosphere and this could take the form of a capsule similar to the Apollo Command Module or a craft capable of some degree of atmospheric flight like NASA’sSpaceShuttle.Theoptimalreusabledesignwaslikelytobeacompromise,capableof re-entry for a human cargo or a disposable system for an unmanned launch. 9.1.2. Air Launch – [LP02] Space Tourism Company Virgin Galactic is planning to begin launching payloads into orbit from a launch aircraft flying at Mach 0.8 and 15.24 km altitude in the near future (Virgin Galactic, 2012). It may have been feasible to use a reusable craft to lift the payload delivery system to an intermediate altitude and velocity (known as in rocket analysis - See Appendix section A.1) which may have reduced total launch costs. The analysis of these concepts required a consideration of the technical feasibility of a large launch aircraft and the potential cost savings that it would bring. Air launch concepts would require a significant proportion of the development budget to be spent on the ‘mothership’ and so Financial Feasibility was a key factor in the analysis of these concepts. 9.1.2.1. Captive Subsonic Air Launch – [LP02-01] A captive launch system is one where the spacecraft is carried by a larger craft during the initial stagesofflight,thetwocraftthenseparateandthe‘mothership’returnstoalandingsitewhilethe spacecraft begins a boosted transition to orbital flight. 9.1.2.2. Captive Supersonic Air Launch – [LP02-02] It may be feasible to significantly increase the separation altitude and, more importantly , by carrying the spacecraft to supersonic speeds. There would be significant engineering challenges involved in the separation of two craft at supersonic speeds, for example the SR-71 blackbird suffered several fatal accidents while trying to separate from an unmanned drone at Mach 3.2 (Rich & Janos, 1998). 9.1.3. Space Plane Concepts – [LP03] Space plane concepts are those systems that operate from a runway and utilise aerodynamic devices to take-off like an aircraft before going on to achieve orbit. The propulsion system may still be rocket based. Historically there has been a perceived efficiency gain from using a Spaceplane concept that can take-off like an aircraft and then attain orbit with one or two staged
Tim West Section 9 - Launch Philosophy Page 32 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 propulsion systems. The closest attempt to date has been the Space Shuttle however it launched using a rocket and was considered in this report under concept LP01-02. 9.1.3.1. Jettisonable Atmospheric Flight System – [LP03-01] A similar concept to the air launch system involves launching the spacecraft from a traditional airfield with a propulsion system and lifting devices that would allow a traditional take-off. The craft would then climb to an altitude and velocity similar to the other captive launch systems where the atmospheric flight system would be jettisoned and recovered and the spacecraft would continue to orbit. This concept was considered because it may have mitigated the mass penalties of the large self-contained aircraft required for the captive launch concepts above. 9.1.3.2. Ground Assisted Launch – [LP03-02] A concept being proposed for future airliners is to expand the catapult system used on-board military aircraft carriers. The system in development by the United States Navy is known as EMALS (Electromagnetic Aircraft Launch System) and is capable of accelerating a 45,000kg aircraft to 240km/h in 91 metres (Schweber, 2002). A larger system has been proposed for airliners (Robinson, 2012) and it could have been expanded to a payload launch system providing initial launch thrust, reducing the fuel required for launch and take-off mass. 9.1.3.3. Boosted Ground Launch – [LP03-03] In a similar method to the ground assisted launch above, additional thrust could have been provided through solid rocket boosters similar to those used on the Space Shuttle where they provided 85% of the launch thrust. Like the Space Shuttle, these would then be jettisoned at some point during the assent. 9.1.3.4. Unassisted Ground Launch – [LP03-04] It could have been feasible to launch the entire system from runway straight to orbit and it has been introduced in the Inception Report as the Single Stage to Orbit (SSTO) concept. This concept would have had a very high degree of reusability as all the hardware is recovered at the end the mission and it may have been a way to maximise uptime and hardware usage. 9.1.4. Exotic Concepts – [LP04] Significant consideration was given to exotic launch concepts that could have provided a significant reduction in launch cost and would have been a strong competitor to expendable launch systems. 9.1.4.1. Electro Magnetic Orbital Launch System (EMOLS) – [LP04-01] EMOLS concept proposed a launch system that accelerates a launch vehicle to orbital velocity using electromagnetic force at ground level. The assumption is that the high initial capital costs would allow for significantly reduced launch costs. Further analysis of this concept was required to assess its Financial Feasibility of constructing such a project in 2022.
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9.2. Launch Philosophy Concept Elimination The initial launch philosophy Down Selection assessed and eliminated the concepts that, on further research, appeared unsuited to a reusable launch vehicle. This was because they either required considerable special to type infrastructure which increased project risk to an unacceptable level and reduced Financial Feasibility, or they presented a significant technical challenge that was unlikely to be overcome in the 10 year timeframe laid out in the brief. In accordance with the Down Selection methodology the concept elimination mechanism was used to eliminate the concepts that were below the Threshold Value of 50% of the available score. 9.2.1. Subsonic and Supersonic Captive Air Launch Systems The suggested potential benefit of a captive launch system was that if the craft was accelerated to an initial velocity and altitude, the orbital vehicle would not have to provide initial launch thrust sufficient to lift and accelerate the take-off mass, making it smaller and lighter. This method of launch is currently used by Virgin Galactic to launch a small craft but to carry a useful payload of between 25,000kg and 40,000kg a much larger Carrier Craft would be needed. The Stratolaunch launch group proposed to use an air launch system that could launch a 6,100kg payload into orbit (Belfiore, 2012) which was used a reference for the Down Selection of air launch concepts. Using the concept evaluation methodology and analysis shown in Appendix Section A.2 it was determined that the Stratolaunch concept appears to be flawed, offering a very low payload fraction despite launching at an intermediate altitude. By scaling the system proposed by Stratolaunch to the lower end of the payload range defined for this report an approximate assessment of the carrier aircraft was made and it was found that a 25,000kg payload required a bespoke aircraft with a take-off weight of approximately 3.5 times that of the Airbus A380. The development of this large carrier craft would have a very low Technical Viability and very high project risk which would significantly reduce the financial viability of this concept.
A supersonic launch platform was also deemed an unsuitable (see Appendix Section A.2.2) concept because of the lower Technical Viability. The exotic materials required to build the carrier aircraft, the high fuel fraction required for supersonic flight and the challenges associated with vehicle separation at supersonic speeds mean it may well have been an impossible vehicle to construct. Both of these air launch systems were found to fall below the concept Threshold Value because of the large financial investment required, limited Technical Viability of large aircraft and the very high development project risk of developing two concepts, the rocket and the ‘mothership’ in parallel. 9.2.2. Sea Level Captive Launch Systems It had been proposed that a sea level captive launch system could a feasible alternative to vertical launch systems for Reusable Rockets or SSTO Spaceplanes (Nobuyuki Tomita, 1999). Operating
Tim West Section 9 - Launch Philosophy Page 34 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 in a similar way to captive air launch systems, initial flight velocity is attained using a ‘mothership’(inthecaseof (Nobuyuki Tomita, 1999) a Ground effect vehicle2 or Ekranoplan) which would accelerate the craft to an intermediate velocity where the launch craft would separate and continue to orbit. By not having to accelerate from rest using internal fuel sources, payload fraction may have been improved.
The analysis in Appendix Section A.2.3 found that although the Sea Level Captive Launch was technically more feasible than the air launch concepts, the financial viability and additional project risk of developing two systems meant that it fell below the concept Threshold Value. 9.2.3. Electro Magnetic Orbital Launch System (EMOLS) Building on the concept of Captive Launch systems the Electro Magnetic Orbital Launch System (EMOLS) could be used to accelerate a payload to orbital velocity requiring a minimal launch craft which would reduce payload cost significantly. It had been suggested that there was a significant market for non-rocket ground based launch system (Coopersmith, 2011). The concept being proposed by StarTram (Powell, et al., 2004) was used for analysis. It proposed to erect a levitating EMOLS that would have a muzzle altitude of ~20km which would project the payload at near orbital velocities requiring a minimal propulsion system to transfer into a stable LEO. The analysis for an orbital velocity EMOLS based on the StarTram documentation is presented in the Appendix section A.3.1. It was decided that large infrastructure projects such as the EMOLS system are extremely sensitive to the launch payload market, which were unlikely to grow sufficiently to make this project viable (see the Payloads and Markets section), and therefore the Financial Feasibility of a project of this scale was found to be low. StarTram make significant assumptions about the low cost of the superconducting cables and magnets necessary to construct the system and the technical challenge of building a 20km levitating structure. Additionally the potential projects risks are not considered in the financial report (Powell, et al., 2004) which would likely increase the overall project cost considerably, further reducing the financial viability. Although the potential savings are significant, the low technical feasibility and marginal financial viability meant that the EMOLS concept fell below the concept Threshold Value. 9.2.4. Summary of Launch Philosophy Concept Elimination Three launch concepts were found to rank below the Concept Elimination Threshold. The feasibility scores for each are summarised in Table 7.
2 A ground effect vehicle uses the vortex interaction between a wing and the ground at an altitude of less than one span to generate efficient lifting surfaces with significantly reduced induced drag. The concept is explained more thoroughly in Appendix A.2.3
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Area Air Launch Air Launch Captive Launch EMOLS (0 – 5) where 5 is better Subsonic Supersonic at Sea Level Technical Viability 2 1 3 2 Financial Feasibility 1 1 1 0.5 Development Program Risk 1 1 1 0 Environmental Impact 1 0.5 2 1 Reusability 4 3 3 5 Total (of a possible 25) 9 6.5 10 8.5 Table 7: Concept Elimination Numerical Ranking The two remaining main concepts, Vertical Launch and Space Plane were taken forward for further analysis. 9.3. Down Selection Between Reusable Rockets & Spaceplane The down selection between the reusable rocket and Spaceplane concepts was conducted using the down selection methodology and assessed the relative benefits and drawbacks of a typical reusable rRocket or Spaceplane. A comparison between the relative performances was made using an analysis of typical flight operations and payload mass fraction which were derived from a numerical analysis of the flight profiles and predicted launch masses. There has been a historic belief that with modern materials and propulsion technologies a Spaceplane could be used to deliver payloads more economically than a rocket. The key argument for a Spaceplane was that it had lower infrastructure requirements and a higher system uptime and did not require a complex overhaul between flights, similar to an airliner. The aim was to make multiple launches with the same system making it much more financially feasible than a comparable expendable rocket (Koelle, 2002). The down selection process considered each of the six down election criteria (Technical Viability, Financial Feasibility, Development Program Risk, Integration, Environmental Impact and Reusability) and compared the results for both of the concepts. 9.3.1. Technical Viability Comparison Most single stage to orbit (SSTO) or two stage to orbit (TSTO) Spaceplanes have tended to use atmospheric aerodynamic effects (such as lift and RAM compression) to offset the structural mass penalty of a system that has to carry a very high structural mass fraction to orbit. This fraction would typically include a traditional landing gear system and flight control systems, a heat shield and a large empty tank mass. Some examples of SSTO or TSTO systems include NASA X-33, ESA Saenger, NASP and, the HOTOL/Skylon concept. Most Spaceplane designs have some common elements; an integrated structural heat shield, small wings optimised for the craft empty mass and landing, air breathing engines to lower oxidiser fuel mass requirements and a central cargo bay. The layout of Skylon is shown in Figure 12 as an example:
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Figure 12: A typical Spaceplane layout (SKYLON) (Reaction Engines ltd, 2012) Despite the air breathing engines reducing the requirement for Liquid Oxygen (blue sections in Figure 12) the majority of the interior volume of Skylon was used to store Hydrogen fuel. A traditional expendable rocket uses stages to reduce structural mass in steps during the launch (See Appendix section A.1) but a Spaceplane carries most of this volume to orbit and requires sufficient structural strength to support the empty volume during the high structural loads of re- entry and landing. From the market research conducted in the Inception report, the only mission that was identified as requiring a craft that can re-enter the atmosphere and therefore requires a heat shield and landing systems, was the delivery of a human payload. For this mission some of the excess mass is offset in a Spaceplane because the heat shield that the craft uses for re-entry can also return human payloads. Payload research also found that although a large proportion of annual launches require the capability of launching humans, due to the predicted rise in space tourism, a significant amount of the heavy lifting required for space infrastructure and interplanetary travel require a large one-way craft. A satellite or space infrastructure launch does not require a re-entry component and so the optimal payload delivery from a technical point of view is to deliver the payload to orbit with a minimum mass solution that is likely not recoverable. A Spaceplane is inferior to a rocket in this case because a large fraction of the launch mass must be carried to the payload delivery point which would consume more fuel and reduce the payload fraction significantly. The technical challenge associated with a large, reusable heat shield also has a significant impact on operational costs which was discussed further in the financial feasibility section.
The market research defined a payload requirement of between 25,000kg and 40,000kg to a LEO orbit. This payload range that can easily be achieved with a medium size rocket like the NASA Delta IV or Falcon Heavy system with a payload fraction of between 3% and 5%. Table 47 in Appendix section A.4.3 contains some representative Spaceplane configurations that were derived by scaling the payload, mass fractions, and dimensions from the HOTOL/SKYLON concept to
Tim West Section 9 - Launch Philosophy Page 37 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 the payload range defined in the inception report. The predicted payload fraction for SKYLON is comparable to the equivalent rockets despite the structural mass increase. This is predominantly due to the air breathing engine system which reduces the amount of liquid oxygen propellant that needs to be carried on board, however this system could be incorporated into a staged Reusable Rocket where it becomes feasible to use a more complex and expensive engine technology because it is reused. This is assessed further in the Propulsion Section. From a Spaceplane mass fraction analysis it was found that a payload of 40,000kg makes a Spaceplane unfeasibly large to build. This is because the mass of main structural elements and landing gear becomes non-linear with payload mass for aircraft with a large Maximum Take-off Weight (MTOW). A 25,000kg Spaceplane could be a feasible solution but it is likely that the key technical components could be incorporated into an equivalent reusable rocket meaning that from a technical point of view the Spaceplane does not offer any benefit above a reusable rocket. 9.3.2. Financial Feasibility Comparison As a Spaceplane is an integrated and new approach to payload delivery it will cost approximately 80%3 more to develop, build, test and prove than an equivalent reusable or expendable rocket system. To be more financially feasible than a reusable rocket a Spaceplane would of had to offset its higher development and manufacturing costs by minimising turnaround time and offer higher system availability by minimising the time between flights. Provided that a reusable rocket was recovered economically it was considered unlikely that the inspection, overhaul and preparation costs of a Spaceplane would be significantly lower than a staged reusable rocket. The inherent system modulatory of a staged reusable rocket meant that it was likely that the servicing and overhaul to operate a small fleet of rockets would be cheaper than a fleet of Spaceplanes to deliver the same annual payload. The heat shield of a Spaceplane would have to be made of a non-ablating material that could be used for multiple flights. Maintenance of this large shield is likely to prove very costly as it was on the Space Shuttle and its performance would be critical to the successful operation and survival of the whole system. Inspection of the modular heat shields on the upper stages of a reusable rocket would be simplified considerably by the lower heat shield area and reduce criticality. Although the loss of a stage during re-entry would be a significant cost, it would not disrupt the operations of the rocket fleet because a module could be swapped in. This would also mean that the re-entry components would not require a human safety rating, unlike a Spaceplane, as they would always be unmanned. 9.3.3. Development Program Risk The development program of a Spaceplane was projected to cost significantly more than a reusable rocket. Historically, higher development costs imply an increased project risk. Three key
3 This is discussed in the Operations and Costs section
Tim West Section 9 - Launch Philosophy Page 38 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 areas were identified that show the Spaceplane has significantly higher project risk then a reusable rocket:
As a Spaceplane must be operated as a fully integrated system without the modularity and inherent redundancy of a staged reusable rocket it must be proven to a higher degree than a reusable rocket. A Spaceplane project would have a higher level of project risk as a result. The air breathing propulsion system for a Spaceplane would likely be quite challenging to develop, requiring risk reduction activities which would increase costs. A reusable rocket would be able to operate efficiently using current generation, off the shelf, engines and although a new propulsion system would be developed to further improve the efficiency and keep the system competitive over a predicted 30 year operational life, the development risk would be reduced because the reusable rocket propulsion system was not on the critical path. Expendable Rocketry is a well understood area that has been continuously developed since the 1950s. Significant design experience could be easily incorporated into a reusable rocket whereas a Spaceplane would have to be developed with little or no historical experience which increases risk as well as likely possibly leading to an over designed vehicle. 9.3.4. Integration, Environmental Impact and Reusability Integration of human payloads into a reusable rocket is a simpler task. As discussed above, although a Spaceplane would be able to offset some of the mass penalty of a heat shield and landing systems when carrying a human payload on unmanned missions it presents a significant drawback. When carrying a human payload a reusable rocket would have to incorporate a heat shield and landing systems into the payload mass and is the more efficient solution as the mass penalty is only incurred on the missions that require it. The environmental impact of both systems was found to be minimal compared to current generation of reusable rockets. Both systems offer a similar level of reusability and both would like run on renewable Hydrogen and Oxygen rather than carbon based fuels. The level of reusability is very similar in both the systems, Spaceplane likely reuses slightly more than a reusable rocket which would discard some small elements during launch but there was not a significant difference identified between the two. 9.3.5. Summary of Down Selection between Spaceplane & Reusable Rocket The conclusion of this section (Table 8) was that there are very few areas where a Spaceplane is the superior solution to a reusable rocket. Although it does make some considerable advances over an expendable rocket system as it was unable to provide sufficient benefits and performance over a reusable rocket to overcome the considerable increase in development costs and risk.
Area Spaceplane Staged Reusable Rocket (0 – 5) where 5 is better Technical Viability 2 4 Financial Feasibility 2 4
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Development Program Risk 2.5 5 Environmental Impact 4 4 Reusability 5 4 Total (of a possible 25) 15.5 21 Table 8 Summary of Down selection between Spaceplane & reusable rocket 9.4. Optimum Vertical Rocket Launch Philosophy 9.4.1. Comparison Methodology The Saturn V launch system was chosen as a comparator for the vertical launch concept. Although designed in the 1960s it was quite advanced and achieved a respectable payload fraction of ~4% to LEO (Powell, et al., 2004) which then moved onto the moon. There is a significant amount of data available for comparison and analysis. The analysis began with an assessment of mass fractions scaled from those of the Saturn V and incorporated information from the propulsion section. The scaled mass fractions were then modified as appropriate to include reusable elements like landing systems and heat shields, modern material choices and, predicted propulsion system fuel consumptions. 9.4.2. Staged Reusable Rocket A typical LEO rocket like the Saturn V comprises of two launch stages and a small payload delivery system. It was necessary to determine the optimum number of stages for launch and this was done using an assessment of the rocket equation in Appendix section A.1. Most reference material indicates that a two stage rocket offers the most efficiency, however to optimise efficiency a 3 stage system was selected. Initially a typical 2 stage expendable rocket that lifted was modelled using analysisandtheSaturnV’smassfractionsgivingapredicted launch mass of . The equivalent three stage rocket was predicted to weigh , 13.4% lighter. A three stage model was then adopted for a number of reasons outlined below, including reduced predicted launch mass. 9.4.2.1. Optimum Level of Stage Recovery Once the three stage model was adopted the optimum level of stage recovery was determined. The recover mechanisms were defined as follows. For 1st Stage recovery, at an altitude just above the atmosphere, a deceleration manoeuvre would be used to reduce the stage velocity and begin the descent. Before entering the atmosphere the rocket would initiate a braking burn to prevent excess aerodynamic heating from damaging the system as it travelled through the atmosphere. As 2nd and 3rd stages tend to separate at approximately orbital height and velocity (184 km and 6.84 km/s for Saturn V 2nd Stage) recovery would require a thermal re-entry system similar to that of the Space Shuttle or the Apollo crew module. All stages would land tail first on a prepared area at the launch site. Selection between the stage recovery options was conducted based on the mass increase required to allow the recovery of each stage and the operational cost of refurbishing a stage.
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9.4.2.1.1. Recovery of the 1st Stage 1st Stage separation on the Saturn V occurred at 67km. The structural mass fraction of the Saturn V Hydrogen stages is 8% ( NASA History Office, 2006) but SpaceX is aiming to reduce that to 4% in a new series of rockets due to enter service by 2015 (Musk, 2012). It was estimated that the structural mass fraction derived from the Saturn V would be increased by 6% from 4% to 10% to account for the structural mass required to cope with the stress of a landing system and the landing systems. This percentage increase was justified by an investigation into modern aerospace alloys in the Structures section. From a consideration of the budget it was determined that approximately of fuel would be required to decelerate the craft to a safe landing speed. An iterative process was used to calculate the change in mass and it was found that for a typical Kerosene and Liquid Oxygen 1st Stage the launch mass would increase by 11.1% and the payload fraction would decrease from 7% to 6.20% compared the equivalent Hydrogen and Liquid Oxygen system. The operational costs associated with preparing a stage were calculated and are summarised in the operational cost section. The cost of preparing a stage was found to have a minimal impact on the cost when compared to the hardware costs of a stage. As a result the cost of reusing a stage was deemed as acceptable and was incorporated into the design. 9.4.2.1.2. Recovery of the 2nd and 3rd Stage The 2nd and 3rd Stage would separate above the atmosphere and at high velocity and so would require a re-entry heat shield. Analysis of re-entry systems found that the mass of a heat shield is 10% of the mass re-entering the atmosphere (see section 15). Like the 1st Stage, the structural mass fraction was increased from a baseline value of 4% to 10% to incorporate the increased structural mass and 10% of the empty mass was added to represent the heat shield. Using the same process as for the 1st Stage analysis, the increase in launch mass was found to be 19% higher than the equivalent expendable rocket and 11% heavier than reusing the first stage alone. This was deemed an acceptable increase in launch mass as it is offset relatively easily with a minor increase in fuel load but increased the reusability of the launch system greatly. 9.4.3. Staged Reusable with Returning Orbiter It had been assumed that the third stage would deliver the payload to LEO. For a satellite payload the most efficient rocket launch system is one that deploys the satellite directly because the final mass to orbit is minimised, improving efficiency. A capsule with re-entry capability and landing systems was deemed to be most suited to the delivery of payloads that require the ability to renter the atmosphere and based on the market research conducted in the inception report this was most likely to be a human cargo. A conclusion of the re-entry and payload analysis sections was to use a returning component for human payloads and so this concept was incorporated as a payload option.
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9.4.4. Conclusions on the Optimal Rocket Concept From a numerical analysis of the launch performance, operational cost, market requirements and orbital transfer performance, the optimal reusable payload delivery system launch philosophy was found to be a fully reusable rocket that used three stages to attain a 200km LEO and transferred payloads to the correct orbit from there with an engine incorporated into the payload mass. 9.5. Creation of a Rocket Family From the market research in the Inception Report, a payload range of 25,000kg to 40,000kg was specified. The intention was to refine the market research and select a payload value within range. Market research conducted in the Payload section of this document concluded that the 40,000kg payload was the most financially beneficial to adopt as it could carry 2 large Geostationary communications satellites and propulsion systems (up to 19,000 kg each) in one launch as well as large amounts of Space Infrastructure in the future. The 40,000kg configuration would have likely been considered too large to service LEO Space Stations and manned missions which do not require a large launch mass. The decision was made to propose creating a two rocket family that had an extremely high degree of commonality, allowing both markets to be served while retaining a favourable payload fraction of 4.08% and 6.53% respectively as well as offering whole system reusability. It was concluded that the optimum solution was to use a common first stage which could be used on both a 25,000 kg and 40,000 kg launch system and that the upper stages of each variant would have a very high degree of commonality with only the stage height and number of engines varying based on the and mass requirements of each stage. The two variants were designated Variant A for 25,000 kg and Variant B for 40,000 kg. The number of engines was calculated in the Propulsion section based on the selection of the thrust augmentation nozzle and an outline of the stages is presented below. The predicted mass fractions for the rocket family are shown in Appendix section A.4.4. 10. Propulsion Launch system propulsion was a key design area that was directly linked to launch philosophy and selection of a launch philosophy strongly influenced the selection of propulsion system. In anticipation of this, the propulsion concepts were grouped by the launch philosophy they were most relevant to. 10.1. Propulsion Concept Analysis The concepts below were selected for analysis based on the requirement to begin building a reusable launch craft in 10 years time, as identified in the design brief. The final concept was likely to be an amalgamation of the concepts introduced below and consideration was made on the future development of engine technologies.
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10.1.1. Rocket Technologies – [PR01] The PR01 concepts were most applicable to the reusable rocket (Concept Group LP01) but a Spaceplane would also require a rocket system to attain orbit. These concepts build upon or improve current generation rocket technology and are analysed numerically in the same way as a traditional rocket. 10.1.1.1. Current Rocket Technology – [PR01-00] To assess the suitability and competitiveness of the final concept it is necessary to assess current propulsion systems. It is also possible that a COTS (Commercial off the Shelf) propulsion system is the most financially viable option for this Payload Delivery System. 10.1.1.2. Optimisation of Bi and Tri Propellant Systems – [PR01-01] There may have been benefits in adopting a different fuelling system from the traditional Liquid Oxygen and either Liquid Hydrogen or Kerosene. Many fuels offer a higher specific impulse than kerosene or Hydrogen (see Appendix section A.1) and more investigation was required to assess the viability of these alternative propellants. 10.1.1.3. Thrust Augmentation Nozzles (TAN) – [PR01-02] It has been proposed that by injecting fuel into the supersonic region of a rocket nozzle overexpansion thrust losses could be reduced (Scott Forde, 2006). In general, rocket nozzles are optimised for partial vacuum performance which means that at sea level the exhaust is over- expanded causing oblique shocks at the nozzle exit, reducing engine efficiency which in this report is represented by specific impulse, . TANs aim to use propellant to reduce the effective size of the nozzle at sea level, preventing this thrust loss giving potential thrust to weight ratio improvements of 25% (see Figure 34and Figure 35 in Appendix section B.1). 10.1.1.4. Aerospike Engines – [PR01-03] Previously used on the now cancelled NASA X-33 SSTO concept, Aerospike Engines may have been suited to most reusable payload launch concepts. Like TANs, Aerospike engines improve off design performance (losses due to overexpansion) by using ambient air pressure to match the effective engine nozzle the ambient air conditions (NASA, 2000). The design of an Aerospike engine is considerably different to a traditional bell rocket nozzle and comprises a high number of smaller combustion chambers arranged against a half nozzle. Although this engine technology is less mature than other rocket systems there is the potential for significant fuel burn and mass savings. 10.1.1.5. Dual-Expander or Dual-Throat Engines – [PR01-04] Like TANs or Aerospike engines, dual expander engines use multiple expansion chambers to optimise the performance through a range of altitudes and offer good vacuum and sea level performance. This is achieved by using a set of secondary expansion chambers that use the outer edge of the nozzle as well as a core combustion chamber that allow the engine to minimise overexpansion losses at low altitude by switching between them (Manski, et al., 1997).
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10.1.1.6. Solid Fuel Rockets – [PR01-05] Solid fuel rockets provide thrust by reacting two stored solid propellants. Although the of a solid fuel rocket is less than a liquid rocket system, solid fuels are often used in launch boosters. A solid fuel booster has a simple, robust and cheap design which makes it suitable for a potentially disposable booster system. Solid fuel systems tend to be difficult to throttle, shutdown or reignite although there are some concepts available to allow this to occur safely. 10.1.1.7. Hybrid Rockets – [PR01-06] Hybrid rockets that use a solid fuel that reacts with a gaseous oxidiser are not expected to provide sufficient thrust payload to orbit (N.A. Davydenko, 2007). There are a number of other applications in space or during re-entry where a hybrid rocket is more suitable. 10.1.2. Atmospheric Oxygen Propulsion Systems The propulsion concepts below are likely suited to a system that launches from sea level with some amount of atmospheric flight performance achieved through wings. 10.1.2.1. Air-Breathing Rockets – [PR02-01] An air breathing rocket draws oxygen from the atmosphere during the initial launch phase rather than using an on-board reserve. The most well-known Air-Breathing concept is the Skylon SABRE engine which significantly reduces the fuel requirement and launch mass (Norris, 2011). Despite testing being carried out on a development engine, there was a significant amount of technical and financial risk associated with a wholly new engine system. 10.1.2.2. (Super Sonic) Ram Air Compression Engine - (SC)RAM Jet – [PR02-02] Similar to the air breathing rocket a (SC)RAM jet could provide additional thrust for the atmospheric stages of launch although a rocket system would be required at altitudes >25km. (SC)RAM Jets could offer high specific impulse with low mass and complexity however they are typically used on aircraft where they provide sufficient thrust to allow wings to provide lift. 10.1.3. Non-Rocket Propulsion 10.1.3.1. Pulse Detonation Engines (PDE) – [PR03-01] Pulse detonation engines provide an extremely high with a relatively simple and light weight design. PDEs provide thrust from a repeating series of explosions. Although current subscale prototypes are have demonstrated an of ~1000s (2.5 times that of a Hydrogen Rocket), the Technical Viability of PDEs needed to be assessed further. 10.1.3.2. Nuclear Thermal Rockets – [PR03-02] The most environmentally compatible nuclear launch systems are those that use the heat from a nuclear source to energise a working fluid (often Hydrogen or nitrogen) into an exhaust stream providing thrust. If the working fluid is heated using a heat exchanger then there are no radioactive products in the exhaust. Early Nuclear Thermal Rockets provided a specific impulse three times that of the equivalent dual fuel rocket and could be used as a rocket propulsion system (Schnitzler, et al., 2009).
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10.2. Propulsion Concept Elimination In accordance with the Down Selection methodology the following concepts were eliminated after failing to meet the concept Threshold Value. 10.2.1. Solid Rockets Hybrid and Solid rockets both require a solid fuel element which is built into the structure and combusted to create thrust. The launch philosophy analysis concluded that to maximise reusability and reduce overhaul costs the first stage should be reused. Solid rockets are quite difficult to reignite and reuse and were considered unsuitable as a main propulsion system for the rocket concept although they could be used as a booster. The specific conclusions on the viability of solid rockets are in Appendix section B.2. 10.2.2. Pulse Detonation Engines Pulse detonation engines were found to provide an extremely high and although currently only at the concept phase, the engines are light and simple to construct (Li, et al., 2011). Significant issues were identified during the technical assessment of PDEs which resulted in it being eliminated from the propulsion concept selection. The engines tend to cycle in the 10-40Hz range and would be required to generate ~1.2 the take-off mass of the rocket at launch. This would have generated an extremely high vibration force through the craft. Although this could be mitigated by using multiple engines running in anti-phase the PDE cycle frequency is difficult to control precisely and was likely that the acoustic stresses on the craft would have been considerable although it could be damped for any human cargo (Gong, 2011). The low technical maturity of PDEs was also a key factor in its elimination, current prototypes are only available at laboratory scale and it was deemed unlikely that a full size prototype could be developed in 10-13 years which contributed to a high level of project risk and low financial feasibility. 10.2.3. Nuclear Thermal Rockets The analysis of Nuclear Thermal Rockets (NTRs) came to similar conclusions as the one conducted on PDEs. Although the theoretical benefits of NTRs are quite considerable the high environmental impact and limited technical viability of these engines combined with their expendable design made them unsuitable for a reusable system. These engines are more suited to interplanetary exploration rather than regular cargo transportation to LEO (Klein, 2004). 10.2.4. Air-Breathing Rockets The SABRE engine operated on SKYLON uses atmospheric oxygen to fuel the engine for the initial stages of flight. Although the research on launch philosophy concluded that a SKYLON type space plane is not the optimum launch philosophy for a payload delivery system, the air breathing technology could have been applied to a reusable rocket system. The main source of data for the down selection was the European Space Agency assessment of the SKYLON program conducted in 2011. It outlined the performance of SKYLONs engines and the analysis of the report concluded that the air breathing engine had a low technical viability for a reusable rocket.
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The high mass fraction of the Oxygen capturing engine intakes mean that the engines have a very poor thrust to weight ratio compared to rocket engines. At launch the SABRE engine produces an Engine Thrust to Weight ratio of ~9:1 compared to up to 155:1 for a rocket engine (ESA, 2011), (Huzel & Huang, 1992). This low T/W does not overcome the mass savings of carrying less oxidiser during launch and means that SKYLON has to take-off using wings rather than using the more efficient direct assent profile of a rocket. Another considerable technical challenge identified was the requirement to offset the moment loss at the air intake. A Rocket Engine generates thrust by accelerating a mass flow and if this flow had been decelerated through the intake then extra fuel would have been required to reenergise it. The financial feasibility and development risk of the SABRE engine and future derivatives was deemed acceptable. Scale models of SABRE engine pre-coolers are currently undergoing risk reduction testing and ESA has stated that the concept presented no significant technical issues for space plane operation. However the positive development program progress and financial viability was not sufficient to overcome the low T/W ratio that lead to a low level of technical viability. 10.2.5. SC/RAM Assisted Engines It was suggested that a SC/RAM engine could provide useful thrust augmentation during the launch phase because of the relatively simple and light design. The theoretical range for a RAM then SCRAM jet is 700s-1500s as the Mach increases which is significantly higher than the offered by rocket engines (Heiser & Pratt, 1994). However as with the air-breathing rockets above, the maximum thrust to weight ratio of a SC/RAM jet is approximately 30:1 which is not sufficient effectively to augment a vertical rocket launch. Additionally, the SC/RAM jet can only operate within a comparatively narrow speed range of M1.5 to M6 (where the decreases rapidly) and cannot be used at launch where the highest thrust is required. The other down selection criteria scored similarly to the air breathing rocket. The development risk was deemed lower because a SC/RAM jet is less complex and separated from the critical design path because it would be a thrust augmentation system. However, like air-breathing engines the technical viability was too low for the concept to be continued.
Table 9: Propulsion Concept Elimination Summary
Area Solid Pulse Nuclear Air- SC/RAM (0 – 5) where 5 is better Rockets Detonation Thermal Breathing Engines Engines Rockets Rockets Technical Viability 1.5 2 1 1 1 Financial Feasibility 1.5 1 1 2.5 3 Development Program Risk 3 0.5 1 2.5 3 Environmental Impact 2 2 0 2.5 2.5 Integration 2 1 3 2.5 2.5 Reusability 1 3 0 2.5 2.5 Total (of a possible 30) 11 9.5 6 13.5 14.5
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10.3. Optimum Traditional Rocket Technology This section aimed to select the optimum fuel and engine type for the reusable launch systems outlined in the launch philosophy section. The elimination of the non-rocket engine concepts left down selection between a set of similar concepts that resulted in the optimum rocket technology being selected. 10.3.1. Optimum Bi or Tri Propellant Fuel System Current and previous generation rockets have tended to use Liquid Oxygen as an oxdiser and
either Liquid Hydrogen or rocket grade RP-1 Kerosene as a fuel. The purpose of this section was to assess if a superior fuel existed and which fuels are most appropriate for each of the stages of a reusable rocket. Fuel selection was made from a consideration of the handling and
storage characteristic of the fuel compared to the Specific Impulse benefit over Hydrogen or
Kerosene. Operational was found by using an efficiency coefficient derived from the
comparison of current engines to convert a theoretical into an operational value which allowed the calculation of predicted fuel burn. 10.3.1.1. Benchmark Performance of Hydrogen and Kerosene Specific Impulse (momentum gained per unit fuel) was used to compare the chemical potential of various fuels. Other key areas that were considered included tank volume, fuel safety and ease of handling and fuel cost. Table 10 outlines the performance and relative densities of liquid Hydrogen and kerosene:
Table 10: Summary of the Performance of and RP-1 (Huzel & Huang, 1992) Oxidiser Fuel Isp at Sea Level s Isp in a Vacuum s Density [ Comments 389 455 67.8 Must be kept at cryogenic temperatures RP-1 300 358 820 Cheaper than Hydrogen Table 10 makes the engineering challenge very clear. The optimum fuel is not necessarily the
most energetic but the one that offers the optimal compromise between , tank volume.
10.3.1.2. Alternative Fuel to and RP-1 An alternative fuel would have had to provide a significant performance increase or much lower cost to be deemed a better solution than the relatively benign RP-1 or highly energetic Hydrogen. From analysis of other fuels (data presented in Table 48 in Appendix section B.2) Liquid Hydrogen was considered the optimal solution. It offered a very high specific impulse without the
drawbacks of more exotic fuels (Huzel & Huang, 1992). Some fuels offer a higher but often incorporate a solid like beryllium making storage and combustion difficult or have significant handing issues like hydrazine which is hypergolic (combusts on contact with an oxidiser). The
fuel that came the closest was and which offers an higher than Kerosene with a high density and required cryogenic rather than hyper-cryogenic storage which reduces fuel tank mass as less insulation is required. When the analysis had been concluded, Hydrogen was selected. This was because it can be generated from renewable sources and the impact on vehicle
Tim West Section 10 - Propulsion Page 47 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 mass from the increase in tank insulation (including embrittlement treatments) and volume is minimal - it is was found to be the lowest mass solution. It could also be used as an engine coolant to increase engine life reusability. 10.3.2. Pressure Compensating Engines A pressure compensating engine overcomes overexpansion losses by varying the nozzle geometry as the rocket climbs. The specialist engine is only required for the first stage of the rocket which has to operate through the atmosphere. Three types were considered in this report and they all offered similar benefits, allowing near vacuum to be maintained from launch for a minimal mass penalty. As the form and function of these engines was quite similar, the analysis focused on Thrust-to-Weight Ratio. 10.3.2.1. Aerospike Engines From research conducted into the Aerospike engine, a maximum sea level Thrust-to-engine- weight of can be achieved (Hall, 2011). This is a significant increase on some current generation engines but it comes at the expense of a complex, monolithic engine design (see Figure 36 in Appendix Section B.1.2). A similar engine was successfully tested during the X-33 program but was not selected for this reusable rocket because of perceived difficulties in fitting the engine and incorporating it into multiple stages. 10.3.2.2. Dual -Expander Engines Dual-Expander engines were found to give a poorer thrust-to-weight ratio of (Manski, et al., 1997) but offered a more traditional and modular design than the Aerospike engine. 10.3.2.3. Thrust Augmentation Nozzles (TAN) TAN were found to provide an excellent sea level Thrust-to-Weight-Ratio of up to 155 (Scott Forde, 2006) and offer the same traditional, modular design of the Dual-Expander engine. The semi-traditional design allowed the engines to be mounted in different sized groups on the main stages and could be easily supplemented with Commercial off-the Shelf engines if the development program fell behind that of the main rocket, reducing project risk. The individual engine was also deemed to be more resilient to damage from malfunction than the combined nozzle of the Aerospike engine, improving safety. The same engine core could be used on all three of the stages, the same engine with a vacuum nozzle for the 2nd and 3rd Stage and with a TAN nozzle on the 1st stage. 10.3.2.4. Summary of Pressure Compensating Engines The three engine types were assumed to cost a similar amount to develop and test and the down selection was made on a consideration of performance, application to multiple stages and safety in the event of a failure. Table 11 summarises the down selection.
Table 11: Summary of Pressure Compensating Engines Area Aerospike Dual-Expander Thrust Augmentation (0 – 5) where 5 is better Engines Engines Nozzles Technical Viability 3 2 5
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Financial Feasibility 4 2.5 3 Development Program Risk 3 2.5 4 Environmental Impact 2.5 2.5 2.5 Integration 2 2.5 4 Reusability 2.5 2.5 2.5 Total (of a possible 30) 17 14.5 21 10.3.3. Number and Size of Engines As thrust augmentation nozzles were selected for their higher performance and modular design, it was sensible to assess whether one engine type could be used for the whole rocket. The number of engines varies from rocket to rocket, Russian designs typically use a high number of smaller engines and United States designs tend towards a lower number of higher performance engines. It was decided to use 9 engines like the SpaceX Falcon 9 system. This allowed the same engine to be used on all the rocket stages with a good degree of agreement between the required thrust and the thrust available from the engines. Table 12 summarises the data and that on average there is good agreement between the thrust requirement and the common engine type.
Table 12: Thrust requirement leading to Number of engines Stage Thrust Required Engines Required Number of Engines A-Variant - - - Stage 1A 15 MN 5.12 9 Stage 2A 3.41 MN 1.16 3 B-Variant - - - Stage 1B 23.4 MN 8 9 Stage 2B 2.79 MN 2.8 3 Stage 3B 2.50 MN 0.85 1 For the A Variant rocket which is expected to carry human payloads, a significant degree of redundancy was added to improve system safety. In operation the engines would be shut down or throttled back to avoid over stressing the launch craft or crew at Mach 1 and at the end of the stage burn. The mass penalty incurred from carrying the extra engines in the A - Variant was not significant and was partially reflected in the 2.45% reduction in payload fraction in A-Variant compared to B-Variant. This loss is more than compensated by the very high degree of commonality between the variants, both in terms of the stage design and the engine sizing. From the thrust requirement generated from the analysis above an approximate engine size was defined from comparison with other engine types and used in the reusable rocket CAD and figures. 10.3.4. Calculation Process for Fuel and Stage Mass Fuel mass estimates were calculated using a analysis which calculated the mass of fuel required based on the empty weight of the craft, the velocity requirement and the performance of the engine. A structural mass estimate was defined by a mass fraction which related the structural mass to the fuel requirement. This calculation was preformed over a number of iterations until a stable value of the stage mass was achieved. The mass fractions were initially derived from the Saturn V (S. C. Krausse, 1969) and then modified to better reflect more modern attempts at a
Tim West Section 10 - Propulsion Page 49 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 reusable rocket (Musk, 2012). The validity of these figures was confirmed by comparing the predicted rocket masses for the test cases against the actual rocket take off masses and a good agreement was found between the two. 10.3.5. Engine Investment Plan The thrust augmentation nozzle was selected for its superior Thrust to Weight ratio and relatively conservative design. The development of the engine was be broken down into two areas, the development of the engine core which houses the combustion chamber and supply systems which was incorporated into all the engines on the rocket and the thrust augmentation nozzle which was required only on the first stage. The first element, the design of the common components, was deemed to be similar to other engine types like the SpaceX Merlin 1. The second element, the development of the nozzles, would be more challenging. Combustion instability, the combustion of fuel in the nozzle rather than the combustion chamber, was a significant challenge in the Apollo program (Yang & Anderson, 1995). The combustion of Hydrogen in all the stages increases complexity but has been achieved before (e.g. Rocketdyne RS-68) and so the additional investment required for the engine, above traditional development costs, was assumed to be approximately 20% in accordance with the increase in cost of a whole reusable rocket system above an expendable rocket (see Figure 37 in Appendix section B.5). Assuming the development baseline costs of 120,000 man years for a reusable engine of this size (Pratt & Whitney Rocketdyne, 1980), increased by 20% suggests that the total development cost for the engine system would be 10,000 man years or $2.5 Billion. This would need to be invested over a typical development cycle over the course of the next 10 years.
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11. Structures and Fuel Systems
The spacecraft structure is going to be broken down into distinct main areas, each with their own individual needs. These areas are the cryogenic fuel tanks, the outer shell of the spacecraft, internal beams/supports and panels. Each area will have the potential technologies evaluated and down selected to provide the most feasible option. The main drive for material selection is to select high performing, lightweight materials that will not fail with repeating use. Therefore cost is not the driving factor in choice, although it is still a viable stake during the selection process. The cryogenic fuel tanks will take up a large proportion of the vehicles volume; the emphasis here is to make them safe for the handling of very low temperature fuels. Careful material selection is needed due to the potential safety implications associated with many materials at cryogenic temperatures and at high pressure. The materials must also be very strong to support the large volume and vastly changing loading conditions as fuel is used up. The spacecrafts outer shell also requires material with a very good strength to weight ratio and high stiffness. It has additional strains of being exposed to the harsh environments associated with space and atmospheric entry/exit, so must be able to cope with this kind of cycling. The internal load bearing structure needs to be exceptionally strong and stiff as it will support the vehicle and the payload during manoeuvres. A material with a very high strength to weight ratio is needed and structure optimised for the many loading conditions and fatigue cycles is essential. The beams and supports used for this structure should have their geometry engineered to maximise stiffness according to the loading directions. 11.1. Exterior Skin of the Fuselage [ST01] There are a number of options available for the exterior fuselage structure, Table 13 summarises the advantages and disadvantages of them. Truss structure – A truss structure consists of welded bars positioned in appropriate directions to resist loads, they can be bulky but are strong. Monocoque – A monocoque design uses a stressed skin to support most of loads experience, it is both strong and lightweight, but can suffer from buckling Semi monocoque – A semi monocoque design uses a stressed skin to support load, but is reinforced with stringers and frames to aid in resisting other types of loading, eg buckling.
Table 13 - Summary of Fuselage Structures Structure Advantages Disadvantages Monocoque Stressed skin supports most of the loads Low tolerance for loads and Saves weight and maximises volume deformations on the surface Aerodynamic Buckles easily More suited to smaller applications Semi-monocoque Stressed skin supports most of load Slight space penalty compared to Stringers give a large increase to the monocoque approach skins stiffness under torsion and bending Stringers and frames add additional loads costs
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Frames prevent buckling Truss structure Truss can be designed to effectively Lack of streamlined shape resist stress from any direction – hard to Requires lots of welding buckle Can be bulky
11.2. Exterior Skin Down selection Table 14 investigates the initial feasibility
Table 14 - Fuselage Down Selection Technical Financial Development Integration Environmental Reusability Total Viability feasibility Program risk Truss 1 5 1 1 NA 2 10 structure Monocoque 4 4 2 5 NA 3 18 Semi 5 3 4 4 NA 5 21 monocoque
The truss structure has therefore been eliminated from the concepts as it scores less than 50%. It is an old fashioned method, traditionally used on small aircraft and not applicable for large, reusable spacecraft. 11.3. Monocoque versus Semi-monocoque In order to determine the optimal choice, a structural idealisation has been evaluated of an example spacecraft in its launch configuration, in order to determine if one option has a significant mass penalty over the other. It has been modelled as a cantilevered cylinder with a uniformly distributed mass as shown in Figure 13. Table 15 shows the example problem requirements and data and shows the arrangement.
퐸퐼 Lateral Rigidity : 푓푛푎푡 3 Equation 1 푀퐵퐿
퐴퐸 Axial Rigidity: 푓푛푎푡 Equation 2 푀퐵퐿
Where E is the modulus of elasticity, A is the cross sectional
area, I is the area moment of inertia of the cross section MB is the distributed mass and L is the length.
Figure 13 - Cylindrical Model
Table 15 - Cylinder Data Geometry of cylinder Length : 10 m Diameter : 2 m Spacecraft uniformly distributed Mass : 2000 kg Load Factors Axial : 2.5 (steady state) + 4 (transient) = 6.5 Lateral : 3
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Rigidity First axial frequency > 25 Hz First lateral (bending) frequency > 10 Hz Material Properties (aluminium 7075) Young’sModulus 71 GPa Poisson’sratio 0.33 Density 2800 kg/m3 Ultimate tensile strength 524 MPa Yield strength 448 MPa
Moncoque Axial Rigidity From equation 2 and the data provided, the cross sectional area can be determined to be 2.817x10-3 m2. Thus, the required skin thickness can be determined;
√
√
Lateral Rigidity From equation 1, it can be found that the required area moment of inertia for the cylinder is 8.982x10-
3 m4. The thickness can be determined to be 2.86x10-3m ( ). 3 This shows that he lateral rigidity requirement is more critical, the minimum cross sectional area needed in order to satisfy this is
Equivalent and applied axial loads The spacecraft weight must be multiplied by the load factors in order to obtain the expected loads. Table 16 shows the applied loads, taking into account the load factors as presented in Table 15.
Table 16 - Applied Loads Load Type Load Factor Limit Load Axial 6.5 127530 N Lateral 3 58860 N Bending moment 3 294300 Nm
The combined axial, lateral and bending loads on a thin walled cylinder can quickly be evaluated using the equivalent axial load equation (Wiley J.Larson, 2004);
Equation 3
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Using equation 3, where P is the axial load, M is the bending moment and R is the radius, the combined loading can be estimated to be 716130 N or a stress of 39.79 MPa.
Tensile Strength To size the cylinder to meet the requirements for tensile strength, the equation for stress is used;
Substituting in the known values gives the required thickness to be 2.175x10-4 m.
Buckling Stability It is important to analyse the buckling stability, as this is a common mode of failure associated with monocoque structures. As previously deduced, the cylinder must withstand compressive loads of up to 716130 N. The following equations for elastic cylinder buckling are needed;
Equation 4
Equation 5
Equation 6 Substituting in known values to equations 5 and 6 (thickness being 2.86x10-3m from lateral rigidity requirement) gives to be 1.169, to be 0.379. Thus the critical buckling stress from equation 6 can be calculated as 46.18 MPa. Finally, the load is deduced;
This is above the estimated combined loading and is a good indication that the cylinder will not buckle. Hence for the example spacecraft of 2000 kg uniformly distributed mass, its required monocoque fuselage would weigh;
Semi-monocoque
With a diameter of 2 m the cylinder has a circumference of 2 m. By incorporating 12 stringers and 11 frames, the angle between each stringer will be 30o. Figure 14 shows the arrangement.
Figure 14 - Stringer Arrangement
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An assumption has to be made when evaluating this approach with respect to the material thickness, as the acoustic equations are bulky and complex. It has been assumed that a standard gauge thickness of 1.250x10-3m (1.25 mm) will be suitable to handle the acoustic loads encountered(Wiley J. Larson, 2004)).
Stiffness As found earlier, in order to meet the requirement for axial frequency, a skin thickness of is needed, showing that the standard gauge thickness is sufficient. For the bending stiffness, the required area moment of inertia was calculated to be 8.982x10-3 m4. This would be covered by a combination of input from the skin and the stringers. Equations 7 and 8 are used for this;
Equation 7
∑ Equation 8
Where is moment of inertia of the skin, r is the cylinder radius, t is the wall thickness, is the moment of inertia of the stringers, is the moment of inertia about the stringers centre of mass, A is the cross section area of each stringer and d is the distance of each stringers centre from the cylinders centre. From equation 7, the skins input can be found to be 3.927x10-3 m4. The remainder to be supplied by the stringers is 5.055 x10-3 m4. Table 17 shows moment of inertia based on the area of the stringers.
Table 17 - Moments of Inertia Distance from Neutral axis (m) Number of stringers at distance ∑ 0 2 0 0.5 4 A x 1 0.866 4 A x 3 1 2 A x 2 Total = 6 x A Using equation 8 and Table 17, it can be deduced that the cross sectional area of each stringer must be 8.425x10-4 m2. The total combined area from skin and stringers is 1.80x10-2m2. Stability of skin panels The elastic stress required to buckle a curved skin panel is given by;
( ) Equation 9
Where t is the thickness, E is young’s modulus, b is panel width, v is poisons ratio and k is a geometric constant (55)(Wiley J. Larson, 2004). Substituting the values into equation 9 gives a
value of 20.54 MPa . However, earlier the estimated loading was calculated as 39.79 MPa, indicating that this option would fail. The margin of safety is negative;
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The skin thickness must be increased, but the stringer area can decrease at the same time to keep the overall cross section area at 1.80x10-2m2. Table 18 shows an iterative approach to a safe solution.
Table 18 - Iterations for Skin Thickiness Iteration Thickness Skin area I skin I Stringer Total MS (mm) (MPa) (m2x10-3) (m4x1 stringers area area (N 0-3) (m4x10- (m2x10- (m2x10- x10^5) 3) 4) 2) 1 1.25 20.54 7.85 3.93 5.06 8.43 1.8 3.69 -0.48 2 1.35 22.18 8.48 4.24 4.74 7.90 1.8 3.98 -0.44 3 1.45 25.59 9.11 4.56 4.43 7.38 1.8 4.60 -0.36 4 1.55 29.24 9.74 4.87 4.11 6.85 1.8 5.25 -0.27 5 1.65 33.13 10.37 5.18 3.80 6.33 1.8 5.95 -0.17 6 1.75 37.27 11.00 5.50 3.48 5.81 1.8 6.70 -0.07 7 1.85 41.65 11.62 5.81 3.17 5.28 1.8 7.48 0.04 8 1.95 42.28 12.25 6.13 2.86 4.76 1.8 8.31 0.16 It can be seen from Table 18 that a skin thickness of 1.85mm is adequate with stringer areas of 5.28x10-4m2. The total weight of this would be;
( )
There is no significant weight penalty between the two options, however both approaches would also be subjected to lateral loading. This would add additional design criteria as the structures would buckle with lower applied forces. The semi monocoque design would have a lower weight penalty due to its additional stiffness from stringers. The benefits of the semi monocoque become greater as the vehicle becomes larger. The semi monocoque design is the more feasible option because of this and with no significant weight penalty has been selected for the fuselage. 11.4. Fuel tank structure [ST02] Spacecraft fuel tanks are widely made up of alloy skins with frames to provide support and stability. The structures incorporate stringers and stabilizing frames, the benefits of which have been explored in 11.3. Typically, pressure fed systems are used for rockets which deliver low to moderate levels of thrust and total impulse. These simplified systems reduce the overall weight of the propulsion system, increase reliability and is an attractive option for spacecraft application (Larson, 1999). Liquid storage systems can be complex as they need to ensure that liquid is expelled from the tank instead of gas or vapour. There are a few ways of doing this which are summarised in Table 19; Artificial gravity induced by spin – The spacecraft is spun to generate centrifugal forces and thus keeping the liquid at the outlet pipe Positive Expulsion Tank – There are a number of positive expulsion devices that mechanically separate the liquid propellant from the pressurising gas Surface Tension Device – Rely on exploiting the surface tension forces of the propellant to keep the outlet covered with liquid
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Table 19 - Summary of Fuel Tank Structures Structure Advantages Disadvantages Artificial No pressurising gas or subsystem Large tanks become more difficult to gravity required spin induced by Large tanks spinning could spin the craft spin in the opposite direction Fuel slosh must be controlled Positive Many different options available Generally contains more valves and Expulsion Brilliant life cycle mechanisms than other options Tank Stops gas from flowing into the propellant piping under multi directional acceleration or spinning. Can accurately control the liquids centre of gravity Surface No pressurising gas or subsystem Best suited to low acceleration Tension required environments Device Very effective in micro gravity Inertia can overcome surface tension conditions Less effective as tank size increases Passively manage propellants
11.5. Fuel Tank Structure Down Selection Table 20 shows the down selection for different tank types.
Table 20 - Down selection for fuel tank structures Technical Financial Development Integration Environmental Reusability Total Viability feasibility Program risk Artificial gravity 2 2 3 1 N/A 10 18 induced by spin Positive Expulsion 10 6 9 8 N/A 9 42 Tank Surface Tension 6 8 5 8 N/A 8 35 Device The artificially induced gravity option is unfeasible as it would alter the dynamics of the spacecraft due to the large fuel tank sizes and would not work as effectively under combined acceleration situations. The surface tension device is attractive, but would not be feasible for large tanks due to the reduced effectiveness by having large surface areas and high accelerative forces upon launch. Therefore the positive expulsion system has been selected as it overcomes these challenges and works well under a variety of conditions. There are a number of options with regards to a positive expulsion system; these can be seen in Table 21
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Table 21 - Positive Expulsion Mechanisms Option for tank Advantages Disadvantages Rolling Low cost Internal welds need inspecting diaphragm Low mass Low Metal diaphragm High volumetric efficiency High mass No ullage volume High cost No sliding seals Only optimises for special envelopes High Rubber Extensive database Not compatible with certain propellants diaphragm Low Great expulsion efficiency Unlimited cycle life Proven design Metal bellows Great compatibility High mass Hermetically sealed High cost No sliding seals Limited in cycling capabilities Low volumetric efficiency Piston Adapts easily to growth High cost Low Low volumetric efficiency Possible failure of sliding seals Critical tolerances
Out of the above devices, a number of them have safety and service life issues, which render them hazardous for the scope of a reusable vehicle. Both the metal and rubber diaphragms however show great performance with regards to cycling and service life and would be a viable option for use. The rubber diaphragm however is both cheaper and lighter than the metal diaphragm, so offers the most benefits for the reusable systems, although the metal diaphragm would need to be implemented if the fuel type being used is degenerate to rubber material to avoid failure. 11.6. Mass and Shape Estimations To estimate the mass and determine the optimal geometric shape, basic equations for sphere and cylinder geometry have been used along with the following equations for pressure vessel stresses;
Equation 10
Equation 11
Equation 12
Where is hoop stress, is longitudinal stress, r is the radius, t is the tank thickness and P is tank pressure. Table 22 shows the volume of alloy material needed and its corresponding mass for the construction of different volume tanks. It also shows geometric values for sizing (Aluminium 2090-T3 used for calculations data available in Appendix C.3). Table 22 is for spherical tanks.
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Table 22 - Mass Estimation of Alloy Material for Spherical Tank Volume (m3) Radius (m) Surface area Tank thickness Alloy volume Mass of alloy of alloy (m2) (mm) required (m3) (kg) 100 2.879 104.188 1.577 0.164 425.500 200 3.628 165.388 1.987 0.329 851.000 300 4.153 216.720 2.274 0.493 1276.500 600 5.232 344.021 2.865 0.986 2553.000 800 5.759 416.752 3.154 1.314 3404.000 1000 6.204 483.598 3.397 1.643 4255.000 Table 22 considers cylindrical tanks for when radial distance is limited in the vehicle. It assumes the use of half-sphere end caps.
Table 23 - Mass Estimation of Alloy Material for Cylindrical Tanks Volume Radius Length Surface area Tank thickness Alloy volume Mass of (m3) (m) (m) of alloy (m2) (mm) required (m3) alloy (kg) 100 2.5 1.760 106.180 2.738 0.291 474.504 200 2.5 6.853 186.180 2.738 0.510 1041.837 300 2.5 22.131 266.180 2.738 0.729 1609.170 600 2.5 32.317 506.180 2.738 1.386 3311.170 800 2.5 37.410 666.180 2.738 1.824 4445.837 1000 2.5 47.596 826.180 2.738 2.262 5580.504 Table 23 looks at the effects of reducing the radius and increasing the length of the cylinder to retain the same volume
Table 24 - Mass Estimation of Alloy Material for Cylindrical Tanks Volume Radius Length Surface area Tank thickness Alloy volume Mass of (m3) (m) (m) of alloy (m2) (mm) required (m3) alloy (kg) 100 2 5.291 116.755 2.190 0.356 519.805 200 2 13.249 216.755 2.190 0.475 1087.138 300 2 37.122 316.755 2.190 0.694 1654.471 600 2 53.038 616.755 2.190 1.351 3356.471 800 2 60.955 816.755 2.190 1.789 4491.138 1000 2 76.911 1016.755 2.190 2.227 5625.805 Table 22, Table 23 and Table 24 show that as the length increases and the radius decreases, the weight of the tank gets greater for a given volume. It shows that the spherical option (if feasible) will give the lowest mass and that a cylinder with a larger radius will have a lower mass than a cylinder of equal volume, but a smaller radius and greater length. There are 2 reasons for this. Firstly, the sphere has the lowest surface area to volume ratio. Appendix C.1 shows that as the number of faces a volume has increases, the lower the surface area to volume ratio. Secondly, from equations 10 and 11, the cylinder and sphere will fail due to hoop stress, however the hoop stress for a sphere is half that of a cylinder, allowing for thinner materials to resist the same forces when using a spherical shape. These estimates only take into consideration the main structural alloy material, but provide a good guide to how the geometry of the vessel should be designed. If the spacecraft were to be designed, consideration for propellant management systems, internal baffles and plates, exterior skins and supports must also be considered.
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11.7. Internal Stiffeners and Support Structures Some parts of the spacecraft would be subjected to greater loads, such as supporting the payload during manoeuvres, supporting fuel tanks and bracing the spacecraft during high acceleration. These loads would be predictable during vehicle design and simulation and hence additional support structures such as trusses and beams can be orientated efficiently to support loads. The use of composite materials with specific fibre orientations could be adopted as multidirectional strength is not as significant as other structures due to the specific orientations and loading directions. Some beam applications will still be subjected to multi-directional loading however. A beams resistance to bending (its stiffness) is related to a beams second moment of area. Eq 13 shows the second moment of area and eq 14 shows the parallel axis theorem, which shows that the further the material is from the neutral axis, the greater the second moment of area and hence the greater the stiffness. Appendix C.2 uses the parallel axis theorem and lists the second moment of area for some common beam cross sections.
∫ Equation 13 Equation 14 From the parallel axis theorem and C.2, beam stiffness is greater when more volume is further away from the neutral axis and hence more efficient, where efficiency means a beam will deflect less for the same overall cross sectional area. From this and Appendix C.2, it can be deduced that for beams with bending in only one direction, the T shape is optimal, the I beam cross section is most efficient when there is bending in 2 directions parallel to each other (eg up and down). The square cross section is most suitable when there is bending in two or more directions adjacent to each other. However, when there is bending in multiple directions not adjacent or parallel, the circular cross section should be chosen. It can also be seen that making a shape hollow would allow greater outside dimensions and further increase beam stiffness, however the thickness must be adequate to support the stresses and torsion forces imposed. Table 25 suggests what geometry beams to use for a variety of applications.
Table 25 - Structures using Beams Structure Beam geometry Selection criteria I beams are best for resisting bending in parallel directions, Semi-monocoque multiple stringers with this cross section around the fuselage I beam stringers will reduce the effects of lateral forces and increase the buckling stability. Compressive members would be at risk of buckling, a circular Payload and fuel profile will reduce the chances of this. Other profiles are more tank support Hollow circular prone to buckling in non ideal loading situations due to trusses limited axis of symmetry. Reinforcing As the profile opening changes, the loading will change also openings in producing many different loading orientations fuselage (eg Hollow circular doors on cargo bay)
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11.8. Panels Panelsoftenincorporatestructuresdescribedas‘sandwiches’. These consist of a lightweight core that is resistant to shear which is bonded strongly to outer face sheets which resist in plane bending, tension and compression. The driving force behind this is to increase a structures second moment of area and hence increase its stiffness and resistance to bending without adding any significant mass. The sandwich approach essentiallyhelpsthestructureactmorelikean‘I’beam. Figure 15 shows how beneficial this approach can be, equations 13 and 14 have been used to calculate second moment of area, a skin density of 2900 kg/m3 and core density of 80 kg/m3.
Figure 15 - Sandwich Panel Comparison There are two different families of core available for use in the construction of sandwich panel, these are detailed below and evaluated in Table 26;
Foam core – An isotropic and homogenous foam of varying density, It offers the same stress bearing capabilities in all directions Honeycomb core – This is a structure itself used for a core, it is made up from a nested array hexagon shapes that form closed cells.
Table 26- Sandwich panel core types Family of core Advantages Disadvantages Foam Has the same properties in all directions Although some foams are stronger Large surface area for adhesive bonding than common honeycomb materials, they are not as great as the metal or composite honeycomb cores Honeycomb Lighter weight than foam core High manufacturing costs Hollow cells offer excellent sound and Higher material costs vibration damping Not able to retain fasteners very Can be made from composites layered effectively increase directional efficiency The honeycomb core is the strongest and lightest option when the loading directions are known and not in many directions. It could give highly significant weight minimisation by using a composite skin with fibres aligned in the directions of loading and should also be used in cases where vibration suppression is necessary. The foam core is most suited to applications where loading is multidirectional and of vary magnitudes; it can cope better with multiple directions of loading.
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Appendices C.4, C.6 and C.7shows properties for a range of cores, along with a description of each type. 12. Materials 12.1. Introduction to Materials Spacecraft structures must be designed with the ability to endure many environments from their manufacture to mission end. With regards to a reusable vehicle, the mission end is ultimately the vehicles termination after its working life is completed, or after a pre-designated number of mission hours. This section aims to look at the requirements of the different structures and select the most appropriate material to for their fabrication. Table 27 lists some typical requirements for structures during the different phases of the vehicles lifetime, for a reusable vehicle, these will repeat.
Table 27 - Typical Structure Requirements over a reusable launch vehicles lifetime Lifetime Phase Source of Requirements Manufacture and Stresses and strains applied during the manufacturing process assembly Handling fixtures or container reactions Handling and Reactions to mechanical lifting equipment Transportation Changing and/or different environments associated with different land types, sea and air Testing Vibration and acoustic testing Test fixture reaction loads Prelaunch Mechanical handling during preparation sequences and pre-launch checks Environmental factors (e.g wind against an empty and unpressurised tank) Launch and climb to Steady state booster accelerative forces ascent Vibration and noise Pyrotechnic shock from jettisoning hardware and explosive bolts Thermal stresses from environments Vertical launch – high unidirectional (lateral) forces Horizontal launch – high thermal stresses due to longer time in atmosphere, forces are more balanced bi-directionally than vertical launch Re-entry and landing Aerodynamic heating from atmosphere Transient wind and landing loads In designing a spacecraft structure, optional materials, types of structures and methods of construction all need to be considered. Typical spacecraft so far are made up from a mixture of metallic and non- metallic materials. The metallic materials in general are homogenous and isotropic having constant properties throughout their composition and the same properties regardless of direction. The non metallic materials are often made from a combination of materials and often do not exhibit isotropic or homogenous properties. There are a number of issues to investigate when selecting an appropriate material to use; these will vary depending on what part of the spacecraft is being looked at. The areas to be considered are;
Material Strength Corrosion resistance Material Stiffness Ease of fabrication Density Material availability
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Fracture toughness Ease of fabrication and manufacture Ductility Electrical conductivity Cost Melting point/max operating temperatures Thermal conductivity Performance at low temperatures (eg cryogenic for fuel tanks) Thermal expansion
There are a number of materials that have proved highly effective for space craft design and have been used commonly in the construction of spacecraft structures, however the most extensively used materials are alloys of aluminium. Aluminium alloys exhibit great strength to weight ratios, they are readily available, low in cost and easy to machine. Compared to steel used in many other construction industries, aluminium alloys have greater strength to weight ratios and a similar stiffness to weight ratios. However, if a structure requires denser or harder properties, steel, titanium or titanium alloys are normally chosen. Table 28 shows the advantages and disadvantages of some of the most commonly used materials in the aerospace industry;
Table 28 - Advantages and Disadvantages of common Aerospace materials Material Advantages Disadvantages Aluminium High strength to weight ratio Low strength to volume ratio Ductile Low hardness Very easy to machine Has a high coefficient of thermal Low density; efficient in compression expansion Available in sheets, plates, forgings and castings Steel High strength High density – not so efficient for Has a wide range of obtainable strengths, stability hardness’s and ductility via different Many are hard to machine treatments Magnetic Available in sheets, plates, forgings and castings Heat High strength to volume ratio High density – not so efficient for resistant Strength is retained at high temperatures stability alloys Ductile Not as hard as some steels
Magnesium Low density – highly efficient for stability Susceptible to corrosion Low strength vs volume Titanium High strength versus weight Hard to machine Low coefficient of thermal expansion Poor fracture toughness if solution treated and aged Beryllium High stiffness vs density Low ductility and fracture toughness Low short transverse properties Toxic Composites Can be customised for high stiffness, high Expensive for low production volumes strengths and extremely low/negligible Strength dependant on workmanship, coefficients of thermal expansion individual proof testing often required Low density Laminated composites are not as well Good performance in tension performing under compressive forces Brittle and can be hard to attach
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12.2. Advanced Materials There are also some advanced composites and alloys that offer additional attractive benefits, they are not used as much as the materials detailed in Table 28 as the technology to produce some of them is still immature, meaning they are not as readily available and more expensive to acquire. However the additional costs could be worth the investment, as if they increase the vehicles reusability, this could create great costs saving by increasing the number of missions the vehicle can achieve. It is also worth noting that these materials will become cheaper as the demand for them increases and the market matures. For a full description and examples of these materials refer to the inception report. 12.3. Material Selection for Structures This section uses the structures requirements to select the most suitable materials for use. For a complete list of all the selected and researched materials properties, see Appendix C.3. Cost estimates are available for widely used materials only. 12.3.1. Fuselage As the fuselage will be subjected to multi-axial loading, harsh exterior environments and must be strong to avoid penetration Titanium Ti-6Al-6V-2SN annealed has been selected. It is denser and more expensive than an aluminium alloy, but offers a greater strength to weight ratio and has very good corrosion resistance properties.
Youngs Yield Melting Density Hardness Cost Material Modulus Strength point (g/cm3) (MPa) Estimate($/lb) (GPa) (MPa) (K) Titanium Ti- 1900- 6Al-6V-2SN 110.3 1350 4.54 367 7.85 1922 Annealed In order to further increase the fuselage stiffness, sandwich panels should be incorporated using a foam core and the above material as a skin. The foam core has been selected to be a dense form of Polymethacrylimide, which provides the highest shear and compressive strength of the foam cores.
Range Max Cost Shear Shear Compressive of operating (relative Foam Strength Modulus Strength density temperature to (MPa) (MPa) (MPa) (g/cm3) (K) foams) Polymethacrylimide 0.03-0.3 0.8-7.5 19-290 0.8-16 413 High
12.3.2. Cryogenic Fuel tanks The cryogenic fuel tanks need a material that can handle very low temperatures without significant hindrance of material properties. The most effective materials for this are aluminium-lithium types which can handle the cryogenic temperatures while still outperforming traditional alloys of aluminium, although they are more expensive. The alloy selected is Aluminium 2090-T83.
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Youngs Yield Melting Density Hardness Material Modulus Strength point Cost Estimate (g/cm3) (MPa) (GPa) (MPa) (K) 3 – 5 > Aluminium conventional 76 520 2.49 176 833-923 2090-T83 alloy(Scanlan, n.d.) Additionally, a polymer matrix composite overwrap will function to reduce the weight of a fuel tank. It can be used to provide the extra strength required for a cylinder compared to a sphere. This would mean the fuel tank hemi-spherical ends would not have to be manufactured with a thicker alloy to compensate for the double hoop stress in the cylinder section, as the composite would support it and reduce the overall weight. The composite selected is uni-directional carbon fibre embedded in an epoxy matrix, with the fibres aligned in the direction of the cylinders hoop stress.
Youngs Yield Strength Yield Strength Cost Density Material Modulus adjacent to parallel to Estimate (g/cm3) (GPa) fibres (MPa) fibres (MPa) ($/lb) Standard Carbon fibre in 135 50 1500 1.6 7.85 epoxy matrix (uni directional)
12.3.3. Support Structures Fuselage stringers Stringers would take most of the loads when their axis of symmetry is parallel to loading direction. If it is not exactly parallel, then the load would be spread more evenly between adjacent stringers with an off axis loading direction. Due to the main loads being parallel to axis of symmetry a standard carbon fibre layup has been chosen to be an outer skin of a honeycomb sandwich panel that make up the beams, as some uni-directional strength must be sacrificed for off axis loading.
Youngs Yield Strength Yield Strength Cost Density Material Modulus adjacent to fibres parallel to fibres Estimate (g/cm3) (GPa) (MPa) (MPa) ($/lb) Standard 7.85 Carbon fibre in 70 600 600 1.6 (Warren, epoxy matrix 2010) (fabric)
The honeycomb core has been chosen to be an Aramid core (HRH10 Nomex) due to its superior stabilized compression strength and good shear properties.
Stabilised Shear Strength Shear Strength Cell Size Density Material Compressive L direction W direction (mm) (g/cm3) Strength (MPa) (MPa) (MPa) HRH10 Nomex 3 15 3.5 1.9 0.144
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Payload and fuel tank support trusses The truss structures would be orientated to take tensile and compressive forces, the main risk of failure is due to buckling, so a high compressive strength in all directions is important for truss members under compression. Organic matrix composites have been selected for this, for tensile members, unidirectional carbon fibre in an epoxy matrix has been selected for its great tensile strength and low density. For compressive members a uni directional boron composite with an epoxy matrix has been selected for its superior compressive strength.
Youngs Ultimate Compressive Ultimate Compressive Density Material Modulus Strength adjacent to Strength parallel to (g/cm3) (GPa) fibres (MPa) fibres (MPa) Standard boron fibres in epoxy matrix (uni 200 280 2800 2 directional) Reinforcing openings in fuselage (eg doors on cargo bay) Titanium Ti-6Al-6V-2SN Annealed, the same material to be used for the fuselage offers the best strength to weight ration with isotropic and homogenous properties, making it suited to multi axial loading conditions.
Stable Antenna and Vibration Sensitive Parts. A Gr/Al discontinuously reinforced metal matrix composite has been selected for this application, this composite shows good compressive strength and moderate stiffness. More importantly, it has a very low coefficient of thermal expansion which will maintain components positions; it also has excellent electrical conductivity for signalling and transmission. Any mounting structures for these parts should be a sandwich panel design with a honeycomb core due to its vibration damping properties, the honeycomb core should be HRH10 Nomex, which is detailed above.
Material Density Youngs Ultimate Ultimate Coefficient Electrical (g/cm3) modulus Tensile Compressive of thermal Resistivity (GPa) Strength Strength expansion (m-ohm- (MPa) (MPa) (10-6/K) cm) Graphite Al GA 2.45 88.7 76.8 202.6 6.5-9.5 6.89 7-230 Non-Reusable’s For non reusable parts that will be not used, but still require a good mechanical properties, a traditional aluminium alloy has been selected as cost has a bigger influence on the decision. The inter stage connectors are a typical example of non-reusable, but high demanding components. The material selected is Aluminium Al7075-T6.
Youngs Yield Strength Density Hardness Melting Cost Material Modulus (MPa) (g/cm3) (MPa) point (K) ($/lb) (GPa) Aluminium 71.7 503 2.81 175 908 0.85-1 Al7075-T6
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12.4. Investment Plan Discontinuous metal matrix composites have shown great use for spacecraft structures due to their great mechanical properties and in particular their low coefficients of thermal expansion and high electrical and thermal conductivity, making them particularly useful for multipurpose structures. However, the continuous metal matrix composites uses are limited, despite their superior mechanical properties and ability to perform at high temperatures. They have the potential to outperform conventional alloys used for aerospace structures and offer reduced structural weight as a benefit and are generally accepted as the ultimate in terms of mechanical properties and commercial potential (CMT, n.d.). However, use is not widespread due to the problems associated with the ease of manufacturing, inspection, scale up and cost. Typically they can cost an order of magnitude greater than conventional aluminium alloys (Panel on Small Spacecraft Technology, 1994). Investment into the manufacturing processes would make them more affordable, readily available, reliable and repairable, it would also trigger more data to be collected on their performance in the field, as currently their use has been very limited and this inadequate information discourages its use. Figure 16 shows some of the mechanical advantages continuous metal matrix composites have over conventional metal alloys. If they became more viable, they would serve as a great replacement for many crucial structures, mainly the fuselage skin and its stringers and long runs. Aluminium metal matrix composites alone can offer mass reductions of up to 60% and strength and stiffness increases of up to 200% when compared to Figure 16 - Continuous Metal Matrix Composites conventional alloys (CMT, n.d.).
Currently there are a number of ways of manufacturing these materials, these include; diffusion bonding, electroforming, vapour deposition, rolling, extrusion, plasma spray, pneumatic impaction simultaneous growth of the reinforcing and matrix materials. Investment money should be used to improve these processes and making them more efficient and widespread, investment should also be made into processes for rapid testing to build a reliable database and further encourage their use. 12.5. Current Financial Estimate of Implementation The main structural material for the fuselage is estimated to cost 7.85 $/lb. The continuous metal matrix composites are up to an order of magnitude (10) greater than that of traditional alloys, but offer mass reductions of up to 60% and strength and stiffness increases of up to 200% when compared to conventional alloys(CMT, n.d.). Assuming the use of metal matrix composites for component
Matthew Chapman Section 12 - Materials Page 67 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 fabrication, while maintaining the original component weight, it can be said the components life will be 200% longer. This would mean that the can be estimated as (3.33) times greater than conventional alloys. Metal matrix composites will not be viable for this project until the costs comes down. The minimum costs for viability would be when the magnitude of the
ratio is reduced to less than or equal to alloy version;
Where X is the magnitude of additional cost compared to the alloy. Solving the equation shows that in order to be viable for the project the cost of metal matrix composites cannot be greater than 3 times the cost of the alloy ( $23.55/lb). The required reduction in price is $54.95/lb, but investment to make this happen has currently not happened, this is due to the fact that most other projects do not use the expensive titanium alloy chosen for this project, which itself is multiple times more costly than aluminium alloys. In order to become viable for replacing aluminium alloys, the cost must come down significantly. 12.6. Investment Analysis An initial estimate of a reduction in cost of $1/lb for every £10 million has been taken. Adopting the same process as above but for aluminium alloy (using the data in the materials tables in Appendix C.3 to see that titanium is 1.7 times stronger than aluminium for a given mass and costs $1/lb), it can be seen that the cost/lb would need to reduce to $5/lb to compete with aluminium alloys. This would represent an investment cost of $735 million. To compete with the price of the titanium alloy for this project, the investment would be $549.5 million. Table 29 shows the effects of replacing the alloy, but minimising mass instead of prolonging component lifetime for both variants of rocket, in order to increase the payload capacity.
Table 29 - Effects of Replacing Alloy Rocket Type Structural Total weight Payload Weight of Weight New Mass fraction (kg) mass alloy (50% reduction payload (%) fraction of of using mass (%) structural composite fraction mass) (kg) (kg) (%) 40 T variant 12.97 956000 4.18 61996.6 24798.64 6.78 25 T variant 9.23 612000 4.08 28243.8 11297.52 5.93
Looking at the extra payload, it can be seen that the launch would yield an additional $12.4 million profit for the 40 T variant and $5.65 million for the 25 T variant assuming an average cost of $5000 / kg launch cost. The investment to replace the titanium alloy would pay for itself in 44 launches of the 40 T variant or 98 launches of the 25 T variant.
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13. In Orbit Operations Mission orbits A reusable spacecraft would need to operate in low Earth orbits. The lowest orbit that the spacecraft would be able to maintain would be 200km altitude above sea level - to achieve this a tangential velocity of 7.784km/s would be required (Wertz & Larson, 1999). The aim is that the spacecraft would be able to deliver the maximum possible mass of payload to 200km. Above this altitude the payload mass would have to be reduced to allow for the extra fuel to be carried, as more fuel would be required to reach high altitude orbits. The maximum altitude that the spacecraft containing humans would need to be able to operate at is 1,000km. Above this altitude the Van Allen radiation belts around the Earth (caused by charged particles getting caught in the Earth’s magnetic field) would cause too much damage to the spacecraft and payload to warrant long exposures or frequent visits to higher altitudes (Sellers, et al., 2005). To achieve a circular orbit at 1000km the spacecraft would need to be travelling at 7.35km/s. The South Atlantic anomaly is caused by the Earth’s magnetic centre being offset from the Earth’s physical centre and therefore the Van Allen belt comes down to an altitude as low as 200km over the South Atlantic (Sellers, et al., 2005). As the spacecraft orbits around the Earth it would have to travel through the South Atlantic Anomaly. This would cause radiation damage to the spacecraft and would increase the likelihood of failures. Therefore the spacecraft should be built to withstand the space environment.
13.1.1. Launch to Orbit There are 3 ways of getting the payload to the required orbital altitude. This section will look at the 3 options and determine which is most suitable for a payload delivery system. 13.1.1.1. Direct insertion into the desired orbit- [OP01] The spacecraft can directly insert into an orbit of the correct altitude from the launch site, however this requires a large amount of fuel to overcome the force of gravity. Once the spacecraft reaches its orbital altitude it needs to have a vertical velocity of zero and a very high tangential velocity (an angular velocity around the Earth). Once the orbit has been achieved, small changes in angular velocity can be used to change the spacecraft’s centripetal force, in turn changing the orbital altitude. For an orbit to be maintained, the centripetal force must equal the gravitational pull of the Earth. Therefore achieving a low attitude orbit is much more efficient than achieving a higher orbital altitude as the higher the altitude the greater horizontal force needed to be applied during the launch process. Once an orbit has been achieved a transfer can be used to change the orbital altitude or change the inclination of the orbit if required.
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13.1.1.2. Orbit transfer for the entire spacecraft - [OP02] A more efficient method is to launch the spacecraft into a low altitude orbit such as 200km and then use a transfer orbit to reach a higher orbit. A variety of different transfers can be used to change orbital altitude. The most efficient method of changing orbit altitude is a Hohmann transfer. This transfer requires much less fuel to reach the orbit which reduces the weight of the spacecraft at launch and therefore this reduces the operating costs of the spacecraft making it more profitable (Bate, et al., 1971). However, there is a small time penalty for transferring - direct insertion would be quicker. This method is suitable for human payloads operating below 1000km. 13.1.1.3. Orbit transfer for the payload and support module only [OP03] This method of using an orbital transfer can be made even more efficient by only boosting the payload to the higher orbit using a kick motor to transfer the payload rather than the spacecraft (Wiesel, 2010). This has the added advantage that the main part of the spacecraft could then re- enter into the Earth’satmosphereifitisnolongrequiredbythemission;thisinturnwouldreduce the time between launches. The spacecraft would need to be able to incorporate payloads with boost/kick motors if the spacecraft is going to be able to launch payloads to GEO (35880km) and other orbits above a1000km as the spacecraft would not be able to operate above this altitude. 13.1.1.4. Hohmann Transfers The Hohmann transfer is one of the most efficient methods of transferring between two orbits. As a small change in the spacecraft orbital velocity can have a large effect over on its orbital altitude. The orbital velocity of the spacecraft in a circular orbit and the radius of its orbit are directly
linked: therefore:
To transfer between orbits an elliptical orbit is used, as the spacecraft travels along the elliptical orbit, the gravitation pull of the earth would change the spacecraft’s velocity. If the spacecraft is gaining altitude the earth gravitational field would slow the spacecraft until it reaches the maximum altitude (apogee). When the spacecraft reaches the apogee of the transfer orbit (the furthest point away from the earth) it velocity is less than that for a circular orbit of this altitude therefore the spacecraft would be pulled back towards the earth by gravity. As the altitude decreases, the velocity would increase due to the pull of the earth’s gravitational field, causing potential energy to be transferred to kinetic energy. To circularise the orbit at the high altitude a second increase in velocity is required. Figure 17 shows the velocity during a Hohmann transfer.
Figure 17 Graph to shows the change in velocity during a Hohmann transfer
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13.1.1.5. Efficient transfers Figure 18 below shows the most efficient launch and transfer profile in green and inefficient launch and transfer profiles in red. It shows that in the majority of cases it is most efficient to achieve a LEO orbit of around 200km and then complete a transfer from this orbit.
Figure 18 - Various launch profiles with green being most efficient It is possible to calculate the maximum payload that could be transferred to an orbit altitude as the maximum mass of the payload and transfer module must be equal to the maximum payload that can be lifted to 200km which is known. The spacecraft would be able to lift between 25 and 40 tonnes to a 200km LEO orbit. This means that the mass of payload that can be lifted to a higher orbital altitude can be calculated using the assumption that the transfer is taking place in the equatorial plane. Graphs showing how the payload mass is affected by orbit altitude are shown in Appendix D.1. If the mission requires additional transfer to orbits with higher inclination then the mass that can be lifted to that altitude would be further reduced. These calculations would need to be done on an individual customer bases. If the payload needs to be re-entered so that it can be landed on the Earth, it is estimated that the heat shield would need to be 25% of the mass of the re-entering object. Therefore the mass that can be taken to an altitude and then re-entered reduces much more quickly than the mass for a one way payload. As heat shields improve it may be possible to reduce the weight of the heat shield and increase the weight of the payload. In orbit sources of disturbance Several sources cause disturbances that change the spacecraft attitude and orbit, more information about these can be found in Appendix D.2. Attitude Determination and Control System (ADCS) The ADCS electronics can be split into three distinct areas: the sensors that are used to take measurements from which the spacecraft attitude can be calculated, the actuators which are used
Robert Taylor Section 13 - In Orbit Operations Page 71 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 to change the satellite attitude and lastly the processing unit (ADCS control unit) that processes the data from the sensors and makes adjustments using the actuators. The ADCS control unit will be covered in the electronic section of the report. A spacecraft would need to be able to change its attitude in all 3 axis, to allow payloads to be released into space and to allow docking with space stations and other spacecraft. The first Down Selection process looked at the type of system that would be needed to stabilise the spacecraft in 3 axes. This therefore ruled out the option of the spacecraft being spin stabilised [EE05-04] as this would prevent payloads from being deployed. Also the spacecraft cannot be gravity gradient stabilised [EE05-02] as this would stop the spacecraft from being manoeuvrable. To ensure that the spacecraft is stable in all three axes and able to manoeuvre the spacecraft would need to be 3 axes stabilised [EE05-03]. This would require sensors and actuators as explained below.
13.1.2. Navigation sensors To navigate in space. the results of at least two navigation sensors have to be used to calculate an exact attitude, as one sensor would be unable to give a specific location. Therefore spacecraft carry more than 2 systems to improve accuracy and give the spacecraft a high level of redundancy. 13.1.2.1. Inertial Measurement Unit (IMU) – [EE05-17] During launch many launch systems use an Inertial Measurement Unit (IMU) which uses a combination of accelerometers and gyroscopes to calculate the craft's change in position. An IMU is used as it is very difficult to take measurements from outside the spacecraft during launch duetothehighlevelofvibrations.IMU’s are ideal for use over short time periods however over long time periods the drift that all gyros encounter makes them inaccurate. In many modern rockets such as Ariane 5, optical (laser ring) gyros are used. Laser ring gyros work by passing light from a laser around a loop of optical fibre. Two beams are used in each loop, sent in opposite directions. Changes in orientation affect the way the two signals interfere, allowing the change in spacecraft/rocket attitude to be detected. IMU can also be used in conjunction with other sensors to generate a predicated position when other sensors are unable to provide data (such as a sun sensor if the satellite is in eclipse) however this is not entirely accurate and a correction would need to be made when the sensor is able to provide accurate data again.
Once the spacecraft is in space and maintaining an orbit there are many combinations of the sensors that can be used to determine attitude. 13.1.2.2. Global Positioning Systems (GPS)- [EE05-07] The global positioning systems work on signals sent from satellites at approximately 20,200km altitude and if the spacecraft is operating below this altitude it should be able to receive GPS signals and therefore calculate its position above the Earth’ssurface.ThereforemostmodernLEO
Robert Taylor Section 13 - In Orbit Operations Page 72 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 spacecraft and satellites have GPS capabilities. GPS can also be used to calculate the altitude of the vehicle above the Earth’ssurface. 13.1.2.3. Horizon Sensors - [EE05-09] Horizon sensors can calculate the Nadir vector (the line between the spacecraft and the centre of the Earth). To make sure that the sensor can track the Earth 24 hours a day, it is common for horizon sensors to track the Earth using the thermal infrared part of the spectrum (The thermal emission from CO2 is 15µm). This allows the horizon sensor to track the edge of the atmosphere continually allowing a continuous reference point (Fortescue, et al., 2003). The only disadvantage is that the atmosphere ends gradually meaning that there is a small error in the sensor reading. 13.1.2.4. Star cameras – [EE05-08] Star cameras are currently the most accurate sensor with which to determine a spacecraft attitude and are capable of calculating the attitude of the spacecraft to within 1 arc second. To do this a largeamountofprocessingpowerisrequiredtoidentifythestarinthestarcamera’sfieldofview. The stars are identified by comparing them to a database of stars stored on the spacecraft. To make the system as efficient as possible the spacecraft database only contains stars that the spacecraft is likely to see. To do this the desired attitude and orbit needs to be known prior to launch. One of the major disadvantages of a star sensor is its complexity and size. The introduction of Charge Coupled Device (CCD) into modern star sensors has helped to reduce the mass, volume and power requirements of star cameras. They are typically around 3.25kg and consume around 10W of power (Fortescue, et al., 2003). Sun sensors, Magnetometers and Doppler ranging were also considered but were deemed not to be suitable for this project, more information about these can be found in appendix D.3.
13.1.3. Navigation sensors conclusion The system will use IMU, Horizon sensors, GPS and Star cameras to calculate its position as these would be able to provide accurate data for the spacecraft. Actuators Actuators are commonly separated into two categories - those that produce external torques like a gas thruster and those that produce internal torques like a reaction wheel. As well as attitude control actuators, it is common for a spacecraft to have a kick motor to control the speed at which the spacecraft is travelling through space. This is a large thruster on one end of the spacecraft; this is used to change the spacecraft velocity, allowing orbit changes to be made. Liquid Bi-propellant thrusters are often used for kick motor and attitude control as the thruster is suitable for both applications; mono-methyl hydrazine and di-nitrogen tetroxide (MMH+N2O4) are often used as these chemicals produce a hypergolic reaction. This means that they react in contact with each other and therefore do not need an ignition system. They are also capable of producing high levels of thrust. The main disadvantage of this arrangement is that the chemicals are very corrosive and
Robert Taylor Section 13 - In Orbit Operations Page 73 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 if they come in to contact by accident, it is like to cause an impromptu explosion. When the ADCS is combined with the kick motor it is often referred to as the Attitude and Orbit Control System (AOCS). Thrusters used for attitude control are normally used in pairs to ensure that the spacecraft rotates around its centre of mass. If a single thruster is used, the satellite position would also be effected when the satellite is rotated.
13.1.4. Hot Gas Thrusters - [EE05-13] Hot gas thrusters release hot gas as a propellant. This can be done with a monopropellant such as hydrazine N2H4 - when this chemical is passed through a catalyst it breaks down into N2, H2 and NH3 gases which are not from the reaction. This releases a lot of energy, producing around 10N of thrust.
13.1.5. Cold Gas Thrusters Cold gas thrusters are normally used in pairs to ensure that the spacecraft rotates around its centre of mass. The thrusters expel pressurised gas into space causing an equal and opposite force on the spacecraft causing the rotation. Cold gas thrusters are capable of creating large forces, however when the supply of gas (normally stored as a liquid) runs out the thrusters are no longer able to stabilise the spacecraft, meaning that they are suited to short missions or as a reserve system for emergencies. However they are able to produce large torques. This torque can be maximised by placingthethrustersasfaraspossiblefromthespacecraft’scentreofmass.
13.1.6. Pulsed Plasma Thrusters (PPT) - [EE05-16] Pulsed plasma thrusters are a new type of thruster that is currently under development. The thruster works by using a high voltage to create a spark next to a Teflon fuel block. The spark causes thrust to be created. STRAND1 (a satellite being launched by SSTL and SSC) is due to be launched early in 2013 and will test pulsed plasma thrusters. One of the main advantages over traditional thrusters is that the power to operate the thruster can be generated in space from solar panels or other power sources removing the need for cryogenic fuel tanks needed for traditional cold gas thrusters. However, the main disadvantage is that the spark creates an electromagnetic field that can interfere with other electrical subsystems on the spacecraft which can introduce errors or cause a system reset.
13.1.7. Magnetorquers - [EE05-06] Magnetorquers are electromagnets carried on the spacecraft. When a current passes through the magnetorquer a magnetic field is created. This field interacts with the Earth’smagneticfield, and the magnetorquer would experience a rotation force as the magnetic field of the magnetorquer would try to align with the Earth’s magnetic field, creating a torque on the spacecraft. Magnetorquers are ideally suited to low altitude orbit as this is where the field interaction is largest. As the altitude r increases the magnetic field reduces by a factor of 1/r3. They are capable of small torques making them ideal for attitude control. However, they are only able to produce
Robert Taylor Section 13 - In Orbit Operations Page 74 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 small amounts of torque and have limited performance close to the equator as the field direction of the Earth is orthogonal to the spacecraft meaning that there is no force between the fields.
13.1.8. Reaction Wheels- [EE05-11] Reaction wheels have a zero momentum bias, which means they are at rest under normal conditions. To rotate the spacecraft around an axis the reaction wheel spins in the opposite direction, the reaction wheel is then brought back to rest to stop the satellite rotation just as the required angle of rotation has been achieved. As reaction wheels have zero momentum bias they only require power when the spacecraft attitude is being changed. However they are only able to produce relatively small forces.
13.1.9. Momentum Wheels Momentum wheels have a momentum bias meaning they spin continually providing gyroscopic rigidity to the spacecraft, meaning that they have a constant power requirement from the spacecraft. To change the satellite attitude around one axis the rate at which the wheel is spinning is varied creating a rotational torque.
13.1.10. Control Moment Gyros- [EE05-12] Control moment gyros work by having three or more wheels mounted in gimbals. When the gimbal is rotated, a force is produced on the spacecraft as the spinning wheel resists the movement as it has gyroscopic rigidity. (Sellers, et al., 2005)
13.1.11. Actuators conclusion The system would use a combination of PPT, cold gas thruster, and control moment gyros to control its attitude and a Bi-propellant kick motor to control its altitude. 14. Electronics
The solutions for the electronic subsystems were selected using the down selection criteria. For the electronics reliability is also a key factor and this was also be used to down select solutions. If the system is reliable this would reduce the development program risk and the environmental impact as less systems would need to be produced. Reliability would also be a key factor that effects the systems reusability. Attitude Determination and Control System (ADCS) Control unit It is common for the ADCS to have a processor built in to the control unit for 2 reasons. Firstly the ADCS requires a large amount of processing to carry out the calculations required. Due to the number and complexity of these calculations, these should be separated from the main processor otherwise delays in processing of critical data may be encountered. Secondly if the ADCS system has its own processor it can provide redundancy for the main processor (Sellers, et al., 2005).
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On-board data handling (OBDH) The OBDH is the main controller on the spacecraft, it receives and issues commands to the rest of the spacecraft system and processes data from the spacecraft. There are 2 ways in which the OBDH can be managed.
14.1.1. Central Data Handling - [EE01-09] The first option is that all data is processed in the main central processor, this has some benefit as it reduces the volume and power required by the electronics (ESA, 2012). This leads to the system having a single point of failure and therefore a spare processor has to be carried on the spacecraft. One of the main draw backs of this system is that all of the data needs to be sent to the main processor, increasing the amount of data sent over the satellite bus. An integrated OBDH may contain other subsystems such as the communication, ADCS and TT&C but to ensure a high level of redundancy this is not suitable, as each of these subsystems should be separate, so that only one system is affected by a failure. A traditional computer (as used on Earth) could act as a centralised data handling system, although it is not suitable for use in space at it is too vulnerable to the radiation environment. The clock speeds are much higher than a spacecraft requires and therefore this makes the spacecraft vulnerable to errors.
14.1.2. Distributed Data Handling- [EE01-08] The second method is to make each subsystem ‘intelligent’ meaning that each subsystem can process its own data and then report the results to the main processor, reducing the amount of data sent over the communication bus. This is sensible as the systems such as the ADCS require a lot of processing power and therefore often need their own processor. A second advantage is that one of the subsystem processors can be used as a reserve for the main processor meaning that a spare processor is not required. Thirdly, as the processing is spread between several processors the speed at which data can be processed is increased without having to increase the clock speed of the central processor. A relatively low clock speed is desirable in space as it reduces the chance of errors. Reducing the clock speed is used with other processes to make component radiation tolerant.
14.1.3. First down selection Summary for the OBDH subsystem During the first down selection it was decided to go with a distributed data handling system, as this was deemed to be the most reliable type of system and the least likely to suffer from a catastrophic failure. This reduces the risk to the project during development. There are many types of technology that could be used to form a distributed data handling system.
14.1.4. Processors (microprocessors) - [EE01-05] Processors (microprocessors) can be used in space and offer a high level of computing power. There a variety of microprocessors available, some of which are radiation tolerant. Many space missions have custom microprocessors hardware designed for their specific requirements.
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14.1.5. FPGA with software processors- [EE0-04] Another processor option would be to use a FPGA for the main processor as this would allow a standard software core to be used to perform the on board computing. This has the benefit that the system is likely to have some heritage (been flown before in space). Also using a FPGA gives a large degree of flexibility as the software can be changed very easily, allowing updates through the systems life time.
14.1.6. Analogue computing- [EE01-07] Analogue computing was used in many launch systems including Apollo before the invention of digital electronics. Analogue computers do not have a clock signal; instead they process the data in real time giving the solution almost instantly, which is a major advantage over digital systems. Analogue computers have to be designed to complete a certain task and therefore are not flexible and they are also very complex. It would be unrealistic to build an analogue computer to control a modern spacecraft. Modern analogue computers can be implemented inside FPGA’s and can greatly improve performance of small tasks, although they tend to take more gates in the FPGA than traditional digital designs, which can be a disadvantage when trying to fit multiple processors into an FPGA. Designing an analogue process into a FPGA allows a combination of analogue and standard digital logic to be used together in the same system.
14.1.7. Second down selection summary for the OBDH subsystem During the second down selection, it was decided to use a combination of microprocessors and FPGA’s as this would provide a reliable system with a lot of processing power. Building a reliable system is key to making the system reusable. This solution is deemed to be highly technically feasible as it is a combination of tested technologies. The microprocessor which was going to be used for the main data handling computer and each of the intelligent subsystems would contain a FPGA. This was later changed due to the cost of the microprocessors affecting the projects financial viability; this is explained later in the section. Components Commercial Off The Shelf (COTS) [EE01-01] components are standard components. These are perfectly acceptable especially if redundant systems are built into the spacecraft. The main problem with COTS components is that they are vulnerable to failures caused by radiation which is much more common in the space environment. Therefore they are well suited to missions with a short life time or missions that are expected to receive a low radiation dose. They are not rated to fly and therefore not suitable for missions carrying human payloads. Although COTS components may be able to be used in certain systems (non-critical), as the system is going to carry humans, all critical systems should be built using components that are rated to fly. There are many types of components that are rated to fly such as Military off the
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Shelf (MOTS) components [EE01-03], which are designed to survive harsh conditions such as flight and radiation environments. Therefore, it is possible that MOTS components could be used in the design as they have gone through the required certification. Although, some of the components that are required by the system are not available as MOTS components and new radiation hard components may have to be developed.
Radiation hard components [EE01-02] are designed and built to be radiation tolerant. This means that the manufacturer rates them to be able to withstand a known level of radiation. Radiation hard components are often more expensive, larger and slower. A number of techniques can be used to make components radiation tolerant. First, using silicon on insulator or silicon on sapphire technology would reduce the chances of latch up (a condition that can destroy the processor instantly as the latched transistor shorts the power rail to ground causing dangerously high currents to burn the wires in the device). Insulation barriers can also be built in between transistors on the processor. Reducing the clock frequencies of the device and increasing the size of the transistors can also help to protect from bit flip (inverting data bits) and other signal events. The results of these changes to the design are that the device would be larger and require more power. Finally, as the processor would have to be designed specially, the cost per processor would be significantly higher. There are a variety of radiation hard processors on the market many of which are based on IBM R6000, MIPS R3000, and IBM/ Motorola PowerPC 603 & 750 architectures. One of the most reliable radiation-tolerant processors today is the PowerPC 750 RadLite. It can run at up to 133MHz. There are radiation hard FPGA that incorporate the same techniques. These are also more expensive than the standard FPGAs but more affordable than processors as they can be used for more applications.
14.1.8. Third down selection for the OBDH subsystem Having examined the many options available it was decided during the third down selection that the best way to implement the distributed architecture (selected in down selection 2) was to use a custom designed radiation hard microprocessor for the main processor unit. This unit will have Triple Modular Redundancy (TMR) to ensure reliability, a key down selection criteria. All other subsystems will include radiation hard FPGA allowing data to be processed quickly in each subsystem. All FPGAs will have TMR, either 3 sets of code inside one FPGA or 3 chips if necessary depending on the size of the software core. If any subsystem can gain an improvement in performance from using a small analogue computer this may be included in the FPGA design. The entirety of the design will be built of radiation hard components wherever possible.
14.1.9. Adding additional redundancy and error checking Information about redundancy and error checking can be found in Appendix D.4.
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14.1.10. OBDH Conclusion During the integration phase, the cost of current microprocessors and FPGAs was examined to estimate the cost of future components. It was found that RAD750 power- PC a type of radiation hard microprocessor costs approximately $200,000 (Vis, 2012). A radiation hardened FPGA was estimated to cost approximately $800; therefore it was decided to replace the microprocessor in the main on board computer with an FPGA to improve the financial feasibility. FPGAs are deemed not to be as resistant to the space environment as microprocessors; therefore the 25000kg variant of the system would have 5 times redundancy rather than 3 to mitigate the risk (this still works out less than the cost of the microprocessors). One of the advantages of using the FPGA is that it increases the level of future proofing as the software for the FPGA can be updated over time making the system more flexible and allowing modifications. 14.2. Electronics reliability The power on hours expected for each of the electronics subsystem over its life time would be less than half the power on hours experienced by equivalent electronics in satellites. Therefore the main factor that would affect reliability is the shock and vibration that the electronics would experience during launch therefore the mountings would require investment. Communications The communication link is a vital part of the spacecraft as this allows communication with Earth. The communication system is used to send telecommands, telemetry and other mission data between the spacecraft and Earth. To create a two way communication link between the spacecraft and a ground station on Earth, 2 carrier frequencies are required. By convention the uplink carrier to the spacecraft from the Earth station is at a higher frequency than the downlink carrier to Earth station from the spacecraft. The data signals can then be modulated with the carrier to send the data. There are 3 types of modulation that can be used to form a communication link; amplitude, frequency or phase modulation. The method of modulation that is used depends on several factors including the data rate and the distance that the link will need to cover. A diagram to show the different forms of modulation is in appendix D.5. To increase the system reliability and add redundancy, it is advisable to have two communication links [EE03-02]; this would include a highly reliable link for critical Telemetry, Telecommand and Control (TT&C) which will have a relatively slow data rate. A high speed data link could also be included that could be used for TT&C and other non-critical communication. This would require the satellite to use 4 carrier frequencies. An additional communications system could be built into some of the payloads such as a human support module to add additional redundancy.
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Communication frequencies are allocated by the International Telecommunication Union (ITU). Therefore, the ideal frequencies in the selected communication bands will be specified with the knowledge that the ITU may allocate a slightly different frequency to the project. The spacecraft would communicate with existing Earth stations and communication systems, as the setting up of new Earth stations would take a lot of time and resources. Earth stations need to be situated in an area where there is unlikely to be interference from other sources and require high quality data and power connections to ensure reliable operation. All data would be sent using the Open Systems Interconnection (OSI) model. This specifies a standard method of communication between the software and hardware layers of the system, shown in Appendix D.6.
14.2.1. Main communication system The main communication system could be implemented using a variety of different frequencies. However, it is very common for such systems to use C band (3.8 and 8GHZ) [EE04-11] as this offers a good compromise as this range of frequencies travels through the Earth’s atmosphere with the least refraction. Circularised polarisation is often used as this is easily sent and received. The ideal frequencies to use in C band are 4GHz downlink and 6GHz uplink. Another option would be to use an S band system (2-3GHz). [EE04-11] This type of communication system was used on the Shuttle and incorporated 4 antennas of which 2 could be used at any one time. The shuttle had two sets of frequencies - 2106.4MHz for the forward link and 2287.5MHz for the return link and 2041.9MHz for the second forward link and 2217.5MHz for the second return link. This would have allowed two shuttles to operate at the same time (Dumoulin, 2000). This should be considered in the design for the future spacecraft. Due to the low altitude, the time over a ground station is relatively short so therefore in many cases it is more efficient to transmit the signal to a GEO satellite that can relay the signal to the ground. NASA have set up the Tracking and Data Relay Satellites System (TDRSS) to perform this task for satellites and spacecraft operating in LEO orbits, as it reduces the number of ground stations that are required to monitor a mission. This system was successfully used with the shuttle system. Therefore, it would make sense for a future system to use the same satellite network; this satellite network is set up to use S band so therefore the new spacecraft would also need to use S- band using similar frequencies to those used by the shuttle. A bipolar antenna [EE04-05] has a length of ¼ wavelength and is ideal for UHF and VHF communications as it transmits the signal over a wide area making the signal easy to receive as precise pointing is not required. Information about other types of antennas that were considered can be found in Appendix D.10 Another option would be to use a flat antenna on the spacecraft, as this can easily be fitted into the structure of the spacecraft. Flat antennas such as a phased array antenna [EE04-03] have a series
Robert Taylor Section 14 - Electronics Page 80 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 of waveguides evenly spaced along the antenna. The wave front created by the antenna can be pointed electronically; this is done by introducing a time delay between each of the apertures transmitting the signal. This is shown below in Figure 19:
Figure 19 Diagram to show how a phased array can point a signal electronically, from (P-N Designs, 2011)
Allowing the signal to be pointed electronically removes the need for a mechanical pointing mechanism used with other antennas; this is an advantage as mechanical systems can be unreliable. The disadvantage of phased arrays is that they can be relatively large (as the frequency decreases, the size of the array increases as the distance between the apertures has to increase). Secondly phased arrays require large amounts of power. If there is the required level of power on the spacecraft and if the spacecraft is large enough to support the use of phased arrays, this technology could be used. Another option for the main communication link would be to have a laser uplink [EE04-01] from the ground station to the satellite. A laser would require precise pointing to ensure that it hit the receiver therefore it is much easier to implement an uplink than a downlink. Building an uplink laser on the Earth is easier to implement as there is not any restriction on the size, power or weight of the mounting. This would allow the laser to be fired directly at the satellite. If on occasion the laser beam missed the target receiver on the satellite the laser beam would miss the spacecraft and travel into space. However if a downlink link laser missed the target receiver it would land somewhere on the Earth, which is unacceptable therefore a laser can only be used for an uplink. This technique is suited for communicating with relatively slow moving objects such Geo-stationary satellites rather than fast moving LEO objects and therefore this would not be suitable for a LEO spacecraft. 14.2.1.1. Main communication summary A traditional communication uplink would be used to communicate with the spacecraft and the stages, using the VHF link; this would allow the use of existing ground stations. A higher data rate system for the human module would use two S-band phased array antennas and two S-band transceivers to communicate with GEO TDRSS satellites to relay signals to a ground station. Currently S-band phased arrays are under development by Ball Aerospace and Technology in America (Ball Aerospace, 2012). The phased array would allow the system to be built into the spacecraft and remove the need for a mechanical pointing mechanism. The phased array reduces the development program risk and is technically viable; however it would be relatively expensive.
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14.2.2. Telemetry and Tele-command (TT&C) To add additional redundancy many spacecraft have a separate link for telemetry and telecommand data, this link normally operates at a relatively low frequency to ensure that it is reliable. This is normally in either the VHF/UHF bands between 30-300MHz and 300-3000Mz. For this system, 136MHz should be used for the downlink and 148MHz for the uplink; this would provide a simple but reliable system (Chetty, 1991). The main advantage of using this frequency range is that the antenna can be a bipolar antenna that does not require pointing this means that the system will always be able to send and receive signals providing that there is power. If the spacecraft/rocket has stages that are jettisoned and need to be returned to the ground in a controlled way it, it would be sensible for each stage to have its own TT&C communication system, so that the electronics in the stage can communicate with control after separation. Spacecraft Communication Bus The communication bus is responsible for joining all of the subsystems, allowing commands and data to be sent between subsystems. This system needs to be highly reliable otherwise the spacecraft will not be able to function. There are several different arrangements that can be used to connect subsystems, these are described below. A diagram of the three topologies is included in Figure 42 in appendix D.8.
14.2.3. Centralised (Star) topology –[EE02-01] In a centralised (Star) topology all the subsystems are connected directly to the main controller, allowing very fast communication between the control and the subsystems. However it means the communication between two subsystems is limited as all of the data has to pass through the main controller. Therefore the main controller becomes a single point of failure making this topology unsuitable for systems that need a high level of redundancy. Also as the main controller is designed with a fixed number of connections, it is difficult to add extra subsystems, making the design inflexible, as new subsystems add to the amount of wiring significantly (Wertz & Larson, 2010).
14.2.4. Distributed (ring) topology–[EE02-02] In a distributed (ring) topology all of the subsystems are connected in a circle with data being passed from one subsystem into the next. This is not desirable as the communication is very slow and every module becomes a single point of failure. If any module fails then the subsystems of the spacecraft will not be able to communicate with each other causing a catastrophic failure. However, this system is very easy to implement (Wertz & Larson, 2010).
14.2.5. Bus topology–[EE02-03] A bus topology allows all modules to connect directly on to the same communication bus, meaning that any subsystem can communicate with every other subsystems. Therefore, there is no single point of failure and the system allows quick communication between all the subsystems.
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However, the shared bus limits the amount of data that can be transferred at any time (Wertz & Larson, 2010). The connections between subsystems are traditionally made using a system of cables. Another option would be to use wireless communication between subsystems; this has the advantage that it removes the need for heavy and complex wiring for the communication system. However the main disadvantage of this kind of system is that it is very vulnerable to interference. Interference could be caused by a variety of sources including solar radiation, cosmic radiation and human made sources. Therefore, this could only be used for non-critical systems. At the first down selection, the 3 types of bus topologies were compared and it was found that a bus architecture was best suited to link the subsystems in this project. The most reliable implementation of a communication bus is to use a bus topology with cables. Bus topologies are ideal for integrating with other systems as they have a modular design. There are several different bus systems that can be used in space which includes CAN, I2C, SpaceWire and IP1553 all of these were compared and the findings are discussed below.
14.2.6. Controller Area Network (CAN) bus–[EE02-05] CAN uses Low Voltage Differential Signals (LVDS) to send data over 2 wires over the bus. The protocol implements a collision detection protocol with arbitration allowing any subsystem to initialise communication; this allows the system to act as if there are multiple masters. This also allows new subsystems to be added without needing to modify the existing subsystems (ESA, 2012). The CAN protocols allow for error detection to be implemented which is a benefit as it allows more efficient communication. CAN is a standard communication protocol therefore modules already exist that are ready to run which could be used in the system, reducing the project risk. The system is capable of data rates of 1Mbit/sec. This is reasonably fast although not as fast as SpaceWire.
14.2.7. I2C–[EE02-08] The I2C system is designed around having a master that controls the bus (normally the main processor in a spacecraft) and slave subsystem that responds to commands from the master. As all communication has to be initialised by the master, communication is limited as the slave cannot send information on the bus unless the master asks for it. Therefore careful implementation is required to make sure that the master interrogates every slave regularly for additional information. If the master does communicate with each subsystem frequently, it is likely that key information will arrive too late to be useful and therefore is not suited to distributed processing architecture. This means that in many cases if the number of slaves is changed the code for the master will need to be altered, making a modular system difficult to implement. The standard clock rate is 100kz which is relatively slow for a communication bus (European Space Agency, 2012). The bus consists of a data and a clock line.
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14.2.8. IP1553 –[EE02-06] IP1553 bus standard was originally developed to be used in military aircraft but has since been certified for space and has been used in several rockets including Ariane. The major benefit of using this system is that it is designed to be highly reliable and is fully compatible with other aeronautics. This system has a fixed data rate of 1Mbit/sec meaning that it cannot operate as fast as other space bus systems. The bus includes error detection and monitoring techniques such as parity bits and Manchester coding. The bus requires a bus controller that acts as a master and must initiate all communication over the bus. The data is sent over shielded twisted pair cables, two sets of cables can be used to add redundancy as shown in Figure 43 in Appendix D.9
14.2.9. SpaceWire (based on IEEE 1355) –[EE02-07] SpaceWire (based on IEEE 1355) is designed for high performance and is able to provide bus speeds between 2 -200 Mbit/s. The system was designed for ESA (European Space Agency, 2012). The system uses point to point LVDS data connections using twisted pair wires and packet switching to route packets across the network from one location to another. This gives the system a high level of radiation tolerance. The routers allow areas of the system to be isolated from each other, increasing reliability and increasing the amount of data that can be sent over the system at any one time. SpaceWire has been widely used on a variety of missions and has proven to be very reliable - this includes missions launched by NASA and ESA. It is now possible to use SpaceWire and the 1553 systems with optical links between the subsystems which increases the reliability of the system as the chance of interference is reduced. An optical system should be used in future spacecraft, as long as the light source and detectors needed to implement the system do not significantly increase the weight of the electronics systems. Many fields of engineering are currently investing in optical technology. This should lead to improvements that help to reduce the mass and cost of optical links.
14.2.10. Down Selection for the Communication bus In the final Down Selection for the communication bus subsystem it was decided that SpaceWire was a slightly better choice than CAN as it has a higher data rate. SpaceWire and CAN are very similar systems. The SpaceWire bus will be implemented using optical links. The communication bus has a high technical viability, although the optical links increase the development program risk. Optical link are more complicated to implement than traditional cables, therefore this could reducing the financial feasibility slightly. Avionics – [EE11] During the integration phase it was decided that it was not necessary to have a winged vehicle for re-entry and landing therefore avionics will not be required in the traditional sense. However, the spacecraft will need to be controlled in the atmosphere. If the modules that are already built into
Robert Taylor Section 14 - Electronics Page 84 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 the design for spaceflight are not capable of controlling the craft in the atmosphere additional avionics modules may still be required. To read more about avionics see Appendix D.7. Temperature control –[EE09] All spacecraft need a thermal control system to keep its equipment within its specified operating temperature range. If the temperature of the spacecraft goes out of the desired operating temperature, it is likely that components will fail and the spacecraft will be unable to function. The temperature of the spacecraft can either be controlled through passive or active systems. In space the spacecraft can only loose heat through radiation, as there is no atmosphere for conduction and convection is not possible.
14.2.11. Passive temperature control Passive control is built into the design and structure of the spacecraft to control the temperature of the spacecraft without any intervention. Even though the spacecraft cannot lose heat through conduction, heat can be conducted through the structure of the spacecraft [EE09-13] to help equalise the temperature between the hot and cold side of the spacecraft - this is a passive control method. The heat that enters and leaves the spacecraft can be controlled by using coatings [EE09- 12] on the outside of the spacecraft that have different levels of absorptivity and emissivity. The spacecraft should be designed so that it loses slightly more energy than it gains as this is a safer failure mode if the active temperature control systems failed. Multi Layer Insulation (MLI) [EE09-14] is designed to reduce the radiation exchange between the spacecraft and space. It is made up of many layers of aluminized plastic films (Mylar or Kapton) separated with layers of low conductance spacers such as Dacron netting. Heat pipes [EE09-11] and capillary pumped loops use the principles of evaporation and capillary wicking to transport heat energy over short distances (Fortescue, et al., 2003). This method is often used to cool electronics processors.
Passive control is only suitable for small structures where the heat energy only has to be transported short distances. Larger structures are likely to generate significant heat gradients across the spacecraft unless the temperature is actively controlled due to the amount of equipment producing excess heat energy and the distance that it needs to be transported before it can be radiated. Therefore at the first down selection, it was decided that modules that would spend significant time in space (more than 12 hours) would require an active temperature control system.
14.2.12. Active Temperature control Active control systems require the temperature of the spacecraft to be monitored, so that the control system can heat or cool the spacecraft to keep the spacecraft within the desired range of operating temperature. This can be done in two ways; either the spacecraft can be designed so that
Robert Taylor Section 14 - Electronics Page 85 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 it would go cold unless the control system keeps it warm enough or the system can be designed in a way that it would over heat unless the control system provides sufficient cooling. 14.2.12.1. Heaters - [EE09-08] Heaters are used to stop subsystems falling below their designed operating temperature. The simplest heater can be produced by passing a current through a resistor. This can be mounted in Kapton to form a patch heater which can be directly mounted on to surfaces within the spacecraft. Another heating method is to use metal coaxial cable in which the core of the cable forms a heating element. Heaters are easy to implement although require a large amount of power when operational. 14.2.12.2. Mechanically Pumped Loops and Fluid Loops - [EE09-07] Mechanically pumped loops and fluid loops are mechanically pumped so that the coolant can transfer heat energy from one area of the spacecraft to another (this could be a radiator that radiates into space). If several systems need to be cooled, the system can be split into sections (to form fluid loops) with variable values allowing the rate at which coolant passes through a subsystem to be varied. This allows one system to cool several subsystems at different rates. The main disadvantage of such a system is that the pump requires power and the pump could cause vibrations on the spacecraft. However, this system has many advantages including that it is not affected by orientation of the spacecraft and the amount of cooling is variable. Variable conducting heat pipes and louvers and shutters were considered but were deemed not to be suitable as they have a high risk of a mechanical failure, more information about these can be found in appendix D.11
14.2.13. Temperature control conclusion It is likely that the system would not be suited to passive cooling due to its size. Therefore, research will be done into both passive and active methods of controlling the temperature. A combination of heaters and mechanically pumped fluid loops will be used to allow fine control over the temperature of every part of the system. This fine control would allow every module to be kept within its ideal operating range all of the time. This system would be relatively complex making the system technically challenging, however the system is critical to the reusability concept, as without it components may be damaged unnecessarily. System monitoring sensors To reduce the amount of inspection that needs to be completed on the spacecraft between missions, the spacecraft could be fitted with a variety of sensors that would provide information about the health of the spacecraft. For example the system monitoring system would able to gather information about the stress and strain on the structure of the spacecraft from sensors that are built in to the structural components. This information could be used to determine when major maintenance is required on the spacecraft. This should help to reduce the time required for
Robert Taylor Section 14 - Electronics Page 86 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 maintenance, reducing the cost incurred. Implementing this idea would have a mass penalty, as the sensors would add weight to the spacecraft. More information about system monitoring and its benefits can be found in Appendix D.12 Power A variety of different systems can be used to provide power to a spacecraft depending on duration of the mission and the amount of power that is required by the spacecraft. When the size of a power system is being designed the degradation of the power system must be taken into account - at the end of the systems life the power system may only produce 80% of the power it produced at the beginning of its life. The End of Life (EOL) power must still be enough to power the whole of the spacecraft.
14.2.14. Power sources 14.2.14.1. Primary batteries - [EE06-03] Short lower power missions can use primary batteries that are not rechargeable and therefore when the batteries go flat, the spacecraft would become unresponsive. Therefore, these are used for short missions and as emergency backup power. Under normal operation the spacecraft power requirement would make primary batteries unsuitable for this application. 14.2.14.2. Hydrogen fuel cells- [EE06-02] Hydrogen fuel cells work by reacting gaseous hydrogen and oxygen together to produce energy and water (the process is the reverse of electrolysis). The water produced can be used as drinking water for astronauts. The main limiting factor for hydrogen fuel cells is the amount of fuel that can be taken on the mission. Therefore, they are typically used on missions that last less than one month and have high power requirements. Hydrogen fuel cells can produce over 1KW of power. Notably they have been used on all NASA manned missions since Gemini 5 in August 1965 (excluding Skylab and ISS as the duration of these missions was too long). The Shuttle required 21KW at 28 volts and a peak of 36KW for 15 minutes. To achieve this, the shuttle had 3 hydrogen fuel cells producing 7KW each with a peak of 12KW each (Baker, 2011). 14.2.14.3. Secondary batteries - [EE06-05] and solar cells- [EE06-01] For medium duration missions it is common to use a combination of secondary batteries (that can be recharged) and solar cells. There are many different kinds of solar panels. Solar panels work by having cells that contain p-n junctions in the material, causing a depletion region. In the depletion region of the device, visible light creates electron hole pairs. The potential difference across the cell forces the holes and electrons to travel to opposite sides of the depletion region. These electron hole pairs create a current that can be used to power the spacecraft or charge a battery. Combinations of cells in parallel and series can provide a range of voltages and currents. Photons with excess energy (for example UV) cause electron hole pairs on the surface of the material, which immediately recombines without producing a current. This introduces heat into the solar panel, reducing the efficiency. Photons with relatively low energies such as infra-red do not have
Robert Taylor Section 14 - Electronics Page 87 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 enough energy to excite the electrons and therefore pass through without causing any changes. More information about solar cells can be found in Appendix D.13 The power generated must be enough to power the spacecraft and charge the secondary batteries that would be used to power the spacecraft during eclipse. Different types of batteries have different charge and discharge characteristics. For example, Ni-Cd can be fully discharged 21000 times whereas lead acid batteries can only be discharged 1000 times (assuming a depth of discharge of 25%). Currently Li-ion battery technology is starting to be used in space and this is superior to many other kinds of batteries. The only disadvantage is the li-ion responds badly to overcharging and therefore additional circuitry is required to protect the battery. 14.2.14.4. Radioisotope Thermoelectric Generator (RTG) - [EE06-06] Due to the short duration and the number of launches and re-entries that the spacecraft will complete, a RTG is not suitable due to its large mass and the high risk of radioactive material getting into the atmosphere. More information about RTGs can be found in Appendix D.14
14.2.15. Power regulation Once the spacecraft has generated power, this power needs to be distributed around the spacecraft. This can be done in several ways, outlined in Appendix D.15
14.2.16. Power Conclusions The rocket stages would use rechargeable batteries (secondary batteries). These can be recharged on the ground before each launch. As the rocket stages would be in flight for a relatively short time, it is unnecessary to have solar panels to recharge the secondary batteries. Batteries are an area that would require investment to improve the technology to reduce the weight of the batteries. For a human crewed module, hydrogen fuel cells provide the best solution as they are compact and can produce over 1KW of power. To make sure the system is reliable, there should be more than one fuel cell as this would provide redundancy. The human life support system should contain its own hydrogen fuel cell for additional redundancy. The power system should regulate the power bus voltage and each module should have a power switch to protect against failures - this is a hybrid power regulation system. Electronics for the Payload and Life Support Systems If a payload that can support humans is fitted within the spacecraft, this module should contain all the systems required to keep the astronauts alive including a backup power supply. If this is the case, the electronic links to the payload would be very simple, as only a power connection and a connection to the communication system would be required.
14.2.17. Deploying payloads Many payloads would need to be deployed into space from the spacecraft. This would require control systems and possibly a robotic arm or explosives to release the payload, more information about this can be found in Appendix D.16.
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14.2.18. Deploying payload conclusions It was decided during the integration phase that the spacecraft will have the payload on top of the rocket stages rather than within the body of the craft. Therefore, a robotic arm would not be required. The system would instead use a combination of explosive bolts, springs and decoupling rings to deploy the payload from the top of the rocket. Launch Errors Just after the launch of a rocket, a system is needed to check that the rocket trajectory is within the accepted flight plan, if the rocket flies out of the ideal launch trajectory this needs to be detected so that the launch can be aborted if necessary so that the spacecraft can be landed in a controlled way. More information about escape systems can be found in Appendix D.17. There should be a method of removing a human payload clear of a stage 1 or 2 failure. This could be done by moving the whole capsule to safety, a section of the capsule, or some kind of ejector seat. Cost of electronics Estimating the cost of the electronics for this system was very difficult as other rocket systems rarely make available such details about their designs. Secondly the shuttle is the only system that has used the electronics multiple times, however this electronics is 30 years out of date and therefore does not provide accurate data on which to base a cost estimate. Also the conceptual design that we are investigating has multiple stages, each of which will require its own electronics so that it can make the return trip and be reused. Therefore, to produce a cost estimate for the electronics, the cost of current radiation hard components was examined to estimate what the cost of future components will be. It was found that the radiation hard microprocessor RAD750 was estimated to cost $200,000 and radiation hard FPGA around $800, as already outlined above, this lead to a design change to remove microprocessors from the design in favourofFPGA’s.FPGA’s will therefore be used in all subsystems. An estimation of the cost of each subsystem is outlined in Appendix D.18. This shows the costs of a set of electronics for the 25 and 40 tonnes variants. It was estimated that a set of electronics for the 25 tonne variant would cost around $10,410,400 and that the 40 tonne variant would cost $11,684,640. Investment plan For a system to be built following the proposals in this report, investment in several technologies would be required to make sure that they are sufficiently developed for the project.
14.2.19. S-band phased array antenna NASA was considering an S-band phased array antenna for their latest spacecraft and some development has already been carried out by companies in the USA. Using a phased array removes the requirement for a moving antenna mount as the signal can be pointed electronically, improving reliability. These systems currently require a lot of power and hopefully the power
Robert Taylor Section 14 - Electronics Page 89 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 requirements would be reduced as part of the development due to the investment. It is estimated that, as some of the research has already been done, 4 years would be required to modify this design to work with the proposed system. It is estimated that cost would be around $10 million.
14.2.20. Radiation hard FPGA It is highly likely that there would be a radiation hard FPGA on the market that would be suitable for the system application, however there is a risk associated with this assumption and therefore it would be sensible to have the ability to invest if necessary. The design and development of a new radiation hard FPGA would take at least 8 years and cost at least $ 100 million.
14.2.21. PPT thrusters PPT thrusters are currently at the experimental stage. They have been flown on several missions to test their capabilities. They have a relatively long life and are ideal for small attitude changes and therefore would reduce the amount of fuel that is required for the gas thrusters used for large changes in attitudes. PPT will require at least 5 years of development and would cost at least $200,000.
14.2.22. System monitoring development This system would need to be built into the design of every module of the spacecraft; therefore the investment budget for this project would be split across many subsystems. The system monitoring system must be developed with every subsystem therefore the time required for development will be dependent on other modules. The cost of development would add cost to all of the subsystems; therefore the cost of development of each subsystem of the spacecraft would be increased by around 2%.
14.2.23. Electronics mounting Research would need to be done to make sure that the electronics can be mounted in such a way that the electronics can survive the shocks and vibrations of at least 50 launches. The circuit boards may require additional mountings. The mounting may also need to dampen vibrations. This should be a relatively small investment with the research only lasting 2-3 years and costing around $100,000.
14.2.24. Battery technology The stages of the rocket would contain batteries that are charged on the ground which should be able to power the stages during launch and recovery. Investment should be made to improve the charge density of battery technology. This would hopefully help to reduce the size and mass of the batteries required to power the system. Batteries would require 12 years of development. Considering the number of other industries that are developing battery technology (such as the car industry) a risk could be taken by not investing as other industrial sectors could develop the technology that the system would require. However investment may still be required and therefore a budget of $ 100 million is planned.
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15. Re-Entry All re-usable vehicles must make it back to Earth intact, and single use vehicles may need to return a payload. To return safely, the vehicle must be able to slow down to a suitable speed before approaching the ground. The main challenge is how to achieve this deceleration. Each of the re-entry concepts in the concept register has been considered. This section covers how the concepts were analysed and down selected to obtain the final concepts. 15.1. Re-entry concept introduction This section introduces the concepts that were selected for analysis 15.1.1. High altitude deceleration [RE01] A typical re-entry speed is 13Km/s (Mach 40) (J.E. Johnson, 2012). Immediately after deceleration begins, the orbit would decay and the vehicle would be pulled towards the Earth. This means that the vehicle would enter the atmosphere. Travelling at hypersonic speeds in the atmosphere produces tremendous amounts of drag, which also produces large amount of heat. The maximum safe deceleration for humans is limited by the human body to 20g (sitting backwards and upside down relative to the direction of travel) (J. E. Pavlosky, 1974) .The vehicle must be capable of surviving these harsh conditions and also then continue slowing down once aerodynamic drag is not as large at the lower speeds. This section introduces concepts for slowing down from orbital speeds to subsonic speeds. 15.1.1.1. Aerobraking [RE01-01] The atmosphere varies with the altitude from the ground. At sea level, the atmosphere is dense and warm. At the edges of space, the atmosphere is thin and cold (the atmosphere has no defined edge, as there are residuals at very high altitudes). Figure 46 in Appendix E shows how the temperature, speed of sound, pressure and density of the atmosphere vary with altitude. As can be seen, the pressure and density have dropped off to tiny values past an altitude of 40Km. This means that above this altitude, the atmosphere can be considered as absent. Spacecraft can move at ease, and at very high speeds. Basic thrusters can be used to provide initial deceleration, beginning the re-entry phase. Once inside the atmosphere, the air density starts to increase quickly and becomes a factor. As the atmosphere gets thicker and thicker, more and more air molecules collide with the spacecraft. At these hypersonic speeds, very sharp shock waves form around the front of the craft. These form superheated gas (plasma) around the craft. The temperatures are so high that molecular disassociation occurs, breaking the air molecules down. The kinetic energy from the craft is being used to produce this heating, which slows the craft down rapidly (forces of around -8g can be experienced in a typical re-entry) (J. Barrie Moss, 2003). This huge force can tear off any parts of the craft which protrude, so aerodynamics are very important. The plasma also creates a magnetic field around the craft, preventing communications.
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15.1.1.2. Reverse thrust [RE01-02] It is possible to easily face the craft in the opposite direction to the direction of travel and use the main propulsion to directly slow the craft down. This would decrease the crafts orbital velocity, but the craft would quickly start accelerating towards the Earth, so there is not a long time before the craft would have to re-orient for re-entry. 15.1.2. Low altitude Deceleration [RE02] Once the craft has slowed sufficiently, aerodynamic heating becomes less of a factor, and the craft can no longer be considered to be in orbit. This means that traditional aerodynamics become important, as the craft would need to be able to overcome the forces of gravity. The final velocity as the craft approaches the ground needs to be within the landing methods acceptable range. The following concepts can achieve this deceleration. 15.1.2.1. Parachutes [RE02-01] Parachutes are very effective at slowing a craft down, by dramatically increasing the drag. They can be stowed, being deployed when required. They can scale to provide varying amounts of drag for various craft. They work both in horizontal flight or free fall but are not suitable for hypersonic speeds, as the huge forces would tear the parachutes away. They can also be used to slow to aircraft down once on the runway, if landing horizontally. Vertically, the speed obtained by parachutes is usually low enough for touch down if shock absorption is used. Once deployed the parachutes can either stay out or be detached – they cannot be taken back in. Parachutes can either be purely drag producing or can also produce lift. The former is very simple and simply lowers the vertical speed. The latter gives a greater degree of control, allowing the craft to be steered. The disadvantage is that the landing gear will need to be more complex to overcome the horizontal and vertical velocity. 15.1.2.2. Flight [RE02-02] As the craft already has altitude and velocity, it is very easy to make the transition into winged flight. The deceleration is not achieved directly, but by flying the craft has much longer before hitting the ground to slow down. Thrust is usually not required, as the craft does not need to accelerate or climb. Gliding is very common, as this only requires wings and control surfaces. Flight also allows the craft to have control of where to land, with a large range. The glide ratio is very important, and determines how far the craft travels horizontally for a given altitude loss. The space shuttle had a glide ratio of 2 (2 metres forwards for every 1 metre down), which is much lower than that of any traditional glider. The main disadvantage of flight is that the craft will require substantial lift generating surfaces, which will add weight. They may also add drag during launch, as they may not be used. However, they may also be used during launch which would make them more weight efficient. They must also survive the high speed re-entry, meaning they cannot protrude too far and must be adequately strong and heat shielded. Deployable wings would help with this, but would be far more complex.
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15.1.3. Thermal Protection System [RE03] The high speeds during re-entry cause the air hitting the aircraft to be compressed by a large amount, which heats the air up (the friction caused by the air passing the craft also produces heat). Strong shock waves also form around the craft, providing huge temperature increases. The temperatures on the skin of the aircraft can reach over 3000K, which can be sustained for many minutes (J.E. Johnson, 2012). Heat shielding is required to protect the craft and its contents from these temperatures. Various methods are currently used to achieve this, often using a combination of each. The traditional method is to line the surface of the craft with a suitable protective layer. Reducing the in-atmosphere velocity of the craft will reduce the temperatures experienced and therefore the thickness of the thermal protection system needed, reducing the weight of the craft. The thermal protection system also needs to protect the craft from temperature fluctuations in space. The following concepts are capable of providing a thermal protection system. 15.1.3.1. High performance alloys [RE03-01] In shallower re-entries and sub-orbital flights, the maximum temperatures experienced are within the operating temperatures of various high performance alloys. Titanium alloys can operate at temperatures up to 500°C. These would simply form the skin of the aircraft, allowing it to survive the temperatures. 15.1.3.2. Ceramics [RE03-02] Ceramics offer excellent insulation and can withstand very high temperatures. They are suitable for protecting the craft from the very high temperatures during a hypersonic re-entry. Silica (SiC) tiles and reinforced Carbon-Carbon (RCC) were both used on the shuttle underside and nose/wing tips to ensure the shuttle could handle re-entry. One side of the ceramic can be at several thousand degrees Celsius, whereas the other side would only be slightly warm. 15.1.3.3. Ablative heat shielding [RE03-03] One method of protecting a craft from the heating is to utilise a heat shield which vaporises during re-entry. Vaporising the material requires a huge amount of energy, due to the materials latent heat capacity. This absorbs a lot of the heat generated, and keeps the craft protected. This does mean that the heat shield has a high mass, as enough mass is needed to prevent the shield being entirely burnt through. 15.1.3.4. Deployable heat shield One major concern is that aerobraking requires a large surface area to achieve a large bow shock wave and high drag. However if the re-entry craft is exposed during launch, the large surface area will produce undesirable drag, as an aerodynamic profile is preferred. A solution to this is to change the shape of the front of the craft once in orbit. Upon re-entry, the front of the craft could be much blunter. This can be achieved by a deployable structure. This structure would also act as a heat shield, as most of the heat generated from re-entry would be absorbed by this structure (or
James Dobberson Section 15 - Re-Entry Page 93 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 into the airstream passing around the vehicle). As this structure would experience extreme drag forces and heating, it would need to be strongly reinforced and be able to resist high temperatures. It could be either ablative or non-ablative. 15.2. Re-entry Initial Concept Elimination This initial down selection will aim to eliminate all re-entry concepts which are not appropriate for reusable launch vehicles. At this early stage, calculations were not necessary, as the concepts could be eliminated purely on judgement. All concepts were rated against the 6 categories as detailed in the Down Selection Methodology and concepts which didn’t score more than 50% overall were eliminated. Table 50 in Appendix E contains the scores. 15.2.1. Reverse Thrust deceleration [RE01-02] The advantage of this is that the craft would reach the atmosphere at much lower speeds, making the re-entry smoother. This may eliminate the need for/reduce the amount of heat shielding, as the maximum and sustained temperatures would be lower. This would lower the cost and weight of the craft. The disadvantages of this are that more work is needed to do this than a simple ballistic re-entry. As the speed is so high, only a tiny portion of the required deceleration can be achieved before the atmosphere is reached. Fuel would be required, which would have to stay on the craft until this point. This extra fuel would add weight to the craft. If this extra weight is larger than the weight saved in heat shielding, this will not be beneficial. However, if the craft is sensitive, it may not be able to survive a high speed re-entry, so may require manual deceleration. The use of aerobraking [RE01-01] can be 10 times more efficient than reverse thrust, as no work is required to slow the vehicle down, at the expense of extra heat shielding (Sellers, 2005). This concept scored poorly in technical viability, as the fuel required to achieve the deceleration would form such a large amount of the payload volume. This also scored poorly in the financial feasibility due to the reduction in payload capacity and cost of fuel. The development risk scored well, due to it being a very simple concept. The impact on other subsystem was poor, due to the large mass requirements reducing the available mass. The environmental score was low due to the large amount of fuel required for the manoeuvre. Finally, reusability was good because this method simply required refuelling after each flight. This concept scored 15/30 and was therefore eliminated. 15.2.2. Ablative heat shielding [RE03-03] The Apollo re-entry capsules ablative heat shield mass was roughly one third of the total mass. This offers the highest maximum temperature resistance, allowing the steepest re-entries. This is frequently used, as it is a simple system. However, this is only used for single use craft, as obviously the craft would not be suitable for another flight, unless the shielding is replaced.
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The technical viability was average – this concept is well established and simple, but is not weight efficient. The financial feasibility was very low, as the shield is not re-usable, and payload is also reduced by the high weight of the shield. The development risk is not a concern, as this is a well- established concept. The impact on other subsystems scored well, as it is a simple concept, which has minimal impact on the craft. The environmental impact is very poor, as the ablation of the shield cause large amount of pollutants to be released during re-entry high in the atmosphere. The shield also needs to be replaced after each flight. Finally, this concept is very poor for re-usability, as the entire TPS needs to be replaced after each flight. This concept scored 13/50 and was therefore eliminated. 15.2.3. Summary These two concepts were eliminated for different reasons. One is inefficient, whereas the other is not preferable for a reusable craft. 15.3. Lifting or Ballistic Re-entry Down Selection 15.3.1. Introduction This second Down Selection aims to decide between a lifting re-entry with wings or a simple ballistic re-entry. This decision would be based on (and also have an effect on) the landing subsystem, as it would determine either a horizontal or vertical landing. The mass penalty of the wings would be weighed against the benefits of a horizontal landing. The most beneficial option would be chosen. 15.3.2. Flight The presence of wings would have a large impact on re-entry. They allow a much shallow re- entry, making use of aero braking to reduce the re-entry velocity. The craft can also have aerodynamic control, meaning it can land on a specified runway, rather than in a landing footprint. However, there is a large mass penalty from the weight of the wings, fin, control surfaces and extra heat shielding required and the landing will also be rougher. This is less desirable, especially for crewed flights. The size (and therefore added weight) of the wings required is mainly determined by the weight of the aircraft at landing, as this is when the craft is going the slowest. The wings need to provide enough lift to overcome the weight of the aircraft to make a safe landing. The aircraft could land at a higher speed, requiring smaller wings. However, this would require a longer runway, which is undesirable. As the craft is not a dedicated aircraft, it would not be optimised for flight in the same way airplanes are. This means that landing at commercial airports is not possible. The space shuttle has a touchdown speed of 98m/s (Sellers, 2005) The wings would need to function at supersonic speeds to provide gliding capability during the descent. Swept back wings delay the onset of shock waves by reducing the effective airspeed seen by the leading edge, producing better performance at high speeds. Delta wings are an
James Dobberson Section 15 - Re-Entry Page 95 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 improvement on swept wings. They have a larger surface area for a given swept wing length, meaning the wings can be shorter to provide the same lift, which is desirable for re-entry. They also spread the wing load over a larger root, giving more strength. However, they give poor performance at lower speeds so landing speeds would be high. One feature of delta wings is a lifting body becoming simpler, as the wings could be an extension of the craft fuselage. The thick wings also provide storage space for fuel. To find the required wing size, it is necessary to consider the lift required at the landing speed of 98 m/s. This then gives the surface area of the wing required for a given wing coefficient of lift (higher at landing due to pitch up and possible extension of flaps). A sweep angle needs to be chosen, which would then define the wing length and maximum chord length. The sweep angle needs to be low, due to the extreme speeds experienced, and the requirement to reduce protruding components. This spreads the wing area down the length of the body rather than away from it. Once the wing dimensions are known, the structural weight of the wings can then be estimated. For a 40 tonne vehicle, landing at 98m/s, 4.01 m long wings are required. For a 25 Tonne vehicle, 3.18m long wings are required. Both of these are delta wings with a 50° sweep angle. However, the mass of the craft is likely to be significantly less than 40/25 tonnes due to fuel use and dumping of fuel/equipment before re-entry. Also, the body would contribute to lift. This means that this is the maximum length the wings would need to be. The analysis of this is shown in Appendix section F.1. The estimated structural mass of the 40 Tonne wing comes out low (596 Kg) due to the very small wings required. However, the increase in mass arising from the thermal protection system is large. If traditional re-useable ceramics are used, the extra TPS would weigh more than 1.7 Tonnes. If the deployable heat shield is used, the increase in mass would be 6 Tonnes. Also, the deployable heat shield structure will interfere with the aerodynamics of the wing, meaning very little lift will be generated. The spent rocket stages would land vertically near to the launch site, as defined in section 9.4. This means that the site would already have the facilities for a vertical landing (large open plain). Due to this, and the large impact on the other subsystems, the decision to use a simple ballistic re- entry with vertical landing was made. 15.3.3. Summary The decision made to eliminate all concepts involving flight leaves the remaining concepts to go ahead to further analysis. The concepts eliminated were [RE02-02], [LA01-01], [LA01-05], [LA02-01] and [LA02-02].
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15.4. Further Analysis and System Integration 15.4.1. Re-entry Profile A ballistic re-entry is one which simply selects an angle to re-enter at and then waits for the process to occur. It is very difficult to have any control at such high speeds. The re-entry angle is very important, as it would affect the temperatures and forces experienced. Shallower re-entries (<5°) spend longer at elevated temperatures, but see lower maximum temperatures. Entries which aretooshallowrisk“bouncingoff”theatmosphereandreturning to space. More aggressive re- entries (>12°) are risky, as the craft reaches thicker atmosphere much quicker, where the aerodynamic drag reaches the most extreme levels, producing the highest temperatures. A re-entry corridor is bound by a high and low re-entry angle, outside of which the re-entry would fail. The re-entry corridor would be specified for a given orbit and craft, and the craft must have adequate manoeuvring capabilities to stay within it. Figure 44 in appendix E illustrates a re-entry corridor. The craft would be capable of surviving higher temperatures but, as the heat shield would be reused, it is preferable to reduce the loading on it during each flight. A very shallow re-entry would produce a large landing footprint, but would be necessary until the heat shielding has been proved reliable. Therefore, an angle of 5° is suitable for the re-entry. A higher angle of up 10° would be achievable once enough flights have validated the thermal protection system. 15.4.2. Shock Profile It is important to prevent the shock waves from attaching to the craft at any point, which would produce temperatures which are simply too high for materials to withstand. The best way to achieve this is to use a blunt hull. This will produce a bow shock which follows the surface of the hull without ever making contact. This also has the additional benefit of increasing the drag of the craft, producing higher deceleration. As well as preventing contact at the hull, the sharp angle of the shock waves means that they could re-attach further along the aircraft at any protrusions, which would also produce extreme temperatures. If the craft had traditional long wings, the reattaching shock wave would destroy them so it is important to consider the shape of the shock waves which would form over the course of the re-entry and launch. 15.4.3. Design of conventional Thermal protection System Conventional thermal protection systems make use of aerobraking to generate drag and heat, a blunt body to spread the heating and increase drag, and a craft with a thermal barrier coating to withstand the heat. Reusable heat shields are ceramic based, and require maintenance after every flight. They are reliable, and simple to implement. As the space shuttle also re-entered from a LEO orbit, any reusable craft can be expected to require a similar TPS. The weight of this TPS would depend on the area of the crafts surface. Table 51 in section E.1 shows that the weight/area of each insulation type increases with the
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maximum temperatures experienced by that insulation, partly due to thicker layers being used in hotter areas. This makes it very simple to design a reusable ceramic thermal protection system for a capsule. The entire vehicle would be covered with suitable insulation, with thicker, heavier insulation used where the temperatures exceed a materials maximum temperature. The only exception is any leading edges which carry high structural load, making RCC is the only viable material. The calculated weights of the components of the ceramic TPS for a re-entry capsule are presented and compared to the deployable TPS in Table 30 - the total weight is 1.350 Tonnes. As can be seen, the heaviest part of the capsule is the nose cone, which is entirely dependent on the capsule radius. Decreasing the radius would have a much greater reduction in the overall weight than decreasing the capsule length. Figure 20 shows how the TPS would be arranged on the capsule.
Temperatures Insulation Total weight of Craft area TPS Surface area (m2) expected (K) selected insulation (kg) 20 RCC Nose Cone Common <1650 905 (See section A.2) (44.7 kg/m2) Capsule 35 HRSI Bottom Surface 650 - 1370 322 (8π/9X2.5X5) (9.2 kg/m2) (160°) Capsule Sides 22 LRSI Surface 315 - 650 88 (2x 50°) (5π/9X2.5X5) (4 kg/m2) Capsule Top 22 FRSI Surface Surface < 315 35 (5π/9X2.5X5) (1.6 kg/m2) (100°) 6.86 SIC Panels Deployable >1650 652 (See section A.2) (1900 kg/m3) Support 261 Deployable - < 315 Steel alloy structure (Estimated) 78.5 FRSI Capsule Deployable < 315 126 (2πX2.5X5) (1.6 kg/m2) Total 1,350/1,943 Table 30 - Summary of surface ceramic and deployable TPS for capsule Side View Top view Bottom View
RCC
HRSI
LRSI 5 m 5 diameter
FRSI 5 m Length
Figure 20- Summary of Ceramic TPS for capsule
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This is the lowest risk TPS for the project. It has been used extensively and mostly successfully in the space shuttle program. Development costs would be low, although each tile must be individually designed and made to ensure a correct fit. Replacing a single tile costs about $2000 and is time consuming, which was responsible for large turnaround times for the space shuttle. As the development risk of the deployable heat shield is high, a conventional ceramic TPS would be the backup system, to be implemented if the deployable shielding development does not produce good results. 15.4.4. Use of a Deployable heat shield The shield would need to be sufficiently wide to prevent the trailing shock waves from re- attaching anywhere on the craft. This is dependent on the craft length and width. These are largely dependent on the payload bay diameter (as this will have a large influence on the craft width), the fuel tank dimensions and the wing span if the spacecraft has wings, as these would protrude the furthest. US Patent 7837154B2 covers a “deployable heat shield and deceleration structure for the spacecraft. The patent cover page is shown in Figure 48 in E.2. It consists of a series of composite panels attached to a ring structure around the outside of a spacecraft. Once in orbit, the structure unfolds and locks in place, forming a cone shape around the nose of the spacecraft. This shields the trailing spacecraft from the forces and temperatures of re-entry. The development of this patent is covered in the investment plan in section 15.5.
The weight of such a device would be hugely dependant on the outer extremes on the trailing vehicle, such as the tail fin and wings. This, along with the vehicle diameter and panel angle would define the surface area of the panels required. This would then give the weight of the panels, which would be the majority of the weight. The supporting structure would consist of a strong metallic structure, whose weight would be hard to estimate. The weight calculations for the device were performed using a spreadsheet, where various dimensions could be altered. The weight was estimated by calculating the area of the cone required, based on the craft radius, the protrusion required and the cone angle. Then, the weight per area for the material could be used to find the total weight. The nose cone was treated as a dome of the same material and the remaining body was coated in LRSI. The total weight for the 5m long 2.5 m radius capsule is 1.943 Tonnes. The analysis of this can be seen in section E.2. The components weights are specified and compared against the traditional surface ceramic TPS in Table 30. 15.4.5. Parachutes After a ballistic re-entry, an unpowered craft would be travelling in a diagonal trajectory, at supersonic speeds. A series of methods are needed to land on the ground. The final impact velocity is the target for this.
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Parachutes are the superior method for achieving deceleration at these speeds. They need to be deployed from the vehicle, so the top part needs to be a section which can open to deploy the parachutes. Parachutes simply increase the drag coefficient of an object. This means that once a parachute is deployed, the craft would slow to a speed where the drag equals the weight of the craft (if in freefall). For a craft travelling horizontally, the parachute would simply achieve a rate of deceleration, which would continue until the craft is at horizontal rest (which could be free fall). Further chutes may be required once the steady state has been achieved in free fall, or if a horizontal craft needs to decelerate quicker. Bigger chutes achieve a higher deceleration, but deploying them at a high speed can produce too much drag, destroying the parachute, tether or craft. Because of this, the smallest parachutes are deployed first, followed by increasingly large chutes. NASA are developing a new parachute system for the Orion spacecraft for planetary re-entry (NASA, n.d.). Three different types of chute are used – drogue, followed by pilot and then main parachutes. The drogue chutes give the craft stability and deceleration at the higher speeds. The pilot chutes only deploy briefly, after the drogue chutes have been detached. These guide the deployment of the main parachutes. The main parachutes open slowly in stages, to lower the high loads seen on the cords. All of them have a nylon canopy and Kevlar cords and risers. This gives the parachutes a low weight and high durability. Using 3 main parachutes would achieve a final descent rate of 10 m/s. The main parachutes would weigh 309 kg, with the weight of the drogue and pilot parachutes being comparatively small. The analysis of the parachutes can be found in section E.3. 15.5. Investment Plan For the proposed craft, the following technologies would need to be developed before production begins in 2024. 15.5.1. Deployable Heat Shield 15.5.1.1. Description The proposal of a deployable heat shield for re-entry relies entirely on the development of such a device.USPatent7,837,154B2describesa“Deployableheat shield and deceleration structure for spacecraft”andwasfiledin2010.Thisdevicefitsaroundtheoutsideofacircularstructure.It consists of a series of panels which form a cone protruding from the craft when deployed. When un-deployed, the panels are close to the body of the craft, minimising drag. The supporting structure is strong enough to resist the high forces exerted on the panels. Once the device is deployed in orbit, it locks in place, not allowing the device to un-deploy. The supporting structure also still functions in the case of the structure warping under the effect of the heat. The benefits described include:
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50% weight reduction on known heat shield used on current craft It is possible to utilise small capsule spaces which would otherwise quickly become too hot for use. The volume requirement is far smaller than other inflation based deployable heat shields The shape is aerodynamically stable, stabilising the flight path during re-entry, reducing the landing footprint The temperature resistant materials used are non-ablative and the entire structure can be re-used. If damaged, the device is modular, so can be unbolted and replaced.
Figure 21 - Illustration of the device fitted to a cylinder (U. Trabandt, 2010) 15.5.1.2. Development time As can be seen from the patent, this device is well defined. The materials suggested are already developed; meaning the mechanical structure of the device would be the focus. This device does not un-deploy, which would cause problems for landing, due to obstructing the landing gear. If this is not possible after development, being able to eject the device/panels would also be beneficial. As nothing of this sort has been flown before, testing would be required to see how the device affects the craft. It is imperative that the shock waves do not attach at any point, as this would compromise the craft. Due to the very high Mach numbers that would be experienced, testing would have to consist of actual re-entries from orbit of test craft. This is very time consuming, but could be combined with the proof testing of other system components. Due to this, the development is likely to consist of a short design period, followed by a long testing period, relying on being able to launch test craft. The estimated development time for this device is 5 Years. 15.5.1.3. Development cost The majority of the development cost would come from the need to perform real tests of re- entering spacecraft. As the craft would rely on the heat shield to survive, the craft must not be in
James Dobberson Section 15 - Re-Entry Page 101 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 actual use, as the risk would be too high. A variety of craft shapes would need to be tested, such as short or long bodies, with varying shield diameters. This could however be performed on smaller craft. This means that 5-10 launches of test craft would be required to fully test this technology. It would be possible to find launch partners to share launches and reduce the costs, however, this could be time consuming. It would be also possible to test this using the crafts launch device, allowing both to be tested at the same time. The estimated development cost is $500 Million, which is mainly based on the cost of launch. 15.6. Future Proofing 15.6.1. Introduction The need to future proof arises from the length of the project and the developing competition. 15.6.2. Re-entry angle Higher re-entry angles allows for more accurate landings, but increase the strain on the craft from re-entry. The main risk is the deployable heat shield collapsing or the material failing. However, the material used increases in strength by roughly 10% at elevated temperatures of 1600 °C (U. Trabandt, 2010). This means that the main risk is the structure collapsing. It would be beneficial to inspect the structure after every flight for signs or excess loading or failure. This would allow the re-entry angle to be increased on subsequent flights, up to a maximum of 10 degrees. This makes the re-entry process quicker, and would allow a more accurate landing. If there are any signs of damage or wear found on the heat shield in flight which could compromise the structure, the re-entry angle should be reduced to the minimum 5 degrees. 15.6.3. Summary The full capability of the deployable heat shield is not being used at the start of the system life cycle. However, fairly soon, the re-entry angle can be increased to make use of the high level of protection and deceleration provided by the TPS, increasing landing accuracy and decreasing re- entry time. This may make site landings possible.
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16. Landing The final part of a mission is the landing. The craft is near to the ground at a low speed, and it needs to bring itself to a rest on the ground. Once on the ground, ground operations can take over, retrieving the crew or payload and transporting the craft back to the base of operations. This section covers the analysis and down selection of the landing concepts found in the concept register. 16.1. Landing concept introduction The landing concepts found in the concept register will be introduced in this following section. 16.1.1. Landing Method [LA01] There are various methods of landing a craft. The requirement is to bring the craft to rest as smoothly as possible. The landing method would depend largely on the landing location [LA02] and also if the craft is travelling horizontally or vertically. The following concepts all achieve landings which the craft would survive. 16.1.1.1. Runway with landing gear [LA01-01] Landing horizontally requires both wings [RE02-02] and a runway [LA02-01], [LA02-02]. The craft would typically have a much larger horizontal velocity than vertical upon approach, so the craft would need to bring itself to the ground at the start of the runway relatively gently. Then, the craft would need to reduce its horizontal velocity rapidly, before overrunning the runway. Typically, as spacecraft are not adapted well for flight, longer runways are required than usual commercial airports. The runway length required is a factor of how quickly the craft comes in and the rate of deceleration once on the runway.
Landing gear would be required, which would typically consist of a set of wheels with brakes. The landing gear could be reusable, or single use. To survive re-entry, they typically must be deployable. Parachutes would still function on a runway, which can contribute a large amount to deceleration. Additional deceleration is often required, as a very high rate of deceleration is required, which is not desired during flight. Deploying extra parachutes or using wheel breaks can achieve this. 16.1.1.2. Vertical landing with landing gear [LA01-02] Landing vertically is the simplest method, as very little control is required. The main issue is coming down too fast, which would easily damage the craft. Careful control is needed of the touchdown velocity. Also, a large flat landing site would be required to prevent the craft landing on anything. All craft would also require some form of landing gear, as even a very slow descent would put a shock through the craft. This final damping could take the form of an airbag or sprung legs, which would likely need to be deployable to survive re-entry.
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16.1.1.3. Vertical landing with thrusters [LA01-03] An advanced method of landing vertically is to use a series of thrusters to have complete control over the descent rate. This may be required for more sensitive payloads. This would require additional fuel, as well as the additional weight of the thrusters. They must also survive re-entry, which means they may need to be deployable of be capable or surviving on the outside of the craft. 16.1.1.4. Vertical with skycrane [LA01-04] The NASA Mars Science Laboratory (MSL) used a skycrane method to land on the surface of Mars. The main advantage of this was very careful control of the descent rate, as well as keeping the thrusters away from the surface. This would have caused the surface dust to be disturbed, which could have also damaged the rover. 16.1.1.5. Horizontal landing on water [LA01-05] Landing on water horizontally requires either landing gear or a carefully designed craft, which could glide through the water, using the drag to come to a halt. This would allow it to land on any large body of water, where it would then need to be recovered. 16.1.1.6. Landing vertically on water [LA01-06] This is the simplest of all landing methods, as landing gear may not even be required. The water would cushion the impact, and the craft would then float, awaiting pickup. This method is very commonly used, as only a capsule with heat shielding and a parachute is needed to return a payload/crew. 16.1.1.7. Mid-air retrieval [LA01-06] This method relies on a parachute being deployed, or a similar device with a long cord. As the rough landing location of the craft is known, a specialised aircraft can be waiting in the target area. Once visual contact is made of the craft, the aircraft can intercept it in mid-air. Using a special hook, the aircraft can latch onto the parachute of the craft. It would then tow the craft behind the aircraft, ready to be towed in. 16.1.2. Landing location [LA02] Where to land the craft is influenced by several factors. The location would need to be close to the orbit path, as the craft would not have the capability to travel far once re-entry has begun. Also, the landing site must have the facilities required, and a ground crew to welcome the craft. It is also preferable to land the craft as close to where the final destination is as possible (where the craft would need to be transported to). Landing closer to the base of operations would reduce operational costs and turnaround time, as well as being more environmentally favourable. There are also political constrains, such as certain territories not giving permission for a landing. Having the flexibility to land in a large variety of locations would make the craft more attractive to the customer. The crafts landing footprint needs to be as small as possible, as a large footprint could reduce the number of possible landing locations.
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16.1.2.1. Runway on Site [LA02-01] The most attractive location is a runway on site. The craft could land and then taxi directly to the final location, saving on transportation. This would require a craft capable of horizontal landings, and one with a good level of control. This would require a suitable site, with a long runway. 16.1.2.2. Runway [LA02-02] Being able to land at most airports would be a huge benefit, as the amount of landing locations is in the thousands. Runway length would be the limiting factor, which may limit the craft to spaceports and air force bases. The craft would then need to be transported to the base of operations (the space shuttle was carried on the back of an aircraft). The runway operator is likely to charge for the landing. 16.1.2.3. Water Landing [LA02-03] Landing on water is the easiest method for a craft, but is not the best location for ground support, as the recovery of the craft is difficult. The craft would also need to be rinsed thoroughly of seawater before re-use. This does make landing possible almost anywhere in the world, as no special ground facilities are needed, although landing too far from shore would make the recovery difficult. 16.1.2.4. Vertical landing on site [LA03-04] Landing vertically near to the operating base would be favourable for operations and the environment. This is difficult, as the footprint of the landing zone would often be large. This limits the location of the base of operations, as it must be near a large flat plain. 16.1.2.5. Vertical landing [LA03-04] This involves landing on any large plain, and then recovering the craft. The main locations this is possible is dried up lakes and flats. This reduces the risk of landing on anything. The craft and crew then need to be transported back to the base of operation. 16.2. Landing Concept Initial Elimination 16.2.1. Introduction This initial down selection would aim to eliminate all landing concepts which are not appropriate for reusable launch vehicles. At this early stage, calculations were not necessary, as the concepts could be eliminated purely on judgement. All concepts were rated as detailed in the Down Selection Methodology against 5 categories, and concepts which scored less than 50% overall were eliminated. Table 53 in Appendix F contains the scores. 16.2.2. Skycrane Landing [LA01-04] The descent and landing component of the Mars Science Laboratory spacecraft (consisting of the skycrane module and heat shield) had a mass of 2.4 tonnes, which lowered the Rover onto the surface, which had a mass of 0.9 Tonnes (NASA, n.d.). The descent stage had a mass over twice that of the rover, due mainly to the complex sky crane component, which contained 8 thrusters. The gravity on Mars is 2.7 times weaker and the atmosphere is much thinner (NASA, 2010). The
James Dobberson Section 16 - Landing Page 105 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 heat shield would need to be thicker and the skycrane component would need to be more substantial than one used on Mars. This means that a re-entry vehicle with a mass of 10 tonnes would require a descent and landing stage with a mass of over 25 Tonnes. This suggests that a skycrane approach would be inappropriate. The complexity of such a method cannot be justified for a mission landing on Earth, as there are no reasons not to disturb the landing site, as a gentle touchdown is the main priority. The technical feasibility was very low due to the large complexity of the system required. The Financial Feasibility and development risk were also high for the same reason of complexity. The impact on other systems was average as the returning craft would be largely unaffected, although much lighter. The environmental impact was very poor as the crane crashing will hugely affect the landing site. For reusability, this scored low and both the payload and sky crane need to be re- assembled, and the sky crane module may not even survive. This concept scored 11/30. 16.2.3. Mid-air retrieval [LA01-07] The advantages of this are that the craft does not need to worry about the landing, and can be quickly recovered and transported. There is also very little chance of the craft being intercepted if carrying sensitive contents. However, there is a risk of the intercept not being successful. It would be very high risk not giving the craft landing gear as a backup. Also, landing footprints tend to be large, and the small capsule travelling at high speeds would be difficult to find. This means that the landing would be heavily dependent on ground conditions.
The technical viability was very low due to the requirement of the catching vehicle, and the high risk. The financial feasibility was par due to the cost of running the second vehicle. The development risk was poor also due to the requirement of a second vehicle. The impact on other subsystems was average, as the landing and vehicle would be largely unchanged, but slightly adapted for the harsh manoeuvre. The environmental impact was average – the second vehicle had to be operated, but this saved on the transport of the vehicle and no landing site is required. Finally, the reusability is good, as the craft is returned to base as fast as possible. This concept scored 15/30 and was eliminated. 16.2.4. Summary The two landing methods eliminated were due to them being solutions to problems which are not faced by a reusable launch vehicle. 16.3. Landing method selection 16.3.1. Introduction This further down selection follows the decision to not use horizontal flight made in section 15.3. After the initial down selection made in section 16.2, the remaining landing methods are a vertical landing with landing gear, thrusters or a sea landing ([LA01-02], [LA01-02] or [LA01-06]). The landing location would then be determined by the landing method.
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16.3.2. Water Landing Water landings are the most common landing method for re-entering capsules, as the water absorbs some of the impact. The landing can vary a lot, so a series of tests have been conducted by NASA (Jr., 1959). It was found that the impact can vary depending on wind speeds, how calm the water is and how much the capsule is swinging on the parachute line. The worst case scenario is a capsule with a spherical bottom (best for re-entry) impacting at an attitude of 0° with an inclined flight path, which produces impact accelerations of roughly 50g. A capsule with a conical bottom only experienced roughly 10g, with attitude having no effect. These are relatively rough landings, compared with more controlled impacts with landing gear. For a reusable craft, the craft would need to be washed afterwards, and some components may need to be replaced if the craft is not fully watertight. This would increase the turnaround time for the craft being used again drastically, so it would be advisable to avoid a water landing. 16.3.3. Landing Gear or thrusters These methods are both similar, and would involve the craft landing vertically on a flat plain. They would follow the use of parachutes, which are the only remaining concept for deceleration. This means that they are dependent on the final speed achieved. The shock delivered to the payload depends on the weight of the craft, the impact speed and the shock absorption of the gear. To decelerate the 10,000 Kg craft from 10m/s as delivered by the parachutes to rest in 1 second (1.01G) would require shock absorbers 5 metres long (Section F.2). This is a very inconvenient length of shock absorber, almost taking up the craft length. They would also be unstable. They would also need to support the 10 tonne vehicle once at rest, and provide a total of 100 kN of force during the impact, meaning that they would need to be sufficiently thick to absorb the shock. Such devices, likely compensating hydraulic shock absorbers, would have a large weight, due to the dense steel they would be made of. Thrusters however only require a certain amount of fuel to produce the required thrust. The working distance can also be as large as required, as the craft does not need to be touching the ground. They could also be used to guide the craft towards a target or away from obstacles. They are also included in the vehicle emergency escape system used for launch. This means that they are already included and would provide a smoother, more accurate landing than using landing gear. Initially, a site landing would not be possible, as the lading method is unproven, and the landing location should be far away from obstacles. Small landing gear would still be required purely to act as the contact points with the ground, protecting the craft and providing stability. These could take the form of small deployable legs which are behind the heat shield.
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16.3.4. Summary Due to the additional benefits of using thrusters, landing gear ([LA01-02]) have been eliminated, leaving a vertical land landing with thrusters ([LA01-03]) as the landing method. This also leaves a vertical landing off site [LA02-05] as the remaining landing location. 16.4. Further Analysis and System Integration 16.4.1. Combined landing and launch escape system As per NASA standards, an emergency escape system is required when launching humans. This will take form of the very top of the payload (the re-entering part) detaching from the rest of the vehicle and climbing quickly away with the use of thrusters. If the launch is successful, these thrusters and fuel will remain unused, meaning they are available for landing. They would be used after the parachutes to achieve the final deceleration and a smooth accurate landing.
To lift the 10,000kg capsule 3 times the rocket height away in 2 seconds required a thrust of 1.05MN over a 2 secondburst. This is achieved using 4 Rocketdyne RS-88 engines and 688Kg of fuel (Ethanol and LO2). This defines what is available to achieve the landing. It would be possible to use only 118 kN of thrust to decelerate to rest from 212 m/s in 18 seconds. However, the escape system also requires the use of parachutes to bring the craft down to a soft landing. Therefore, it is not necessary to switch to thrusters at such a high speed. Instead, a much smoother controlled landing can be achieved from a lower speed. As parachutes would be used with a final velocity of 10m/s, the craft could decelerate from 10m/s to 0 m/s at a point 1 m above the ground. To achieve a comfortable deceleration of 1m/s over 10 seconds, the following thrust would be required:
This would take a distance of 50 metres, so the deceleration would begin at 51 metres, with the thrusters being adjusted appropriately to guide the craft and adjust the descent rate if required. This fuel use could then be found:
leaving 364 kg spare for the controlled descent from 1 m rest point. 16.5. Future Proofing 16.5.1. Introduction The lifetime of the project is 42 Years – 2 year for inception, 10 years of development and a 30 year operational phase. This means that the project is outdated before it is even operating. The need to future proof arises from still being competitive over the course of the project. Several concepts have future proofing opportunities.
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16.5.2. Site landing Initially, landing with the use of thrusters would be relatively new. Because of this, landing on site accurately should not be undertaken, due to the risks associated with landing in an area which isn’tclear.Thecraftwould need to wait for the recovery crew to arrive and pick it up, which could take a while due to the size of the footprint. As a large amount of landings occur during the operational phase, the landing system would demonstrate its capability. Using parachutes with built in guidance and gimbled thrusters, the capsule could then be steered towards an exact landing location. This would allow the onsite vertical landing method ([RE02-04]) to be switched to. The benefits of this are that the turnaround time will be minimal. The ground infrastructure could immediately begin to unload and prepare the capsule for re-use. However, it is important that the initial re-entry trajectory takes it away from the base of operations. This is to ensure any debris from any failures does not land anywhere populated. The decision to land at the base would be made once parachutes have been deployed. 16.5.3. Summary This improvement to the operation of the craft means that the turnaround time and cost would be dramatically reduced, as the recovery and transport segment of the operation would be eliminated.
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17. Payload and Markets
This section of the report is going to extensively look into the payload. Payload is defined by Loftus Jr. and Teixeira to be all hardware above the launch vehicle-to-spacecraft interface, excluding the payload faring. From this simple definition, the payload would consist of the space vehicle excluding the booster adapter interface (Larson & Wertz, 1999). The space markets that would be facilitated are going to be clearly defined. Also a further analysis would be conducted and the results would provide an understanding of the economy and possible financing through the determination of launch rates per year. This is very important as it would later be utilised to establish whether or not the system would be financially feasible. 17.1. Missions and Market Analysis In the inception report the payload mass was chosen to be in the range of 25,000 to 40, 000kg. To select the appropriate capacity, all the possible space markets were considered. It was recognized that for both active and decommissioned human rated spacecraft, human modules are relatively small in size. As a result, a decision was made to have two separate systems, A-variant and B- variant. The A-variant is smaller in size and human-rated, it has a payload capacity of 25,000kg. The B-variant is larger and would only carry cargo, it has a payload capacity of 40,000kg. From the preliminary market research the primary missions that were selected to be carried out by the final system design are as follows: People to ISS, people to space hotel, space station supply, GEO satellites (and MEO satellites), LEO satellites and space infrastructure. These markets were analysed further and a forecast of the launch rate per year was determined using both historic data and current data. Historic data was taken for systems like space shuttle, Proton K, Delta IV, Altas III and Soyuz (Isakowitz, 1999). Current data was taken for falcon 9 (Malik, 2012).The data that was recorded includes: the number of launches per year and the cost per launch. The values were averaged and are recorded in Table 54 shown in Appendix G.1. Using both historic and present data, the launch rates per year for each market were determined to be 8, 10,8,8,4 and 8 respectively. The launch price, in millions, was found to be $637.5, $133, $ 165, $201, $321.6 and $149 respectively. The average cost per kilo was calculated by dividing the launch price with the payload. In descending order value the average cost per kilo was determined to be $7868.75 for conveying people to space station and space hotel at, $7275.5 for launching small supplies (25,000kg) to Space Station, $7249.5 for space infrastructure projects, $5650.25 for launching GEO satellites, $ 3637.75 for launching large supplies to the space station (40, 000kg). The B-variant would be able to launch more than one satellite at a time - for example two GEO satellites. It would also be able to launch multiple small satellites using the honey comb
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17.1.1. Results
80
60
40
20
0
2024 2026 2028 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 Totalnumber of missions Years in Service
Figure 22: The Forecast of The total number of mission per year in services over 30 years
The forecast of the market trends from 2024 to 2054 is shown in Figure 22. The predictions which were made are stated in this passage. Both the A-variant and the B-variant will come into service in the year 2024. These would launch humans and small supplies to ISS and; satellites to LEO and GEO respectively. Each existing market would have a growth rate of 2% throughout the whole 30 year period (Corporation, Futron; Webber, D;, 2003). The large increase in the number of missions in 2030 (from 8 to 22 missions) is due to the commencing large projects, Space infrastructure including the construction of a Space Hotel. The second rise in the number of missions in 2035 (from 25 to 33 missions) would be as a result of the increase in the number of human cargo commuting to newly built space stations and space hotels. Large supplies to the space stations and also to the hotels would be taking place. The final sudden rise in the number of missions per year in 2045(from 48 to 60 missions) would be due to the doubling of the Space Hotels which would result in more demand for tourists and supplies. The dip in 2050 is due to the drop in the market share as the system would be reaching the end of its service. The shape of the trend is hugely influenced by the global market share which begins at 30% in 2024, rises to 50% in 2030, rises again to 60% in 2037 and finally depreciates to 50% in 2050. 17.2. Volume per mass Here, launch vehicle payload is made up of the entire spacecraft above the booster adapter
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Table 31 shows that the calculated average volume per mass was 0.0.12676 . This can be used to calculate the volume as a function on the mass for the different payloads.
Table 31: Launch Systems Characteristics information obtained by (Larson & Wertz, 1999)
Launch Upper Payload Payload Dimensions Volume Volume Volume System Stage (If to LEO Diameter Length (m3) per unit per unit any) (kg) (m) (m) mass ×10-3 mass (m3/kg) (m3/kg) ATLAS IIAS Centaur-2A 8,640 4.2 12.0 166.25 19.242 0.019242 DELTA II PAM-D 5,089 2.9 8.5 56.14 11.032 0.011032 (7920/25) PEGASUS 460 1.3 4.4 5.84 12.696 0.012696 XL SHUTTLE 24,400 4.5 18.3 291.037 11.928 0.011928 ARIANE-5 L9 18,000 4.5 12 190.85 10.603 0.010603 PROTON D1 20,900 4.1 15.6 205.96 9.855 0.009855 LONG MARCH 13,600 3.8 6 68.047 5.003 0.005003 (CHINA) CZ3B Falcon 9 9000 4.6 11.4 189.456 21.051 0.021051 Average 12511 12.676 0.012676
17.3. Human Module Payload Mass Fractions The human rated A-variant system has the capability of carrying 25 people to space. However, at the beginning of service the system would only be carrying a maximum of 10 people. This is because the system would need to comply with the safety requirements by having a Launch escape mechanism. In 2030, the A-variant will begin to carry 25 people. This system will be safe. Below are the illustrations of the Human module payload mass fractions. The fractions represented in Table 56 in Appendix G.1 are as follows: structures 7.76%, Escape system 6.245, Heat shield 10%, propulsion 31%, electronics 12%, ECLSS hardware 14%, supplies 7.3 % , human cargo (x10) 3.2% and cargo 8.5%. The fractions represented in Table 55 in appendix G.1 are as follows: structures 24%, propulsion 23%, electronics 12%, ECLSS hardware 14%, supplies 19 % and human cargo (x25) 8% 17.4. Seating Arrangements For the A-variant the passengers would occupy the nose cone section of the rocket. The module would have an upper deck and a lower deck. Each deck would accommodate 5 passengers. (no need for seats as no human cargo in B variant).
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17.5. Human Cargo Several manned space missions have taken place up-to-date. The majority of the missions were scientific. Nowadays Space exploration has become commercialized and more and more people are interested in visiting space, resulting in the birth of space tourism. Space tourism is one of the markets that the reusable launch and delivery system will facilitate. As we know, space is a hostile place which lacks the basic resources necessary to support life. An artificial life support system has to be integrated into a manned spacecraft to allow passengers (astronauts, crew or the public) to function and work efficiently. Therefore it is expedient to investigate the requirements of the human and the Environmental Control and Life Support System (ECLSS).
17.5.1. Human Requirements Human requirement expected by ECLSS technical authority are as follows:
Total pressure 70kPa/30 % O2 and Partial Pressure of oxygen (PPO2) 19 kPa (Carrasquillo & Anderson, 2012) Temperature 18.3 – 26.7 degrees Relative humidity (RH) 25-75 % RH
Maximum Partial Pressure of carbon dioxide (PPCO2) for short duration 1.01 kPa Exposure to contaminants SMAC value for 7 days: 15ppm/ 30 mg/m3 Total mass of consumables 30.60 kg per person per day Maximum G-forces: 4g
17.5.2. Environmental Control and Life Support System (ECLSS) The ECLSS includes many subsystems that carry out different tasks. To control the atmospheric conditions in the human payload module an atmosphere revitalisation system, external dust removal system and atmosphere control unit are required. The ECLSS allows the humidity and temperature to be controlled. The ELCSS subsystem is also responsible for water recovery and management this includes managing waste products. The ECLSS also contains fire detection and suppression equipment. Other elements that may be deemed to be part of the ECLSS are food storage and preparation, radiation shielding, hyperbaric chambers and air locks. ECLSS subsystems that are required for the mission would be dependent on the payload and its requirements. A human payload module would require more ECLSS subsystem. For unmanned missions, the environmental control is only considered, which is made up of atmosphere control and supply and, temperature and humidity. In this section the research into appropriate technology will be carried out, and the ‘best’ components will be integrated into a system. The system is heavily based on reports by NASA.
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Space is hostile to humans therefore it is necessary to provide an artificial environment for humans to survive and maintain health in space. The requirements for the ECLSS are listed: Provide life support functions for all the crew members provide life support for the entire mission without resupply from Earth, operate during launch, transit, descent and surface g-loads, reliability, maintainability and safety, provide enough atmosphere, gas composition and pressure, must have necessary Gas Storage for mission duration, must have adequate ventilation, shall provide temperature and humidity Control, must have fire detection and suppression, must provide the entire crew with adequate sources and amounts of food and potable water for a 7days, must provide adequate supply of hygiene water, shall provide psychological support by taking into account crew environment and other human factors and target life support system power usage. System trade factors were considered: Total Launch Mass (system and consumables), safety, reliability and cost. 17.6. Down Selection between Regenerative ECLSS and Non-regenerative ECLSS Non-regenerative ECLSS This is a more traditional method which has been used on many inhabited US space missions. This system requires supplies for the human cargo to be sufficient for the duration of the mission. If feasible, the system could be resupplied with consumables from the Earth. This system is generally made up of simple technologies and is easy to use. This is suitable for short duration mission lasting days or a few weeks. It is not necessary to recycle oxygen, water, or other masses since adequate quantities could economically be carried on short missions, with the simple technologies being significantly cheaper than the cost of recycling system.
Regenerative ECLSS Most emerging exploration vehicles are adopting a regenerative ECLSS as they are looking to conduct deep space exploration with the like of manned lunar mission and the mission to Mars. Regenerative systems recycle waste products in order to recover mass like oxygen and water. This becomes very important as mission duration increases (months – years) as it keeps reduces the storage space for consumables and also reduces the costs of resupplies from the Earth. The down selection method will eliminate concepts that are unsuitable for a delivery mission to LEO, which is has short duration (two days for a single journey). Also methods that will be deemed as unsafe for the human cargo will be eliminated. Technical Feasibility: Both systems have the ability to meet the needs of human cargo. At systems’levelaregenerativesystemweighsmorethananon-regenerative system. (Metcalf, et al., n.d.) quotes that “The totalmass and volume ofairand waterincludingthe associatedsystem hardware required to support crew metabolic needs for short duration missions is normally less
Mukudzei Muchengeti Section 17 - Payload and Markets Page 114 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 than the mass and volume of just the hardware required to reclaim and reuse those same resources.” The capability to carry passengers would have to be compromised in order to accommodate the regenerative ECLSS system hardware. Financial Feasibility: A lot of money would need to be budgeted for the purchase and maintenance (including spares) of the regenerative system. The regenerative system is generally more expensive than the non-regenerative system. This is illustrated in the NASA budget for the ISS ECLSS. NASA has budgeted $233 million for the regenerative system and only $55.2 million for the non-regenerative system (GAO, 2011) Development Program Risk: The regenerative system is highly likely to incur unplanned project costs and delays as the system is complex and being developed. So far the water recovery is being used on the ISS but crew are unwilling to drink the recycled water. On the other hand, components of the non-regenerative system are readily available. Most of these elements especially fire detection methods are already used domestically. Integration: The regenerative ECLSS would require more energy so more demand will be placed on the electronics. Environmental and Reusability: Used up small components, like filters, are expendable. In the non-regenerative system CO2 isventedoverboard.However,inastudybyNASA“Modelling the Effects of Spacecraft Venting on Instrument Measurements of the Martian Atmosphere for an EllipticalOrbit”ElainePetro(NASA/GSFC)andDavidHughes(NASA/GSFC)concludedthat vented didn’t pose a hazard to the equipment. 12-14 July 2011. Bigger components like the condensing heat exchanger can be reused. Table 32 shows a summary of down selection between a regenerative ECLSS and a non- regenerative ECLSS
Table 32 – Summary of ECLSS Down Selection Non-regenerative ECLSS Regenerative ECLSS (0 – 5) where 5 is better Technical Viability 5 5 Financial Feasibility 2 1 Development Program Risk 4 2 Environmental Impact 3 4
Integration 4 3 Reusability 2 3
Total (of a possible 30) 20 18 Based on these results, the regenerative system was eliminated.
17.6.1. Open v Closed loop Both can be used in a non-regenerative ECLSS. Closed loop systems can be used to control and maintain temperature and atmosphere constituents. For example Hybrid LSS can be used such as
Mukudzei Muchengeti Section 17 - Payload and Markets Page 115 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 the temperature control aboard the space shuttle. A closed loop can be used but if the system exceeds its capacity, the open loop evaporator is used to cool the shuttle. (Anon., n.d.)
17.6.2. Analysis of ECLSS Subsystems In the following sections the main subsystems of the ECLSS will be dissected and inherent technologies for different components will be evaluated. All of these technologies are Off-The- Shelf. For each technology the advantages and disadvantages are noted. These facts would be used to down select the technologies using the down selection methodology. The ‘best’ technologies for the respective subsystems will be assembled to form the optimum non- regenerative ECLSS. 17.7. ECLSS 17.7.1. Atmosphere Revitalisation 17.7.2. CO2 removal Non-regenerable absorption are only suitable for open loop operation, sorbent cannot be regenerated.InastudybyNASA“ModellingtheEffectsofSpacecraftVentingonInstrument Measurements ofthe Martian Atmosphere foran Elliptical Orbit” Elaine Petro (NASA/GSFC) and David Hughes (NASA/GSFC)concludedthatventeddidn’tposeahazardtotheequipment. 12-14 July 2011. Adsorption Processes are suitable for closed-loop operation. They are less susceptible to degradation because there are no chemical reactions involved. Contaminants can degrade performance over time but they are desorbed from the sorbent with heat or vacuum.
Technology Description Advantages Disadvantages Lithium hydroxide Removes CO2, trace Capacity- 0.92kg Life cycle: for 7 (LiOH) contaminants and CO2/kg sorbent people changed every other odors. - Costs for 10 people 11 hours, Been used for all US overall mission Very high EMS of space habitats except duration $1560 9333 for skylab.
Sodasorb A mixture of calcium CO2 goes into Capacity- 0.488kg hydroxide with solution and forms CO2/kg sorbent sodium or potassium carbonic acid as “activators”. Water is required for the reactions that consume CO2 Superoxides E.g. potassium Potassium hydroxide Capacity- 0.388kg superoxide, react with absorbs the CO2. CO2/kg sorbent, moisture in the USSR uses method Release oxygen and atmosphere to produce heat enough to cause O2 and potassium fire, hydroxide KO2 causes irritation of the eyes and respiratory tract. Chemical degradation
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affects performance
Molecular sieves Zeolite, principle Water and air save Zeolite has more (4BMS) adsorbents. The features, affinity for water than capacity of molecular reasonable EMS of CO2 sieves for CO2 is 418, ranked 3 and Quite large and temperature and ECLSS community complex pressure dependent concluded that it is Zeolite particle Very common in ducting space vehicles been Contamination of used for multiple downstream missions components Membranes Used to remove CO2 CO2 concentrations on from natural gas, at space habitat are low concentrations usually very low between 4-40%
17.7.3. Down Selection Area LiOH Sodasorb Superoxides Molecular Membranes (0 – 5) where 5 is Sieves(4BMS) better Technical Viability 3 3 2 3 1 Financial Feasibility 4 3 2 3 1 Development 4 2 2 3 1 Program Risk Environmental 4 2 1 2 2 Impact Integration 3 3 3 2 3 Total (of a possible 18 13 10 13 8 25)
LiOH has the highest score in the table. Molecular Sieves are too big and heavy. 17.8. Atmosphere Control and Supply (ACS) The ACS subsystem performs the roles of storage, distribution, conditioning, pressure control of the atmospheric gases, vent and relief capability; and habitat depressurisation or re-pressurisation. Sufficient quantities of oxygen and nitrogen must be available from storage. It is important to maintain the correct composition and pressure of the atmosphere to ensure the safety of human cargo and performance of equipment. The partial pressure of oxygen (PPO2) must be high enough to ensure sufficient absorption during respiration, therefore it is important to monitoring and controlling the pressure level. The total pressure of the pressure habitat required by the ECLSS technical community is 70.3KPa/30% O2 such that The PPO2 is 21KPa. The pressure would need to be adjusted for the human cargo to easily translate from the pressure of the pressure habitat to that of the space station or space hotel. When the method of atmosphere control and supply is decided the following factors must be considered: safety, reliability, the volume available and the amount of power requirement. The
Mukudzei Muchengeti Section 17 - Payload and Markets Page 117 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 location in which the compression/ liquefying of gasses would take place also needs to be considered as this could be done in orbit or at a facility on the ground. 17.9. Storage of Oxygen and Nitrogen O2 Storage for 10 people total mass of oxygen for whole mission is 240kg Technology Advantages Disadvantages Storage Tank size High Simple, reliable, Heavier and larger for larger Higher pressure, thicker pressure gas optimum for amounts wall, higher volume smaller amounts Cryogenic -Lighter than Gas -Sensitive to heat leaks Lower volume for the same Liquid for larger -fluid delivery in mass of fluid, smaller quantities, -higher microgravity is more container storage density, complex safer (lower pressure) Chemical Reaction processes occur to Solid, only 0.38kg oxygen Storage produce O2, vigorous generated per kg of reactions with water, may absorbent. Thus, required explored from friction, heat higher volume to produce or contamination, TOXIC sufficient oxygen4 (cameochemicals, n.d.) Hydrazine Product H2 can be used id Very toxic needed for another process such as propulsion
17.10. Down Selection Area High Cryogenic Chemical Hydrazine (0 – 5) where 5 is better pressure gas Liquid Storage Technical Viability 5 3 1 2 Financial Feasibility 4 3 2 2 Development Program 5 2 2 1 Risk Environmental Impact 4 4 1 1 Total (of a possible 20) 18 12 6 6 The concepts of storing oxygen in chemical storage and as a cryogenic liquid are being down selected. High pressure gas has the highest score. ISS also stores oxygen and nitrogen as high pressure gases in tanks. For a short duration mission the amount of oxygen that would be needed is up to 240kg, this is a small quantity. Storage as a high pressurised gas would be easier and more reliable compared to the cryogenic liquid which has difficulty flowing in microgravity.
17.11. Adjusting Pressure Method Description Advantages Disadvantages Depressurisation by Excess atmosphere is Simpler Mass of make-up venting vented to space gases must be
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sufficient to compensate for leakage, high volume Depressurisation by Excess gases are pumped Mass of make-up More complicated pumping back gases back into storage tanks gases is reduced into storage Re-pressurisation by Make-up gases in storage Easy to operate pumping gases into can be pumped into the the atmosphere atmosphere 17.12. Monitoring Major Atmosphere Constituents Many of the instruments used for monitoring trace contaminants are also able to monitor the major constituents. One approach is to monitor the partial pressures of O2, N2 and CO2; and to sum them to get the total pressure which can be compared with the sensor system as a system check. The O2 and the N2 measurements are then used to control the addition of these to make up for losses. (Tatars & Perry, 2004)
Method Description Advantages Disadvantages Mass spectrometry Monitors Stable , repeatable Cannot distinguish between all gases performance different compounds having the Used on ISS same molecular weight e.g. N2 & Moderate cost CO Complex Requires a roughing vacuum resource to start up CHEMFET/ISFET Detects H2, All solid state Some metals can be deposited on (Ion Sensing Field H2S, NH3 Small size the electrode which can give faulty Effect Transistor) and CO Wide temperature reading. operations Has to be cleaned well after every Low cost use to make sure reliable results are Long term stability read (4.5 kg in air) Easy calibration Low power Metal oxide Detects CO, Low cost Operate at high temperatures 300- o CH4 and Easy to use 450 C combustible Simply to produce Better results calibration is time gas Detects a large amount consuming and expensive of gases Separate analyzers Costs $800- High life cycle, - Many components to read and for each gas (e.g. 1000 10to+50 degrees maintain for a specific O2 Celsius sensor) P860-3O to P860- 5O P860-3N to P860- Costs $800- 5N 1000
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17.13. Down Selection
Monitoring Major Mass CHEMFET/ISFET (Ion Metal Oxide Separate Atmosphere Spectrometer Sensing Field Effect Analyzers for Constituents Transistor Each gas (0 – 5) where 5 is better Technical Viability 4 3 3 2 Financial 2.5 3 2 3 Feasibility Development 3 3 3 3 Program Risk Environmental 3 3 3 3 Impact Integration 4 4 4 2 Reusability Total (of a possible 16.5 16 15 13 30) Mass spectrometer has the highest score. However, combining approaches may provide the benefit redundancy for this critical function. 17.14. Temperature and Humidity Control (THC) Thetemperatureatwhichhumanscanoperate“comfortably”isintherangeof22to24oC. The appropriate humidity (RH) levels are between 25 and 75 percent. Excess heat (produced by electronics, lighting, solar heating and metabolic sources) and humidity need to be controlled to maintain suitable conditions inside the human module. The THC has to prevent condensation from forming on any windows, walls or equipment. (Purser, et al., 1964). Documented in the Shuttle Operational Data Book, JSC 08934, Section 4.6.1.3.1, p. 4.6.1-15, the nominal temperature extreme for 7 and 10 people ascent/entry phases is also given. For 7 people: 25 oC /16 oC and for 10 people 27 oC /16 oC. 17.15. Temperature and Humidity Control Method Description Advantages Disadvantages Atmosphere bypass Removes condensed Accommodates Degradation of control (Inside water from the nominal heat loads at coating due to Condensing Heat atmosphere in zero g the lowest specified exposure to Exchangers) set-point temperature contaminants, (18.3oC) Failure of coating Used in all modern causes carryover of day space habitats. liquid Low initial purchase Material of cost, Exhibit lower construction important fouling characteristics, as wall thickness is High heat transfer thin coefficients Gravity insensitive Long flight history Provides particulate
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and air borne control through HEPA (expandable) filters in the return air ducts Cold Plates liquid or atmosphere- Low operating costs, Heavy, risk of failure cooled “cold plates” elimination of in the presence of to which heat maintenance, and a vibrations generating equipment long lifespan are also is attached among the benefits. On-off Thermostat Detects temperature Low cost Atmosphere control changes and sends Available over a wide on space habitat is signal to heating unit operating range continuous to increase or Slow to react at decrease the temperature changes temperature Large Hysteresis range. Offer levels of imprecise levels of temperature control.
Proportional Integral Maintains the Easy to tune Offset-error must be (PI) Control atmosphere Provides good there for controller to temperature within the stability change output. specified limits for a Responds very rapidly Overshoot- range of heat loads. Dynamically stable oscillations can be undesirable There are several methods of removing excess heat; ventilation pulls the air through condensing heat exchangers removing heat energy. Alternatively liquid or atmosphere-cooled“coldplates” can be used to remove excess heat energy. Heat energy could then be transferred to external radiator, to transfer the energy into space. Evaporating liquid into space could also be used although this require a large amount of fluid and causes pollution.
Temperature Condensing Heat Cold plates On-off Thermostat Proportional Control exchanger Integral (PI) (0 – 5) where 5 is better Technical Viability 4 1 3 4 Financial Feasibility 3 2 2 3 Development 3 2 2 4 Program Risk Environmental 4 3 4 4 Impact Integration 4 3 3 4 Reusability 3 3 2 2
Total (of a possible 21 14 16 21 30) 17.16. Down Selection The condensing Heat exchanger serves many purposes in the ECLSS like humidity control filtration of air by the filters in the slurper bar. The optimum element would be to combine the condensing heat exchanger and the Proportional Integral.
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17.17. Ventilation Ventilation serves a variety of roles firstly it allows the constituents atmosphere to mix well. This enables the provision of sufficient O2 for metabolic requirement (If the RH is too low, the throat and nasal tissues become dry and if too high perspiration does not give proper cooling); and adequate removal of CO2, water and trace contaminants. Secondly it is the primary method of removing heat energy.
To avoid stagnant regions of gas where O2 levels may get too low or the CO2 level too high, there is a requirement for a specific ventilation flow rate. The ventilation flow rate must also meet the requirements for heat rejection to accommodate waste heat generated by people and equipment. The total pressure must also be maintained. Medical doctors at JSC have set face atmosphere velocity requirements for space habitats to a maximum of 0.20 m/s to prevent atmosphere drafts and a minimum of 0.08 m/s to control the CO2 level in the habitat. The use of auxiliary ventilation fans in the habitat will help engineers to meet these needs. (Wieland, 1994) 17.18. Equipment Cooling Method Description Advantages Disadvantages A single fan/heat System has supply and Offers lowest resource Can be challenging to exchanger system return ducts routed to and requirements ( power, balance the from all racks and/ or weight, and volume) atmosphere flows equipment bays properly with complicated distribution networks with varying heat loads Distribution Each fan/ exchanger tied Each separate package Past trade studies system with separately to the liquid can be sized to handle show that this various dedicated cooling loop for waste the expected heat load approach requires rack fans/heat removal offering flexibility and more resources exchangers eliminating the atmosphere balancing problem. Cold Plates liquid or atmosphere- Low operating costs, Heavy, risk of failure cooled “cold plates” to elimination of in the presence of which heat generating maintenance, and a vibrations equipment is attached long lifespan are also among the benefits. 17.19. Down Selection The table below shows how the possible solutions were marked against the Down Selection criteria allowing the ideal solution to be selected.
Temperature Control A single fan/heat Distribution system Cold Plates (0 – 5) where 5 is better exchanger system with various dedicated rack fans/heat exchangers
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Technical Viability 4 1 3 Financial Feasibility 3 2 2 Development Program 3 2 2 Risk Environmental Impact 3 3 4
Integration 4 3 3 Reusability 3 3 2
Total (of a possible 30) 20 14 16 From the table it can be seen that a single fan/heat exchange system would be ideal for this system. 17.20. Fire Detection and Suppression Wieland and Marshall define a FDS as a system that detects fires in a space habitat (closed volume accommodating powered equipment and an open area), alerts the crew by an alarm when a fire arises and provides a method to extinguish the fire. Requirements for the FDS system are as follows: It must be compatible with the ECLSS, non- toxic, and not produce toxic by-products. It is important that the detection of fires is without delay, so that crew will quickly take action to suppress the flame. The FDS elements chosen must be the quickest, most effective and safest. Smoke detectors must be positioned downstream to prevent the smoke from being filtered by the filter.
PY2 LSS- Fire Description Advantages Disadvantages 0 Detection Can see flame -When effectively able to -Senses can be impaired PY2 Crew and smell smoke use the senses humans are by sleep and tiredness 0-01 Senses excellent detectors -no cost UV radiation -Unaffected by solar -Blinded by thick emitted by fire radiation smoke, oil or grease on detected when -Unaffected by hot objects the lens (affects passed through -Low cost sensitivity), PY2 Ultraviolet UV glass and Hydrocarbon vapours, 0-02 Detectors strike a cathode. Chloride vapours Current formed -Subject to false alarms from UV sources -Reference fire at max 15m light-sensitive Detection distance Cost photocell Sensitivity monitors visible Speed of response PY2 Visible radiant energy Reliability 0-03 Detectors of a flame, flame flicker is monitored Monitor flicker Low cost Subject to false alarms Infrared PY2 of flame (0.7-5 (in the presence of detectors 0-04 um) Flickering IR sources) (IR) Will be blinded by fog,
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water and ice or salt layer on lens, sensitivity affect Poor detection and performance for stable flames Reference fire-15m max Monitors heat Linear over wide operating -low sensitivity generated by range -higher cost than Thermiters fire Wide temperature operating thermocouples and range -No point sensing PY2 resistance High temperature operating -Affected by shock and 0-05 temperature range vibration devices Interchangeability over wide -Requires three or four- (RTD) range wire operation Good stability at high temperature Air molecules -cheap -Smoke detectors will ionised by -Effective when properly run out of battery over radiation, maintained time if not wire to Ionization PY2 voltage moves electrical system Smoke 0-06 ions towards -Subject to false alarms sensors electrodes, due to interference current -Respond better to high generated flames detects smoke Simple susceptible to false using Reliable alarm due to dust obscuration accumulation Photoelectr PY2 detectors, light ic Smoke 0-07 scattering Detectors detectors and condensation nuclei counters
-Very low false alarm rate -Blinded by thick Dual -Unaffected by solar smoke, oil or grease on dete IR/UV radiation the lens ctors -moderate cost -Highest immunity to false -Higher cost Tripl alarms -Reference fire -65m e -Highest Sensitivity max (sprectrex-inc, n.d.) IR3 dete -Longest detection range ctors
17.21. Down Selection Area Crew UV Visible IR RTD Ionisation Photoelectric IR3 (0 – 5) where senses detectors Detectors Smoke Smoke 5 is better Sensors Detectors Technical 1 2 4 1 3 2 4 5 Viability Financial 4 3 3 2 3 1 4 3 Feasibility
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Development 1 2 3 2 1 1 3 2 Program Risk Environmental 4 3 4 3 2 2 4 3 Impact Total (of a 12 10 14 8 9 6 15 13 possible 20) The technologies with the highest scores are the Photoelectric Smoke Detectors. Triple IR (IR3) Flame detector scored poorly in the development program risk because it has not been used before in space habitats therefore will require testing before it can be utilised to detect. However, this technology scored very highly in the technical viability section. This technology will be ideal in the next 10 years as it does not get affected by interference. As crew senses are a bonus, they will be incorporated into the fire detection and suppression system. As the Photoelectric is susceptible to false alarms combining it with the human senses will improve this subsystem. 17.22. Suppression Advantages and disadvantages relate to the types of fires that can be extinguished, any toxic compounds produced by the suppressant, and cleanup after a fire, in addition to the basic factors of mass, volume, power consumption, safety, and reliability.
LSS- Fire Description Advantages Disadvantages PY21 Suppression From the Water readily available Not good for electrical fires PY21 food Water -01 rehydration gun Manually Removes combustion by- Loss of atmospheric gases opening products PY21 Depressuris cabin -04 e Habitat outflow valve Portable Used in Suppresses most types of fires Difficult to clean up PY21 Aqueous gel Skylab -03 extinguisher s (Foam) Distributed Space Highly effective at Halon produces toxic by- rack/ Shuttle and suppressing flammable liquid products that are difficult to subfloor Spacelab and electrical fires clean up PY21 mounted -04 Halon on bottles with distribution lines Can be removed from the CO2 is toxic at high levels PY21 CO2 atmosphere by the atmospheric -05 Extinguisher revitalisation system
Is the largest component of the Not as effective as halon PY21 atmosphere Low cooling potential Nitrogen -08 Inexpensive Heavy metal cylinders Compact storage space required when compared to
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Does not fog or impair vision halon on discharge Provides increased retention times to keep a fire extinguished Safer for people
17.23. Down Selection
Area Water Depressurise Foam Halon Co2 Nitrogen (0 – 5) where 5 Habitat Extinguisher Extinguisher is better Technical 1 3 3 4 3 3 Viability Financial 2 3 2 2 3 3 Feasibility Development 2 3 2 1 3 2 Program Risk Environmental 3 2 2 1 3 3 Impact Total (of a 8 11 9 8 12 11 possible 20)
Co2 Extinguisher has the highest score. This method is currently used on the ISS. This method can be used in conjunction with a nitrogen extinguisher as nitrogen will be readily available in the habitat as a major constituent. 17.24. Waste Management The MSFC and Johnson Space Centre (JSC) Space Station ECLSS personnel have identified the technology on the ISS to be state of the art. Key factors to be considered when designing a waste management system are the availability of power and storage volume. 17.25. Requirements: 1. Interface requirements including structural, electrical, fluid and odour requirements. 2. Performance requirements including quantities and types of waste to be processed. This also addresses the collection, processing and storage of waste, which may include faeces, wash water, EVA wastes, vomitus, menses and associated paper waste (tissue etc). Ensuring that contamination of the crew compartment does not happen is a major requirement. 3. Operational (flight and ground) requirements including procedures, man-systems aspects, noise level, and safety requirements. 4. Environmental requirements including temperature, pressure, RH, acceleration, and Vibration requirements.
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5. General requirements including leakage, maintainability and transportability.
Concepts to be considered simply collect and store waste with no further processing or involve limited processing such as dehydration to compact the waste (or to recover the water).
PY22 Technology Description Advantages Disadvantages Bags Sealable bags containing a No power Awkwardness of biocide taped to the buttocks required use Few reliability Waste has concerns potential to escape No atmosphere is bag before it PY22-01 lost due to sealed. dehydration to Crew must knead space the contents to Used as back up mix the biocide aboard ISS and with the waste soyuz Urinal e.g. Toilet functions as a vacuum No odour Crew members Soyuz toilet cleaner. The tanks are problems hold their bowls (See Image connected to a vacuum pump for two days. PY22-02 below) such that urine is sucked into the tanks through the filters and then separated.
Commode- -Urinal funnels specific for Little human Space Shuttle, male and female crew interference ISS toilet (See -Faeces are collected in the More comfortable Image below) commode storage container, and look like a where they are vacuum dried. conventional PY22-03 This kills bacteria therefore toilet reducing odour problems. - Solid waste goes into bags that are stored in airtight containers, until it can be disposed of. The WCS on the ISS will be adopted without the water recovery system. 17.26. Conclusion From the analysis of the various technologies for the ECLS subsystems would be assembled to construct the optimum ECLSS. The atmosphere control would be provided by a mass spectrometer. The atmosphere revitalisation would be provided by the Lithium hydroxide filters. The temperature and humidity control will be provided by condensing heat exchangers and the fire detection will be provided by the combination of photoelectric smoke detectors and human senses. The fire suppression would be provided by CO2 extinguishers. Water would be stored in tanks and food will be stored in foil packets and cans.
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18. Component Mass Estimation 18.1. Mass estimating Relationships (MERs) Mass estimating relationships offer the availability to model each sub-section of the proposed launch system through empirical formulas. These were derived by (Rohrschneider, 2002) using historical data from the NASA Space Shuttle to develop future conceptual launch systems. The equations are described in the appendix H with their associated input parameters. These parameters were calculated from the earlier sections in this report such as body diameter and engine vacuum thrust requirements. This provides an insight on how the reusable launch system would be broken down with each sub-section. It is to be noted that these mass estimates are purely conceptual estimates and would be validated by comparing to the mass predictions as stated in appendix H. Certain parameters that were required by these empirical formulas were not available in this conceptual study to which these were estimated through existing launch systems.
Sub-system Mass Estimates Mass (Kg) Tank Mass 730 Anti-vortex 56 Slash Baffles 60 Inter-tank 2,390 Inter-stage 1,720 Forward Skirt 1,615 Engine Compartment 2,660 Aft Skirt 855 Thrust Structure 15,130 Engines 76,440 Engine Install 4,450 Engine Thrust Vectoring 7,640 Engine Sub System 4,455 Propellant Purge 47 Feed Sub Systems 5,660 Avionics 460 Primary Power Group 1,615 Hydraulic System 265 Table 33 showing the mass estimates for stage 1 for both variants
Further mass estimates of all the stages for both variants are shown in appendix H. These mass estimates from the MERs provided a reasonable estimate compared to the derived mass estimates in appendix H with an error of less that 16%.
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19. Infrastructure
The infrastructure section would provide the necessary ground activities and buildings requirements for the proposed launch system. The infrastructure is to be designed around the conceptually designed launch vehicle as to provide the necessary support for the vehicle on the ground. It is to be noted that this section would provide the infrastructural implications on maximising the reusable launch system concept, other areas such as determining building requirements, test facilities requirements and propellant storage tanks requirements will not be mentioned as it does not contribute to the feasibility of the reusable system. 19.1. Launch Site 19.1.1. Launch Platforms [LF01] 19.1.1.1. Mobile Sea Launch Platforms – [LF01-01] A relatively unused type of launch platform is a mobile sea platform such as a converted oil rig. Due to the large expanse of sea, the location of launch can be optimised for the mission criteria. LaunchingalongtheequatorprovidesthebestlocationastheEarth’srotationprovidesanextra push to the vehicle allowing for less propellant to be used which in turn allows for a higher payload to be lifted. Furthermore, any orbital inclination can be achieved as the location can be varied to tailor the mission criteria. Due to this benefit, launch costs can be reduced for the mission. Moreover, due to the location of the launch out to sea, several benefits in terms of safety and risk can be described. Due to the large distances away from populated areas, failures during launch would not pose any risk to inhabitants. However, polluting the sea may have environmental concerns. The main disadvantage is that the sea platform itself has to sustain the weight of the launch vehiclefromthebearingsea.Thecurrentseaplatform’slaunchessmallsatellitesintoGEOwhile the current proposed launch vehicle would be several magnitudes larger than the current sea launch vehicles. The mobile sea launcher would have to be unfeasibly large to cope with the weight of the proposed launch vehicle. Furthermore, transportation and assembly on the sea platform may prove problematic against the tides and weather. Other sea launch platforms include launches from a submarine however; the size of the submarine would be greatly unfeasible for the size of the launch vehicle. 19.1.1.2. Missile Silos (Land based launch platforms) – [LF01-05] From the cold war era, a number of missiles silos existed where launch vehicles could be launched allowing for a pre-build launch pad. These offered an enclosed space where the exhaust gases from the engines could be vented and controlled. However, from the size of the proposed
William Wou Section 19 - Infrastructure Page 129 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 vehicle, the implication of the size of the missiles silos would be unfeasibly large due to the amount of excavating required for the volume of the launch vehicle. 19.1.1.3. Spaceport (Land based launch platforms) – [LF01-02] The spaceport is a facility capable of launching and receiving launch vehicles with the possible inclusion of runways for aircraft, the most famous spaceport being Cape Canaveral in Florida. Although these are fixed launch sites, they have been traditionally placed at existing military installations. The main consideration when choosing a location for a spaceport is the distance from populated areas or near an ocean from the risk of rocket failure. This restricts the location of the spaceport from the possible optimum launch locations. Further criteria such as transportation and utilities availability play a big role in choosing a spaceport's location. However, this form of launch platform offers the most realistic approach for development and operational use. 19.1.2. Launch Tower The launch pad is there to hold the launch vehicle in a vertical position and to allow for refuelling and pre-launch activities. The Russia Agency's launch pad holds the launch vehicle under four trusses that hold the weight of the vehicle. When the vehicle launches and the weight of the vehicle reaches equilibrium from the thrust force, the trusses fall apart due to counter-balances. NASA's method requires explosive bolts which holds the launch vehicle down when fuelling that are subsequently controlled detonated on launch. 19.1.3. Transportation of vehicle to launch pad and orientation of vehicle assembly [LF02] The method of transporting the assembled vehicle to the launch pads has two implications under the discussion below. The first being the building size requirements due to how the vehicle is to be transported and the time it takes for it to move to the launch pad. The vehicle must be moved to the launch pad as fast as possible due to how the reusable launch system is to be operated. 19.1.3.1. Mobile crawler-transporter (Vertically Assembled) – [LF02-01] Assembling the vehicle vertically upwards as traditionally done by NASA, results in tall assembly buildings that are able to move stages around under ceiling cranes. This assembly method allows for the rocket's structure to only withstand the loads generated from its own weight from nose to tail which reduces the structural mass of the rocket. However, the process of assembling the launch vehicle in a vertical manor can be complicated due to stages dangling from ceiling cranes in the assembly building. Furthermore a mobile crawler transporter is required that is capable of moving the entire vehicle from the assembly building to the launch pad vertically which causes problems when high gusts shift the top heavy vehicle around. 19.1.3.2. Mobile crawler-transporter (Horizontally Assembled) – [LF02-02] Assembling the launch vehicle horizontally as traditionally done by the Russian Space Agency results in buildings that utilise significant plots of land but is convenient since the assembly building has a common height. However, this method causes the structure of the launch vehicle to
William Wou Section 19 - Infrastructure Page 130 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 be much stronger as it has to withstand its own loads horizontally. This results in the structure of the vehicle to be more robust but this method allows for simpler assembly since the stages can be moved around on the ground by stands on wheels. Once assembled, powerful jacks are required to lift the launch vehicle vertically once it has been moved to the launch pad. 19.1.3.3. Modules transported to launch pad and assembled – [LF02-03] Moving an assembled launch vehicle from the assembly building to the launch pad is rather slow and difficult. Another method is to transport the individual stages to the launch pad where they are assembled on the launch pad itself. This reduces the time and difficulty moving a large top heavy launch vehicle on a mobile transporter. However, the assembly building is located next to the launch pad which results in the assembly building having to be moved backwards away from the exhaust gases from the launch vehicle. 19.2. Infrastructure Requirements The type of launch mode that the vehicle requires dictates the infrastructure requirements due to the necessary actions and buildings required for launch, recovery and landing. These can be broken down into vertical take-off – vertical landing (VTVL) such as a traditional rocket system, vertical take-off – horizontal landing (VTHL), horizontal take-off – horizontal landing (HTHL) such as a space plane. Table 34 shows the infrastructure requirements for the various launch types.
Launch Type Infrastructure requirements Vertical Take-off – Vertical Landing (VTVL) -The launch pad has to be restricted for launch -Turn around depends on recovery procedure -Issues on safety due to vertical landing Vertical Take-off – Horizontal Landing -The launch pad has to be restricted for launch (VTHL) -Long runway required for landing -Requires a launch pad and runway (expensive) Horizontal Take-off – Horizontal Landing -Long runway with a high strength concrete to (HTHL) cope with the weight and the launch thrust Air Launch -Specialised custom aircraft required -Possible to use existing airport infrastructure for launch -Transpiration infrastructure required to move rocket plane Table 34 showing the infrastructure requirements for different launch types
It is to be noted with a VTVL that the turnaround procedures depends on how the vehicle is landed and its location of landing. Furthermore, inspection and maintenance of these vertical landing vehicles might be high due to rough landings or sea landings. A parachute recovery near the infrastructure may incur a relatively high turnaround time; however, a parachute recovery into the sea would increase the turnaround times due to maintenance and transportation.
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Moreover,anairlaunchwherearocketplaneis“piggybacked”onanotheraircrafthasseveral advantages in terms of its infrastructure requirements. It allows the main aircraft that uses traditional jet propulsion to take-off from existing airports since the rocket plane is not released until reaching a desired altitude away from any airport. However, it requires certain equipment for daily operations, similar to those needed for the space shuttle since it could not fly between airports under its own power. The current intended design incorporates the VTVL launch mode; this system would require two different launch sites, one for launching the vehicle and another for recovering the stages which would automatically land under its own guidance computers. As stated in Table 34, both of the launch pads would be restricted from personal during launch and recovery for safety. Furthermore, analysis would be conducted to determine the recovery time caused by the recovery of stages. 19.3. Infrastructure designed around a reusable launch system The primary influences that cause high launch costs are due to the operations, maintenance and the infrastructural requirements by the project over its life time (Charania, et al., 2004). As stated earlier, the infrastructure is to be designed around the reusable launch system as to maximise its potential benefits through the reduction of recurring costs due to the ground operations. One such method of modelling the new infrastructure is by using the airline and airport industry as a guide to efficiently map how to turnaround reusable launch systems (Fox, 2001). Like the airport industry, the infrastructure for the launch vehicle must be designed as to allow for multiple launches, maintenance checks and assemblies simultaneously. This provides an increased number of launches per year allowing the launch vehicle to run like an aircraft in an airline manor where the turnaround maintenance time and cost is the deciding factor in profitability.
A large portion of the recurring cost is the maintenance that the launch system requires after each mission. The launch system must be recovered from its landing position and processed as quickly and as safe as possible to a high standard, before it is loaded with its payload and launched again. This is where the system gains its strength by launching the vehicle as rapidly as possible since profits are gained by launching customer's payloads into space. A manufacturing technique in terms of a production line is to be proposed to allow the stages of the rocket to be processed as efficiently as possible. The layout of this infrastructure is shown in Figure 23 as a method of describing this technique. This production line is used due to the modularity of the launch system as it allows each sub-section of the system to be broken down and serviced together. This allows for maintenance lines that have common components that filter
William Wou Section 19 - Infrastructure Page 132 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 down receiving maintenance till they are deemed flight worthy where they can be assembled into the full variant launch system. The maintenance building would to be sectioned off into four sections for each sub-section of the launch system. The first three being the three stages from variant-A and B and the fourth being the engines from all of the stages as they are common in both variants. All of the components from the system will be inspected, maintained and served in a horizontal position down their respective maintenance line where they are to be mated with equipment and worked on by ground staff. The stages and engines would be moved down the maintenance line by small robotic carts and placed on stands providing access to the whole body of the vehicle. Each stage and engine would have a stand which holds the component horizontally in a truss like structure that can be manoeuvred under electronic drive units, controlled by the ground crew. The full process of this will be discussed in the operations section under maintenance. Due to said operational cost, it has been evaluated that the launch vehicle would be assembled horizontally and transported to the launch pad to where it would be moved to a vertical position for launch. This has been chosen as it allows for a faster assembly turnaround with multiple assemblies occurring simultaneously within the assembling building. A vertical assembly would require multiple heavy cranes causing assembly times to increase. Furthermore, a Russian styled launch tower would be used as it provides a relatively fast method of preparing for another launch on successive missions.
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19.4. Infrastructure Layout Figure 23 shows a conceptual schematic of the proposed spaceport maintenance building necessary to maximise the reusable aspect of the launch system.
Figure 23 showing a conceptual Stages moved into Stages Land on maintenance building schematic (Not to Scale) maintenance building Recovery Pads
Stands and equipment mated to Launch vehicle broken each stage for servicing down into sub-sections
First Stage Maintenance
Arrow showing the direction of work flow, from recovery to launch Engine Maintenance pad. Stages move down maintenance lines after successive servicing
Second Stage Maintenance
Payload preparation bay Third Stage Maintenance
Assembly of the launch vehicle and pre-launch preparations
Stages are moved under rocket stands allowing for free movement
Vertical Launch Pad
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19.5. Developing New Launch Sites from Existing Launch Sites – [LF03-01] The infrastructure is an important element to the reusable launch system as it provides a framework and support necessary for each mission. As stated earlier, the infrastructure must be designed in such a way as to reduce the operating cost over the course of the project life. This cannot be achieved by utilising current infrastructure such as from Cape Canaveral. As part of the investment plan for the infrastructure, a new spaceport is to be designed and built around the launch system to maximise the reusability aspect of the launch system. The new infrastructure would have similarities towards the airport and airline industry as they have developed the necessary technology and processes to efficiently maximise the reusability of their vehicles to reach a maximum in their profitability. The infrastructure for the launch system would be a major investment for the project. A cost estimating technique was used to extrapolate historical data from the Space Shuttle program in terms of facilities and ground support cost. This was based off the size of the launch vehicle, size of the integration stages and payload bay sizes using techniques employed by (Zapata, n.d.). Table 35 shows the cost estimates for the proposed reusable launch system in terms of facilities and ground support cost as an investment cost.
Facilities Ground Equipment Payload Processing $38,000,000 $84,000,000 Flight Control/Traffic $6,000,000 $87,000,000 Launch Pad $763,000,000 $763,000,000 Recovery Pad $84,000,000 $88,000,000 Vehicle Turnaround $57,000,000 $381,000,000 Vehicle Assemble/Integration $110,000,000 $69,000,000 Vehicle Maintenance $265,000,000 $465,000,000 Supporting Infrastructure $265,000,000 $465,000,000 Concept Unique Logistics $265,000,000 $465,000,000 Management/Operations $265,000,000 $465,000,000 Table 35 showing the cost estimates for the infrastructure in terms of facilities and ground equipment The total investment cost into the infrastructure is estimated at $5,450,000,000 which includes the facilities and ground equipment cost. During operational service, the infrastructure would be continually upgraded with new technology and procedures to maximise the launch system. 19.6. Operations From the infrastructure section, the operations of the reusable launch system would be designed for the proposed reusable launch system as to provide necessary information on how the system woul be utilised and designed. The main emphasis is on the optimum operational method for the intended launch system to determine whether this is feasible through these methods.
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19.7. Maintenance Operations 19.7.1. Stage Recovery Each stage of the launch system is able to land vertically under its own engines onto the recovery pad next to the maintenance and assembly building. Once landed, the vehicle would be left to cool down for a period of time before a ground crew do a walk around checking for any glaring issues with each stage. Once cleared, an automated recovery vehicle would lift the stage into the maintenance hangar for servicing. The stage would be orientated horizontally and placed onto a stage stand where it can be manoeuvred by ground staff using electric motors. Initially the tanks would be purged to remove the contents from the tanks in a controlled space before entering the maintenance line. Once entering the maintenance line, the engines on each stage will be removed and moved to the engine maintenance line. This allows for common sub-sections of the launch vehicle to be processed efficiently and as quickly as possible.
19.7.2. Stage Maintenance Each stage and engine in the four maintenance lines would be mated with equipment, ground staff and cranes where maintenance can be conducted. Floors above the stage will also be mated as to allow ground staff access to the entire body. Along the maintenance line, there will be stations where a particular activity will occur such as inspection or replacing of components. Once the ground crew have finished with that particular stage, the stage will be moved down the maintenance line to the next station. If stages are deemed unserviceable or require extra servicing, they will be moved off the line into another line for heavy maintenance. This is to stop any disruption in the flow of work which would hamper the maintenance line.
19.7.3. Launch System Assembly After the stages have been deemed flight worthy, the stages will be assembled into one of the two variants by moving the stage stands together. Once all the stages have been mated, an inspection and test team will conduct the necessary flight safety tests before the payload is attached to the vehicle. If the launch vehicle is carrying crew, the crew would be inserted into the vehicle at the launch pad. The assembled and tested launch vehicle would be moved by rail to the launch pad where it would be vertically raised by powerful jacks onto the launch tower. Fuelling and necessary pre- launch activities will be conducted before the crew enter the vehicle for crewed missions.
19.7.4. Component life-time The components of the reusable launch system have a limited life-span as they have to withstand the harsh environments of launch and space. During the maintenance, the vehicle's old components will be serviced with newer components once they have reached their mission life- span. Each component has a flight rated life time of x amount of missions. Three main areas of the vehicle have been identified to be replaced, these being the structure of the vehicle, the
William Wou Section 19 - Infrastructure Page 136 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 engines and the sub-systems including the electronics and avionics which will be used in the turnaround modelling later in this report. This continual change of components for each launch vehicle allows the vehicle to last over the entire project, the project life time being 30 years. This is because each sub-system of the vehicle will be serviced and changed once it has reached its life-span. This allows for only a few vehicles to be manufactured, meeting the market requirements as each of the reusable vehicles are serviced and maintained. A vehicle produced at the start of the project would be an entirely different vehicle after the project life of 30 years as all the components will have been overhauled. From Table 36, each component has a certain mission lifetime. Variant-A carrying human crew will have a higher amount of component replacement to allow for safer missions. The development cost for each of these sub-systems will verify the reliability during test activities (Pearson, et al., 2000).
Number of missions each component can last before servicing Sub-system Variant-A Crewed Variant-A Unmanned Variant-B Unmanned Engines 15 20 20 Structure 25 30 30 Sub-systems 45 50 50
Table 36 showing the mission space for each major component for the different variants From the component lifetimes as shown above, these would be used to determine total amount of component changes throughout the whole project. This would provide a total cost for replacing the components for the whole fleet over the 30 years which will be factored into the reoccurring costs in the finance section. 19.8. Automated Monitoring Systems As stated earlier, the maintenance cycle for the vehicle is critical for its operation. Several methods for reducing the maintenance cycles are to be discussed below:
19.8.1. Integrated Health Management System (IVHM) Due to the large surface area and complex intrinsic sub-systems, inspection of the vehicle would be a lengthy process through visual inspection by ground staff. Inspection historically from the Space Shuttle took a tremendous amount of time to validate its safety especially with the heat shield. An ideal solution would be to have the vehicle identify areas that require maintenance before the vehicle touches down. The ground staff would be organised with equipment and spares before the vehicle enters the maintenance building. This system is recognized as Integrated Vehicle Health Management (Baroth, et al., 2011). This system aims to produce three benefits to the launch system. Firstly, this system increases the safety and reliability of the launch vehicle as it provides detailed information about each sub- section to ground staff. This information uplink allows for components of the vehicle to be
William Wou Section 19 - Infrastructure Page 137 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 changed before they fail once they reach certain tolerances during operations. Secondly, ground maintenance cycles are reduced as the vehicles inspect themselves allowing for a more automated ground servicing and checkout. Finally, the system allows more automation during operations with information data linking to ground controller team which reduces their work load (Jambor, 1997).
19.8.2. Informed Maintenance (IM) Informed maintenance (IM) is a subset of IVHM and is another method of reducing the work load for ground crew. It has several aims to optimise system maintainability and performance. Informed maintenance is a system created by NASA which allows for testability examination, automated maintenance organisation, automated logistics management, paperless documentation and automated diagnostics (Fox, 2001). Both these systems are to be used by ground crews to reduce the work load and increase efficiency during maintenance.
19.8.3. Automated System Checkout (PCCS) A system developed by NASA called the Propulsion Checkout and Control System (PCCS) is aimed at predicting the condition of the launch vehicle than solely relying on scheduled maintenance checks (Patterson-Hine, et al., 2001). The system is able to reduce turn-around time and cost by using simulation models than by using conventional maintenance techniques. The simulation models use real-time sensors, validation; limit checking and failure prediction which feed into IVHM for overall visibility of the launch vehicle's system state. All three of these systems will be integrated into the infrastructure and operations before operational service at a projected cost of $200,000,000. 19.9. Turnaround Maintenance for a reusable launch system The reusable launch system derives its cost effectiveness through its ability to be reused again as it is not discarded after every mission. The cost and time of this maintenance is one of the main driving forces for its feasibility as it produces recurring costs that are dictated through its maintenance. The ability to recover and maintain the system to a launch worthy state in the minimal amount of time and cost allows the recurring operating cost to be significantly reduced allowing for a feasible system. This is because the vehicle generates its profits by missions sending up payloads into space for customers. A similar system is the airline industry (Sethi, 2012). The airlines only generate cost when their passengers are in the air. While the aircraft is under maintenance, the airline is losing money due to the servicing cost (Jiang & Hansman, 2006). One of NASA's problems with the Space Shuttle was the large recurring operating cost from its maintenance requirement of 87 days with a total turnaround time of 159 days (assembly to launch) (Andrews, 2011) due to its complex embedded systems.
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The length of the maintenance on the launch vehicle also dictates the number of reusable launch systems to be produced. With a longer maintenance cycle, an increased amount of launch vehicles is needed to be produced to satisfy the number of launches per year. This because, the vehicle is grounded under maintenance requiring another vehicle to be built for the next mission. Reducing the size of the fleet is necessary to allow the production cost over the course of the project to be at its minimal as it produces a cost effective cycle. An equivalent expendable system like the Soyuz system requires a new vehicle to be built for each mission causing high recurring production costs each year. The reusable launch system in theory should be built once and maintained between each mission to a launch worthy standard where components and sub-systems are replaced. This replacement of components such as the structure and engines would be significantly lower than producing an entirely new system. Although the reusable launch system would require a higher production cost than an equivalent expendable launch system, the reusable launch system's maintenance cost is significantly lower than the expendable launch system's production cost. In order to reduce the recurring operating costs, the vehicle must be designed to allow ground staff to inspect the vehicle and replace parts where necessary efficiently. Through the proposed integrated system monitoring system, the ground staff are able to identify areas where more attention is required before the vehicle's stages touch down onto the pad. This allows ground staff to efficiently organise maintenance tasks in terms of spare parts and equipment to stages that require special maintenance activities. This reduces the inspection time required by ground staff for each stage as the vehicle is able to inspect itself. Furthermore, the vehicle will be designed from an operational standpoint to maximise its reusability and maintainability. Using techniques from the aircraft manufacturing industry, the design of the vehicle will allow ground staff access to all areas of the vehicle that will be tested through virtual reality that aircraft manufactures use today (Staton, 2002). For example, the numerous engines on the first stage will be designed to be easily removed and attached to the vehicle's body as to reduce the time for maintenance. Moreover, due to the commonality and modularity between the two variants in terms of their first stage, ground staffs are able to swap these stages between different variants if a stage is deemed unserviceable or requires longer than expected maintenance. The overall maintenance of the launch system must not be interrupted by unforeseen circumstances as it risks cancelling the mission and thus loses its profitability. The operations of the maintenance must flow seamlessly as to ensure each launch system is ready for the next mission. 19.10. Turnaround Maintenance Modelling Modelling the turnaround time for the reusable launch system allows for estimation into the recurring costs and the length of the maintenance. The full methodology behind this modelling is described in appendix H.1.2. The main areas of the launch vehicle that were considered were the
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thermal protection system (TPS), engines, structure and sub-systems which included the electronics, hydraulics and avionics etc (Snead, n.d.). Variant-A carrying crew had a 20% increase in maintenance servicing as a level of safety for crewed missions. Furthermore, with IVHM, IM and PCCS, the maintenance time are expect to decreases by 35%.
Table 37 shows the estimated maintenance time for each variant and their stage. The estimates are conservative estimates to which a sensitivities analysis will be conducted later in the report.
Stage 1 Stage 2 Stage 3
Systems Variant-A Variant-A Variant-B Variant-A Variant-A Variant-B Variant-B Crewed Unmanned Unmanned Crewed Unmanned Unmanned Unmanned
TPS - - - 260 hours 240 hours 240 hours 180 hours
Engines 2,560 hours 2140 hours 2140 hours 860 hours 710 hours 2140 hours 240 hours
Structure 4789 hours 2430 hours 2430 hours 3505 hours 1770 hours 1650 hours 1340 hours
Sub-Systems 4200 hours 3500 hours 3500 hours 4200 hours 3500 hours 3500 hours 3500 hours
Table 37 showing the estimated maintenance time for each system It is to be noted that stage 1 for all the variants do not require a TPS as they do not re-enter through the atmosphere. Furthermore, the sub-systems have been treated as constants from (Snead, n.d.) due to the lack of information on sub-systems for this study. Due to the proposed maintenance line as specified in the infrastructure section, the work load by the ground staff is split down the four maintenance lines for each stage and engines. This allows the maintenance time to be decreased than working on each system one by one. Table 38 shows estimates for the total cost of maintenance including the cost for replacement components.
Variant-A Crewed Variant-A Unmanned Variant-B Unmanned
Estimated total maintenance time 42 days 38 days 54 days Estimated reoccurring maintenance cost $78,000,000 $47,500,000 $61,000,000
Table 38 showing the estimating maintenance time and cost for all 3 variants 19.10.1. Fleet size Due to the reusable aspect of the launch system, the fleet size is dictated by how long the maintenance cycles are. With variant-A (crewed) taking 42 days to complete its maintenance cycle and with a proposed 22 missions per year for this variant, a fleet of 3 vehicles would be required to satisfy the launch frequency. This of course, assumes that no problems will occur during the maintenance of the vehicle and the component will last till the stated component life- time. The full fleet size for all variants are shown in appendix H.1.5 as it is a function of the mission per year set by the market forecast.
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20. Finance
The Financial Feasibility study was conducted through two methods. The first is the comparison between an equivalent Expendable Launch System (ELS) to the proposed Reusable Launch System (RLS) to determine whether the reusable mechanism of the launch system is viable. The second method was comparing the reusable launch system to the current and past markets of launch systems to determine whether the system is competitive for potential customers to use.
Although most launch systems do not provide profit for the organisation that operate the launch vehicles, the proposed reusable launch vehicle must be able to demonstrate potential either by some return from each launch or by showing a loss penalty for each launch that does not significantly hamper a given organisation. This is to be determined by using the market forecast in the marketing section to simulate operational use for both the systems (ELS and RLS) over the project life of 30 years. Calculations of these costs used were generated by CER (Cost Estimating Relationships) which used the system's mass as a function to generate its cost. These were derived from using parametric analysis that used historical data to predict future launch systems by (Koelle, 2002). The equations are listed in appendix H.1.3. The cost estimates are generated by historical costs from similar launch systems with complexity factors added for safety. A level of effort factor of 8% is added for program management and 15% is added for system engineering and Integration (Rothschild & Talay, n.d.). To mitigate risk, a contingency factor of 15% and a DDT&E (Development, design, testing and evaluation) factor of 10% are added to the recurring costs. It has also been assumed since the CER were derived, there has been a 40% technological improvement to the development and production cost (Snead, n.d.). 20.1. Launch Cost The launch cost for both RLS and ELS is to be determined as the total operating cost for each mission. The market forecast provides prices for each mission as payment to the organisation operating the launch system using current based vehicles. From this information, the revenue from the operations can be determined over the project life span. It is to be noted that the market forecast uses current system prices; the proposed launch system will be operated under those conditions as to determine whether the system generates profit. The launch cost is broken down into smaller defined costs which are derived from (Wertz, 2004) as shown below.
The development cost ( ), although not a recurring cost, is to be divided every year over the 30 year project life time. This allows the launch system's development cost to "bought back" if the launch system generates any profits.
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The production cost ( ) for the ELS is a reoccurring cost as a new launch vehicle is required for every mission. For a RLS, this is a non-reoccurring cost as the launch system can be reused.
The recovery cost ( ) and refurbishment cost ( ) is to be taken together as the reoccurring maintenance cost for the RLS. The ELS does not have these cost attributed to it. The cost of replacing all of the components for the RLS is split over the 30 year lifetime. Therefore, each launch cost takes into account cost for replacement of components for the RLS.
The flight operations ( ) and insurance cost ) is to be neglected as they do not alter to the Financial Feasibility study. Furthermore, a learning factor of 0.85 will be used (Keith, n.d.) to simulate development advancements. 20.2. Development costs The development costs in Table 39 take into account production of a prototype vehicle and test evaluations during the development phase. It is to be noted that variant A and B share the same stage which results in the same development cost for each.
Stage 1 Stage 2 Stage 3 Variant-A Variant B Variant-A Variant -B Variant-B
RLS development $14,600,000,000 $9,400,000,000 $6,360,000,000 $5,500,000,000
Cost
Table 39 showing the development cost for the RLS with both variants
The total development cost for the RLS is $35,860,000,000. This development cost is verified by Figure 37 in appendix B.5 which shows correlated historic development cost for various launch systems as a function of total launch mass. The figure estimates the development cost to be approximated $37,500,000,000 which verifies the development cost using the CER.
The equivalent ELS is shown in Table 40 using the same mass estimates from the RLS.
Stage 1 Stage 2 Stage 3
Variant-A Variant B Variant-A Variant-B Variant B
ELS development $6,140,000,000 $3,870,000,000 $2,550,000,000 $2,200,000,000
Cost
Table 40 showing the development cost for equivalent ELS for both variants
The total development cost for the ELS is $14,760,000,000 which is 59% cheaper than the equivalent RLS which is expected as shown by figure 33 in appendix B.5.
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20.3. Production costs Table 41 shows the production cost to build one RLS. Stage 1 Stage 2 Stage 3 Variant-A Variant B Variant-A Variant -B Variant-B
RLS production Cost $282,600,000 $89,100,000 $154,000,000 $72,400,000
Table 41 showing the development cost for the RLS with both variants
The production cost for Variant -A and Variant-B is $371,000,000 and $509,000,000 respectively. The equivalent ELS is shown in Table 42 using the same mass estimates from the RLS.
Stage 1 Stage 2 Stage 3 Variant-A Variant B Variant-A Variant-B Variant B
ELS production Cost $138,500,000 $43,600,000 $75,500,000 $35,500,000
Table 42 showing the development cost for equivalent ELS for both variants The production cost for Variant -A and Variant-B is $182,000,000 and $249,500,000 respectively. The equivalent ELS for both variants are approximated 50% cheaper to produce than the RLS. 20.4. Variant-A Crewed Missions Variant-A is to be separated from the unmanned missions as higher levels of safety are added onto the maintenance, development and production cost. Figure 24 shows the launch cost for the RLS and equivalent ELS. The market forecast showing the number of missions per year in purple with its corresponding secondary axis on the right of the graph.
Crewed missions for a reusable and equivalent expendable launch system 30 6,000
25
5,000 20 4,000
15 3,000
2,000 10
1,000 5 Number Missions of peryear
0 0 Launch Launch (US Cost Dollars (Millions)) 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 Year Expendable Reusable Missions per year
Figure 24 showing the RLS and equivalent ELS over the project life time undergoing manned missions
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The operational model shows that the ELS's (blue line) reoccurring production cost causes the launch cost to rise with the number launches per year rate. This is because the ELS is manufactured each time a mission is required. The RLS (red line) shows how the launch cost reduces with an increased number of missions per year. This is due to the reusability aspect of the launch system. The launch system only needs to produce a vehicle at the start of its operational lifetime and when an extra launch vehicle is required to meet the demand specified by the market forecast. This is shown by the "jumps" in the launch cost for the RLS as certain years requires a new vehicle to meet the demand which incurs high RLS production cost causing these "jumps". The RLS's reoccurring operating costs are far lower than the ELS's reoccurring production cost which allows the RLS to benefit after year 2034. This is due to the doubling of missions per year as space tourism is forecasted to increase due to the opening of space hotels and the ISS. The RLS benefits from the higher launch frequency as the system becomes more economical with higher launches per year.
It is to be noted that, during the first four years of the operational model, the RLS has a higher launch cost than the equivalent ELS. This is due to the RLS being more complicated to assemble and integrate which results in an overall higher pre-launch preparation cost than the ELS. Figure 25 shows the revenue by the manned missions set by the market forecast. As mentioned earlier, the RLS has a period from 2030 to 2034 where it is more expensive than the ELS due to pre-launch activities. This is shown in Figure 25 as a higher loss during those four years for the RLS (blue).
Cash flow for a manned missions for the reusable and expendable system 16,000 14,000 12,000 10,000 8,000 6,000 4,000 2,000 0
-2,000
2,030 2,032 2,031 2,033 2,034 2,035 2,036 2,037 2,038 2,039 2,040 2,041 2,042 2,043 2,044 2,045 2,046 2,047 2,048 2,049 2,050 2,051 2,052 2,053 2,054
Launch Launch (US Cost Dollars (Millions)) -4,000 -6,000 Years Reusable Crewed System Expendable Crewed System
Figure 25 showing the cash flow for Variant-A-Crewed with reusable and expendable launch systems Once the number of missions per year double after year 2034, both systems make significant profit with the RLS producing the highest revenue when compared to the ELS.
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With a 20% margin for profit, a new price for cost per seat for the ELS and RLS has been made with the figure below. As mentioned earlier, the market forecast uses existing launch systems to estimate the typical launch price for different missions. Figure 26 uses its own launch cost with the stated profit margin to determine its own cost per seat for both systems. It is to be noted how both systems meet the competitors cost per seat after year 2025 with the RLS producing the lowest cost per seat between the two systems. This can further be shown by Figure 54 showing the RLS crewed vehicle against leading competitors in terms of cost/kg.
Comparing Cost Per Seat for manned launch systems
250
200
150
100
50 Cost/Seat Cost/Seat (USDollars (Millions)) 0 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 years Expendable Reusable Soyuz Space Shuttle Space X
Figure 26 showing the cost per seat for the reusable launch system and equivalent expendable launch system compared to other launch systems from (SpaceX, n.d.)
20.5. Variant-A and Variant-B for unmanned missions The cash flow for Variant-A shown by Figure 27, shows a typical trend for the RLS as it is able to produce an appreciable profit rise after year 2030. This is due to the projected increase in missions for the LEO satellites and Space Station supply missions. However, the ELS generates significant loses over its service life. This is due to the estimated launch price from the market forecast which used existing system's launch price. The launch price is not sufficient enough to break-even which causes the considerable loss. Furthermore, the cash flow for Variant-B can also be shown below in Figure 28. Again, the trends are similar for both systems. The RLS is able to steadily increase in profits after year 2030 due to the forecasted start of space infrastructure missions. The ELS's launch price is notably higher than the forecasted launch price from the market section causing the resulting loses over its operational service life.
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Cash flow for variant-A unmanned missions
1300
700 100 -500 -1100
US Dollars US (Millions) -1700 -2300 2024 2026 2028 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054
Reusable Expendable
Figure 27 showing the cash flow for variant-A for both RLS and ELS.
Cash flow for Variant-B unmanned missions 4000
2000
0
-2000
-4000
US Dollars US (Millions) -6000
-8000 2024 2026 2028 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 Years Reusable Expendable
Figure 28 showing the cash flow for Variant B for RLS and ELS
20.6. Financial Feasibility of reusable system The proposed reusable launch system simulated through a forecasted market model shows how initially over the first ten years of operational service (Figure 29), the reusable launch system loses profit in the "valley of death" period of the project. After year 2035, the project significantly gains in profitability as the RLS is able to reduce its low reoccurring operational cost with higher missions per year. The investors are able to notably gain back from their initial investment from the development costs with a 34% gain in profits over the project life span. This is further shown by the RLS generating a reasonable internal rate of return (IRR) of 11% while the equivalent ELS only produces 2% due to the significant loses for Variant A and B for unmanned missions.
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Cash flow for the reusable launch system for all variants 20000
15000
10000
5000
0
2024 2026 2028 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 US Dollars (Millions) Dollars US -5000
-10000 Years
Figure 29 showing the cash flow of the entire project for both reusable launch systems
It is to be noted that although the losses made during the first 10 years are relatively high, the figure above includes the development cost, production cost and total component replacement cost stretched over the 30 years for the proposed reusable launch system. This stated IRR shows the major comparison between the two systems. IRR is defined as the discount rate for a given project that produces a zero net present value (i.e. higher IRR is better whilst neglecting risk). From the statement above, the RLS generates this higher IRR value than ELS due to the relatively low reoccurring operating costs attributed to the RLS. This shows that the RLS is able to provide positive economic return over the project life time when compared to equivalent ELS satisfying the first financial feasibility criteria. This is because the ELS cannot earn back on its initial capital. This ultimately demonstrates the importance of the reusability mechanism in the launch system. Moreover, the profits rise significantly due to the constant launch price used from the market forecasting section; this does not take into account of the learning factor reducing the launch cost. This allows for a reduction in launch price for customer's using the RLS which would flatten out the above figure with a certain profit margin. The second financial feasibility criterion was to determine whether the RLS is capable of competing against the current market of launch systems. Figure 30 shows the launch cost for the proposed reusable launch system with both variants, where it is able to meet the competitor's launch cost at year 2030. It is to be noted that the competitor's launch costs is taken from when this report was created. Once the proposed RLS is operational, the competitor's launch cost would be a degree lower than as specified. However, the RLS continues to reduce its launch cost where it has a good opportunity to meet its competitor's price for customers.
William Wou Section 20 - Finance Page 147 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Launch Cost For all Variants Against Current Market 1200
1000
800
600
400
US Dollars (Millions) Dollars US 200
0 2024 2027 2030 2033 2036 2039 2042 2045 2048 2051 2054 Years Group 2 Space X ULA
Figure 30 showing the launch cost of the RLS for both variants, data from (SpaceX, n.d.).
The financial study above proves that a reusable launch system is not only significantly beneficial to an equivalent expendable launch system but is able to meet current based competitor's launch vehicles due to its reusability. The proposed reusable launch system is a financially feasible launch system which is also shown by the current interests by companies today. 21. Sensitivity Analysis
The main variable for the reusable launch system that controls the launch price is the maintenance and replacement of component in terms of cost and time. A longer maintenance cycle results in a higher production cost as more vehicles are needed to meet the launch frequency. Furthermore, components with a predicted lower life span results in higher component turnaround costs. Both of these contribute to the reoccurring costs for the RLS which define the system. A study was conducted by taking a 25% increase and decrease of these reoccurring costs to determine whether it has a contributory effect to the launch cost. The results shown in appendix H.1.6 shows that there is an approximate 20% change in the overall launch cost by altering the total reoccurring cost by 25%. As predicted, the fleet requirement must increase due to longer maintenance times causing significant jumps in the launch cost in certain years due to new launch vehicles being produced. Overall, from this sensitivity analysis, there appears to be no large changes in the launch cost that warrants a different decision on the feasibility of the reusable launch system. This is due to the reoccurring costs remain significantly lower than the reoccurring production cost of the expendable launch system.
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22. References Launch Philosophy and Propulsion References NASA History Office, 2006. Apollo by Numbers. [Online] Available at: http://history.nasa.gov/SP-4029/Apollo_18-19_Ground_Ignition_Weights.htm [Accessed 7 November 2012]. Belfiore, M., 2012. Stratolaunch. Popular Mechanics, 189(4), p. 62. Coopersmith, J., 2011. The cost of reaching orbit: Ground-based launch systems. Space Policy, May, 27(2), pp. 77-80. Day, D. A., 2011. Gazing Back Through the Crystal Ball. The Space Review, 28(4), pp. 50-52. ESA, 2011. Skylon Assessment Report, Noordwijk, The Netherlands : ESA. Fortescue, P., Swinerd, G. & Stark, J., 2003. SPACECRAFT SYSTEMS DESIGN. In: Chichester: Wiley, pp. 226-231. Fortescue, P. W. S. G. S. J., 2011. Spacecraft Systems Engineering 4th Edition. Chichester: John Wiley & Sons, Ltd. Gong, S., 2011. Development of low frequency pulse detonation engine with high-G mitigating system for aero-spacecraft propulsion. Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 2011, 225(G), pp. 645-656. Hall, J. N., 2011. OPTIMIZED DUAL EXPANDER AEROSPIKE ROCKET, Wright-Patterson Air Force Base, Ohio: AIR FORCE INSTITUTE OF TECHNOLOGY. Heiser, W. H. & Pratt, D. T., 1994. Hypersonic Airbreathing Propulsion. s.l.:American Institute of Aeronautics and Astronautics. Huzel, D. K. & Huang, D. H., 1992. Modern Engineering for Design of Liquid-Propellant Rocket Engines. 3rd ed. Washington: AIAA . Klein, M., 2004. Nuclear Thermal Rocket — An Established Space Propulsion Technology. Albuquerque, New Mexico (USA), AIP. Koelle, D., 2002. Development cost of Reusable Launch Vehicles. Acta Astronautica, 51(1-9), pp. 23-31. Li, J.-L.et al., 2011. Performance enhancement of a pulse detonation rocket engine. Xi’an, Proceedings of the Combustion Institute. Malik, T., 2012. Reusable 'Grasshopper' Rocket Concept Makes 1st Test Flight. [Online] Available at: http://www.space.com/17868-spacex-grasshopper-reusable-rocket-test.html [Accessed 10 October 2012]. Manski, D., Hagemann, G. & Sahick, H. D., 1997. Optimization of Dual-Expander Rocket Engines in SSTO Vehicles. Pergamon, 40(2-8), pp. 151-163. Moscow Top News, n.d. Ekranoplan. [Online] Available at: http://www.moscowtopnews.com/?area=articleCommentationController&action=add&article=88 6 [Accessed 05 11 2012]. Musk, E., 2011. National Press Club Luncheon. s.l., s.n. Musk, E., 2012. SpaceX and the future of space exploration. London, RAeS. N.A.Davydenko,R.G.A.G.,2007.Hybridrocketengines:Thebenefitsandprospects. Aerospace Science and Technology, Volume 11, pp. 55-60. NASA, 2000. Linear Aerospike Engine - Fact sheet number: FS-2000-09-174-MSFC. [Online] Available at: http://www.nasa.gov/centers/marshall/news/background/facts/aerospike.html [Accessed 31 October 2012].
Section 22 - References Page i Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Nobuyuki Tomita, A. V. N. V. V. S. Y. O., 1999. Performance and technological feasibility of rocket powered HTHL-SSTO with take-off assist (aerospace plane/ekranoplane). Acta Astronautica, 45(10), pp. 629-637. Norris, G., 2005. Superjumbo of the 21st Century. s.l.:Zenith Press. Norris, G., 2011. Positive Reaction. Aviation Week & Space Technology, 173(15), p. 74. Powell, J., Maise, G. & Paniagua, J., 2004. StarTram: Ultra Low Cost Launch. Albuquerque, New Mexico, s.n. Pratt & Whitney Rocketdyne, 1980. Rocket Engine Cost Model. s.l.:s.n. Reaction Engines ltd, 2012. Skylon Cutaway. [Online] Available at: http://www.reactionengines.co.uk/image_library.html [Accessed 21 November 2012]. Rich, B. R. & Janos, L., 1998. Skunk Works. LA: Little, Brown Book Group Limited. Robinson, T., 2012. Making operational improvements sexy. AEROSPACE, 14 September, pp. 30-32. Rogers, A., 1998. Modern Spacecraft Design. 3rd ed. London : McGraw Hill. S. C. Krausse, N., 1969. Apollo/ Saturn V Post Flight Trajectory AS-506, Cape Canavral: NASA. Schnitzler, B. G., Borowski, S. K. & Fittje, J. E., 2009. 25,000-lbf Thrust Engine Options Based on the. Denver, Colorado, s.n. Schweber, B., 2002. Electronics poised to replace. How it Works, 11 April, pp. 29-30. Scott Forde, M. B. ,. T. N., 2006. Thrust augmentation nozzle (TAN) concept for rocket engine booster applications. Acta Astronautica, 17-21 September, pp. 271-277. SpaceX, n.d. Falcon Heavy. [Online] Available at: http://www.spacex.com/falcon_heavy.php [Accessed 05 November 2012]. Virgin Galactic, 2012. Virgin Galactic Launcherone Preformance and Specification. [Online] Available at: http://www.virgingalactic.com/launcherOne/performance-and-specification/ [Accessed 12 10 2012]. W.Beny & H.Grallert, 1996. PERFORMANCE AND TECHNICAL FEASIBILITY COMPARISON OF REUSABLE LAUNCH SYSTEMS: A Synthesis of the ESA Winged Launcher Studies. Pergamon, 38(4-8), pp. 333-347. Yang, V. & Anderson, W., 1995. Liquid Rocket Engine Combustion Instability. Volume 169 ed. s.l.:Progress in Astronautics and Aeronautics.
Structures and Materials References A.M Howatson, P. L. J. T., 1991. Aluminium alloys, titanium and titanium alloys. In: 2. edition, ed. Engineering Tables and Data. s.l.:Kluwer Academic Publishers, p. 41. ASM Aerospace Specification Materials Inc, n.d. Aerospace metals. [Online] Available at: http://www.aerospacemetals.com/ [Accessed 15 10 2012]. Beckwith, D. S. W., 2009. Sanwich Core Materials and Technology - Part II. SAMPE Journal, 45(4), p. 52. Carbon Composites inc, 2005. Carbon/Carbon Composites. [Online] Available at: http://www.carboncompositesinc.com/carbon_composites.asp [Accessed 12 October 2012]. CMT, n.d. Metal Matrix Composites. [Online] Available at: http://www.cmt-ltd.com/html/mat_1.htm [Accessed 09 December 2012].
Section 22 - References Page ii Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
HexWeb, 2000. Hoenycomb Sandwich Design Technology. [Online] Available at: www.hexcell.com [Accessed 03 December 2012]. HexWeb, 2000. Honeycomb Sandwich Design Technology. [Online] Available at: www.hexcel.com [Accessed 03 December 2012]. Larson, W. J., 1999. Component Selection and Sizing. In: J. R. Wertz, ed. Space Mission Analysis and Design. El Segundo, Dordrecht: Microcosm Press, Kluwer Academic Publishers, pp. 708-10. Mortimer, C. E., 1975. Chemistry: A Conceptual Approach. In: New York: D. Van Nostrad Company. Odenward, D. S., n.d. What material is used to make spacecraft?. [Online] Available at: http://www.astronomycafe.net/qadir/q2112.html [Accessed 12 October 2012]. Panel on Small Spacecraft Technology, N. R. C., 1994. Materials. In: Technology for Small Spacecraft. Washington: National Academy Press, p. 46. Panel on Small Spacecraft Technology, N. R. C., 1994. Technology for Small Spacecarft. Washington: National Academy Press. Performance Composites Ltd, n.d. Mechanical Properties of Carbon Fibre Composite Materials, Fibre / Epoxy resin (120°C Cure). [Online] Available at: http://www.performance-composites.com/carbonfibre/mechanicalproperties_2.asp [Accessed 20 10 2012]. Rawal, u., 2001. Metal-Matrix Composites for Space Applications, New York: Springer Sciences & Business Media. Sari Katz, E. G., 2012. Encycolpedia of Composites. s.l.:John Wiley and Sons. Scanlan, J., n.d. University of Southampton. [Online] Available at: www.southampton.ac.uk/~jps7/Aircraft%20Design%20Resources/manufacturing/aluminum- lithium.doc [Accessed 03 12 2012]. Warren, D., 2010. Carbon Fiber Cost Breakdown. Low Cost Carbon Fiber Overview, Oak Ridge National Laboratory. Wiley J. Larson, J. R. W., 2004. Structures and Mechanisms. In: Space Mission Analysis and Design. El Segundo: Microcosm Press, Kluwer Academic Publishers, p. 479. Wiley J.Larson, J. R., 2004. Structures and Mechanisms. In: Space Mission Analysis and Design. El Segundo: Microcosm Press, Kluwer Academic Publishers, p. 478.
Orbit Operation and Electronics References Ball Aerospace, 2012. Ball Aerospace Delivers Orion Phased Array Antenna EDUs. [Online] Available at: http://www.ballaerospace.com/page.jsp?page=30&id=447 [Accessed 30 11 2012]. Bate, R. R., Mueller, D. D. & White, J. E., 1971. Basic Orbital Manuevers In-Plane Orbit Changes. In: Fundamentals of Astrodynamics. New York: Dover Publications Inc, pp. 163-166. Chetty, P., 1991. Telemetery, Trackin,Command and Communication Systems. In: Satellite Technology. USA: McGraw-Hill Inc, p. 248. Dumoulin, J., 2000. Orbiter Communications. [Online] Available at: http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts-ovcomm.html [Accessed 3 11 2012].
Section 22 - References Page iii Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
ESA, 2012. CAN - Controller Area Network Bus. [Online] Available at: http://www.esa.int/TEC/OBCDH/SEM8KZEURTG_0.html [Accessed 21 11 2012]. ESA, 2012. Mil-STD-1553, The history and basics of the Milbus 1553b. [Online] Available at: http://www.esa.int/TEC/OBCDH/SEMWJZEURTG_0.html [Accessed 21 11 2012]. ESA, 2012. Onboard Computer and Data Handling. [Online] Available at: http://www.esa.int/TEC/OBCDH/SEM2ZYEURTG_0.html [Accessed 20 11 2012]. European Space Agency, 2012. I2C. [Online] Available at: http://www.esa.int/TEC/OBCDH/SEM8LZEURTG_0.html [Accessed 16 11 2012]. European Space Agency, 2012. SpaceWire. [Online] Available at: http://www.esa.int/TEC/OBCDH/SEMKKZEURTG_0.html [Accessed 24 11 2012]. Fortescue, P., Stark, J. & Swinerd, G., 2003. Antenna Types. In: P. Fortescue, J. Stark & G. Swinerd, eds. Spacecraft System Engineering. Padstow: John Wiley and Sons Ltd, p. 436. Fortescue, P., Stark, J. & Swinerd, G., 2003. Therma Contol, Passive Control. In: P. Fortescue, J. Stark & G. Swinerd, eds. Spaceccraft Systems Engineering. Padstow, Cornall: John Wiley & Sons Ltd, pp. 375-376. Fortescue, P., Stark, J. & Swinered, G., 2003. Types of reference sensor. In: P. Fortescue, J. Stark & G. Swinered, eds. Spacecraft Systems Engineering. Padstow: John Wiley & Sons Ltd, pp. 310- 318. P-N Designs, 2011. Phased Array Antennas. [Online] Available at: http://www.microwaves101.com/encyclopedia/phasedarrays.cfm [Accessed 2012]. Sellers, J. J., Astore, W. J., Giffen, R. B. & Larson, W. J., 2005. Attitude control. In: D. H. Kirikpatrick, ed. Understanding Space. United States of America: McGraw-Hill Companies, Inc, pp. 426-431. Sellers, J. J., Astore, W. J., Giffen, R. B. & Larson, W. J., 2005. Living and Working in Space. In: D. H. Kirkpatrick, ed. Understand Space. United States of America: McGraw-Hill Companies, Inc, p. 95. Sellers, J. J., Astore, W. J., Giffen, R. B. & Larson, W. J., 2005. The Space Environment and Spacecraft. In: D. H. Kirkpaterick, ed. Understanding Space. United States of America: McGraw- Hill Companies,INC, pp. 86,87. Vis, P. J., 2012. PowerPC 750 / RAD750. [Online] Available at: http://www.petervis.com/Vintage%20Chips/PowerPC%20750/PowerPC%20750.html [Accessed 1 12 2012]. Wertz, J. R. & Larson, W. J., 1999. Earth Satelite Parameters. In: J. R. Wertz & W. J. Larson, eds. Space Mission Analysis and Design 3. Hawthorne and New York: Microcosm Press and Springers, p. 985. Wertz, J. R. & Larson, W. J., 2010. Space Computer Systems Computer System Specification. In: Space Analysis and Deisgn. Hawthorne: Microcosm Press, pp. 653-654. Wiesel, W. E., 2010. The Hohmann Transfer. In: Spaceflight Dynamics 3rd edition. Yalonda: Aphelion Press, pp. 79-83.
Section 22 - References Page iv Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Re-Entry References Anon., n.d. Centennial of Flight. [Online] Available at: http://www.centennialofflight.gov/essay/Theories_of_Flight/atmosphere/TH1G1.htm J. Barrie Moss, G. E. D., 2003. Re-entry into Earths atmosphere. In: Spacecraft Systems Engineering. s.l.:Wiley, pp. 233-239. J. E. Pavlosky, L. S. L., 1974. Apollo Experience Report - Thermal Protection Subsystem, s.l.: s.n. J.E. Johnson, M. L. R. P. S., 2012. Multiobjective Optomization of Earth-Entry Vehicle Heat Shields. Journal of Spacecraft and Rockets, 49(1), pp. 38-50. Jr., V. L. V., 1959. Water Landing Impact Accelerations for Three Models of Reentry Capsules, Washington: NASA. K. Bledsoe, M. E. A. M. R. O., 2009. Overview of the Crew Exploration Vehicle Parachute, s.l.: NASA. NASA, 2010. Mars Fact Sheet. [Online] Available at: http://nssdc.gsfc.nasa.gov/planetary/factsheet/marsfact.html [Accessed 11 2012]. NASA, n.d. A Canopy of Confidence: Orion's Parachutes. [Online] Available at: http://www.nasa.gov/exploration/systems/mpcv/canopy_of_confidence.html [Accessed 3 12 2012]. NASA, n.d. Landing 101. [Online] Available at: http://www.nasa.gov/mission_pages/shuttle/launch/landing101.html NASA, n.d. Mars science laboratory: spacecraft. [Online] Available at: http://mars.jpl.nasa.gov/msl/mission/spacecraft/ NASA, n.d. Orbiter Thermal protection system. [Online] Available at: http://www-pao.ksc.nasa.gov/kscpao/nasafact/pdf/tps.pdf NASA, n.d. Orbiter Thermal Protection System Diagram. s.l.:s.n. NASA, n.d. Space Transportation System. [Online] Available at: http://science.ksc.nasa.gov/shuttle/technology/sts- newsref/sts_overview.html#sts_overview R. Olmstead, A. M. K. B. M. E., 2009. Overview of the Crew Exploration Vehicle Parachute Assembly System (CPAS) Generation I Drogue and Pilot Development Test Results, s.l.: NASA. Rohrschneider, R. R., 2002. Develepment of a mass estimating relationship database for launch vehicle conceptual development, s.l.: s.n. Sellers, J. J., 2005. Understanding Space: an Introduction to Astronautics. s.l.:The Mcgraw-hill companies inc. U. Trabandt, M. S., 2010. Deployable heat shield and deceleration structure for spacecraft. US, Patent No. US 7,837,154 B2. Wehrly, D. J., 1987. Low altitude, high speed personal parachuting: medical and psychological issues, s.l.: USAARL.
Payloads and Markets References Anon., n.d. google. [Online] Available at: https://docs.google.com/viewer?a=v&q=cache:rddf04GYmBEJ:snowbears.org/asci512/512casest udies_fall06/Life%2520Support%2520For%2520Deep%2520Space%2520Travel%2520(octavian ).ppt+open+loop+ECLSS&hl=en&gl=uk&pid=bl&srcid=ADGEESihzYk2KGdBXYjlLg_w2aIp RsEZ39Lpa7jY [Accessed 21 December 2012]. Belfiore, M., 2012. Stratolaunch. Popular Mechanics, 189(4), p. 62. cameochemicals, n.d. cameochemicals.noaa.gov. [Online] Available at: http://cameochemicals.noaa.gov/chemical/1529 [Accessed 12 December 2012].
Section 22 - References Page v Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Carrasquillo, R. L. & Anderson, M., 2012. Environmental Control and Life Support (ECLS) Hardware Commonality for Exploration Vehicles, AL: NASA. Corporation, Futron; Webber, D;, 2003. Ascent Study Final Report, s.l.: NASA. Fortescue, P. W. S. G. S. J., 2011. Spacecraft Systems Engineering 4th Edition. Chichester: John Wiley & Sons, Ltd. GAO, U. G. A. O., 2011. GAO. [Online] Available at: http://www.gao.gov/assets/590/587156.txt [Accessed 20 December 2012]. Isakowitz, S. J., 1999. International refence guide to space launch systems. 3rd ed. Virginia: Reston. Janos, B. R. R. a. L., 1998. Skunk Works. LA: Little, Brown Book Group Limited. Keith, L., 1998. NASA Quest. [Online] Available at: http://quest.arc.nasa.gov/people/journals/space/keith/05-25-98.html [Accessed 22 December 2012]. Larson, W. J. & Wertz, J. R., 1999. Space Mission Analysis and Design. 3rd ed. CA: Microcosm Press. Malik, T., 2012. Reusable 'Grasshopper' Rocket Concept Makes 1st Test Flight. [Online] Available at: http://www.space.com/17868-spacex-grasshopper-reusable-rocket-test.html [Accessed 10 October 2012]. Malik, T., 2012. space.com. [Online] Available at: http://www.space.com/11125-nasa-russia-soyuz-deal-spaceflights.html [Accessed 19 December 2012]. Metcalf, J., Peterson, L., Carrasquillo, R. & Bagdigian, R., n.d. nasa.gov. [Online] Available at: http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120012436_2012011692.pdf [Accessed 26 November 2012]. Moore, T. A. & Yamada, N., 1998. Halon Options Technical Working Conference. [Online] Available at: http://www.nist.gov/el/fire_research/upload/R0000296.pdf [Accessed 22 December 2012]. Norris, G., 2005. Superjumbo of the 21st Century. s.l.:Zenith Press. Pettit, D., 2011. airspacemag.com. [Online] Available at: http://blogs.airspacemag.com/pettit/page/5/ [Accessed 22 december 2012]. Purser, P. E., Faget, M. A. & Smith, N. F., 1964. "Manned Spacecraft: Engineeirng Design and Operartion.". New York: Fairchild Publications Inc.. Robinson, T., 2012. Making operational improvements sexy. AEROSPACE, 14 September, pp. 30-32. Schweber, B., 2002. Electronics poised to replace. How it Works, 11 April, pp. 29-30. sprectrex-inc, n.d. sprectrex-inc.com. [Online] Available at: http://spectrex-inc.com/files/sharpeye/presentations/firedetectiontypes_may08.pdf [Accessed 22 Decemeber 2012]. Tatars, J. D. & Perry, J. L., 2004. Spacecraft Cabin Atmosphere Major Constituent Monitoring Using, Off-the-Shelf Techniques, Alabama: NASA. Virgin Galactic, 2012. Virgin Galactic Launcherone Preformance and Specification. [Online] Available at: http://www.virgingalactic.com/launcherOne/performance-and-specification/ [Accessed 12 10 2012]. Wieland, P. O., 1994. Designing for Human Presence in Space: An Introduction to Environmental Control and Life Support Systems, Alabama: NASA.
Operations and Finance References Andrews, D., 2011. The Space Review. [Online] Available at: http://www.thespacereview.com/article/1881/1 [Accessed 12 November 2012]. Baroth, E. et al., 2011. IVHM (Intergrated Vehicle Health Management) Techniques For Future Space Vehicles, s.l.: AIAA.
Section 22 - References Page vi Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Charania, A., Wallace, J., Olds, J. R. & Zapata, E., 2004. Design For Operations of Future Reusable Launch Systems, Vancouver: International Astronautical Congress. Fox, J., 2001. Informed maintenance for next generation space transportation systems, s.l.: NASA. Gstattenbauer, G. J., 2006. Cost Comparison Of Expendable, Hybrid and Reusable Launch Vehicle, Ohio: Air Force Institute of Technology. Jambor, D. B., 1997. System health management/Vehicle health management for future manned space systems, s.l.: NASA. Jiang, H. & Hansman, J., 2006. An Analysis of Profit Cycles in the Airline Industry, Kansas: AIAA. Keith, E. L., n.d. Launch Vehicle System - Reusable. Maryland, Applied Technology Institute. Keith, E. L., n.d. Reducing Space Launch Cost. Maryland, Applied Technology Institude. Koelle, D. E., 2002. Handbook of Cost Engineering for Space Transportation Systems with Transcost 7.0, Germany: TransCostSystems. Loehr, R., 2012. Mishaal Aerospace. [Online] Available at: http://www.mishaalaerospace.com/news.php?id=125 [Accessed 04 October 2012]. NASA, 2012. NASA's Plan To Modify The Ares I Mobile Launcher In Support Of The Space Lauch System, s.l.: NASA. Patterson-Hine, A., Deb, S., Kulkarni, D. & Wang, Y., 2001. Automated System Checkout To Support Predictive Maintenance for Reusable Launch Vehicles, s.l.: NASA. Pearson, J. et al., 2000. Low-Cost Launch System For The Dual-Launch Concept, Brazil: Air Force Research Laboratory. Rohrschneider, 2002. Development of mass estimating relationship database for conceptual design, Georgia: Georgia Institute of Technology,. Rothschild, W. J. & Talay, T. A., n.d. A Near-Term, High-Confidence Heavy Lift Launch Vehicle, Houston: NASA. Scanlan, J., 2004. Method of Calculating Direct Operating Cost, Southampton: Southampton University. Sethi, C., 2012. 2012: SpaceX's Odyssey, s.l.: ASME. Snead, J. M., n.d. Cost Estimates of Near-Term, Fully-Reusable Space Access System, Ohio: American Institude of Aeronautics and Astronautics. SpaceX, n.d. US Air Force Evolved Expendable Launch Vehicle (EELV). [Online] Available at: http://www.spacex.com/EELVBenefits.pdf [Accessed 15 November 2012]. Staton, E. J., 2002. Developing the Operational Requirements for the Next Generation Launch Vehicle and SpacePort, Georgia: Georgia Institude of Technology. Wertz, J. R., 2004. Response Launch Vehicle Cost Model, Los Angeles : Microsm.
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23. Appendices A. Launch Philosophy Appendix A.1. Ideal (Tsiolkovsky) Rocket Equation and Staging The Ideal Rocket Equation describes the velocity gain ( ) based on the change in system mass and ideal nozzle velocity (Fortescue, 2011):
Where: is the change in velocity as a result of the manoeuvre. is the ideal nozzle velocity [m/s] is the initial mass [kg] is the mass at the end of the rocket burn
It can be seen that it is beneficial to maximise the exhaust velocity and minimise the empty mass.
This can be expanded to include multiple stages. For one Stage:
Where: is the specific impulse of the engine is acceleration due to gravity.
For n stages: ∑ Accounting for the losses due to gravity and Aerodynamic Drag
Typically and (Fortescue, 2011)
From an analysis of these equations an assessment of the optimum number of stages can be generated (Fortescue, et al., 2003):
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Figure 31: Assessment of the Optimum Number of Stages (Fortescue, et al., 2003) From Figure 31 it can be seen that as the payload fraction P increase the benefit in increasing the number of stages decreases. For a typical LEO payload three stages is the optimum solution.
A.2. Assessment of the Viability of Captive Launch Systems A.2.1. Subsonic Air Launch Systems Using Stratolaunch (see Figure 31) as a reference for an Air Launch system the payload is extrapolated into the parameters for the lower payload limit for of 25,000kg in Table 43 (Belfiore, 2012).
Figure 32: The Stratolaunch Concept of 2 modified 747s A comparable traditional launch rocket take-off mass is calculated based on the following mass fractions:
System Component Payload Structure Fuel Mass Fraction 4% 10% 86%
Table 43: Extrapolation of System mass from Stratolaunch Characteristic Stratolaunch Concept Lower Limit Payload 6,100kg 25,000kg Scaling Factor - 4.10
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Span 117m 480m Carrier Craft Take-off Weight 324,311kg 1,329,675kg Payload Rocket Mass 220,000kg 902,000kg Total Take-off Mass 544,311kg 2,231,675kg Equivalent Rocket Launch Mass 152,000kg 625,000kg
It is immediately apparent that the Stratolaunch is unlikely to be economically feasible unless significant savings can be made elsewhere during launch and preparation. The payload fraction is extremely low and there appears to be limited benefit in terms of reducing the rocket mass. Scaling the concept up to the lower payload limit of 25,000kg makes the aircraft unfeasibly large and would require a take-off mass approximately 3.5 times that of the A380 MTOW (Norris, 2005). The Stratolaunch carrier aircraft is derived from 2 747-400s (Belfiore, 2012) but for an orbital payload of 25,000kg a wholly new aircraft type would need to be designed, manufactured and tested. This would require a significant financial investment and is unlikely to that an investor or customer could be found this new concept when current rocket technology is superior. A.2.2. Supersonic Air Launch Systems A supersonic launch system would be nearly impossible to build for the specified payload range. The exotic materials required to build the carrier aircraft, the high fuel fraction required for supersonic flight and the challenges associated with vehicle separation at supersonic speeds mean that this concept is unsuitable for this application. The exotic materials and high engine thrust requirement means that this is the least environmentally friendly system of these three. A.2.3. Sea Level Captive Launch Systems Ground Effect Vehicle Theory At low speeds drag on an aircraft is primarily due to induced drag. Induced drag is causes be downwash effects from the vortices at the wing tips. At very low altitudes ground effect causes a large reduction in induced drag and so less thrust is required to maintain flight (Rogers, 1998). It is considered a very efficient way of maintaining high subsonic speeds at low altitudes and hence has been considered as a suitable launch platform in reference (Nobuyuki Tomita, 1999). Large Ground Effect vehicles have been operated by the Russian Navy which could attain high speeds with a small wing and relatively low thrust requirements in the cruise.
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Figure 33: Russian Ekranoplan (Moscow Top News, n.d.) Feasibility of a Sea Level Captive Launch System Like the air launch concepts above the Ground Effect Vehicle (GEV) would be very large and would face significant operational difficulties. The sea state would need to be very calm to allow a successful launch and the recovery of an orbiter launched by this method and this presents significant technical challenges. A suitable launch vehicle would weigh 1000 tonnes and like air launch systems, the financial viability of building a large special to type craft is low. A.2.4. Conclusions on the Viability of Captive Launch Systems Technical Viability: The high altitude air launch systems are almost complete unfeasible for the payload range specified for this project, the launch aircraft is almost unbuildable. The Sea Level system has a higher level of feasibility due to the use of ground effect which allows the structural mass fraction to be higher which in turn makes the craft easier to design and manufacture.
Financial Feasibility: Due to the extremely high initial investment required for the development of two vehicles and then need to create an economical reusable payload delivery system the financial viability of these concepts are low.
Development Program Risk: The development of two systems creates an additional critical path for the development program which in turn increases project risk considerably compared to the equivalent SSTO or Reusable Rocket program.
Environmental Impact: Reusable launch systems have inherent environmental benefit over the equivalent expendable system. Using a captive launch system does little to affect these benefits as the launch profile has little impact on the environmental impact however because of the more efficient cruise phase the sea level launch system is likely to use the least fuel off all three systems and the supersonic launch system would use considerable more fuel.
Reusability: The air launch concepts all require a rocket system to attain orbit. The proposed solution appears to have a low Technical Viability and so there would not be sufficient mass to incorporate reusable launch components.
Table 44: Numerical Ranking of the Captive Launch System Area Air Launch Air Launch Captive Launch (0 – 5) where 5 is better Subsonic Supersonic at Sea Level Technical Viability 2 1 3 Financial Feasibility 1 1 1 Development Program Risk 1 1 1
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Environmental Impact 1 0.5 2 Reusability 2 2 2 Total (of a possible 25) 7 5.5 9
A.3. Assessment of the Viability of EMOLS and Nuclear Launch A.3.1. Orbital Velocity EMOLS Constructing an EMOLS to attain orbital velocity is very expensive. StarTram propose to use a High ‘g’ cargo configuration to minimise launch tube length and constructioncosts associated with making the system Human Payload Compatible5. Currently estimate that it will cost $36.38 Billion to construct and could yield an operational cost of $37.73/ kg (see below) after inflation which is extremely competitive (Powell, et al., 2004).
However this is based on 16,000 Launches/Year for 10 Years which based on the market research in the inception report seems unlikely. A more likely figure of 200 launches per year and an increased life of 15 years yields the following:
Which when adjusted for inflation for 2012 yields a cost of $514.02 which is very competitive compared to a current expendable rocket system payload cost of $2415/kg (SpaceX, n.d.) this also includes the development costs of the system.
A.3.2. Summary of EMOLS Technical Viability: The Technical Viability of the construction of the EMOLS within the next 20 years is low. Although the company behind it present convincing data their financial analysis lacks consideration of the technical challenges associated with such a large structure and the practicality of the extremely large magnetic fields needed to suspend it.
Financial Feasibility: As above. The proposed financial plan makes considerable assumptions about the frequency of launch that the market could generate.
Development Program Risk: The technical risk is extremely high and it is likely that such a large monolithic structure would have significant risk and a difficult critical path.
Environmental Impact: To obtain the rare earth metals required to construct the super conductors a significant increase in mining activities would be required. The EMOLS system would likely be built in a remote location but would require a significant onsite construction facility. Both of these present significant environmental challenges.
5 Human Payloads require launch forces of < 3g. This would make the StarTram system unfeasibly long.
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Reusability: An EMOLS system would significantly reduce the size and weight of the launch payload but it would be an expendable launch system.
Table 45: Numerical Summary of the EMOLS System Area EMOLS (0 – 5) where 5 is best Technical Viability 2 Financial Feasibility 0.5 Development Program Risk 0 Environmental Impact 1 Reusability 3 Total (of a possible 25) 6.5 A.4. Assessment of the Vertical Rocket Launch Systems A.4.1. Expendable Rocket Benchmark Case To create a baseline for a large two stage rocket data from the Saturn V was used:
Table 46: Saturn V Performance Information (S. C. Krausse, 1969) and ( NASA History Office, 2006) Parameter Value Launch Mass 2.886E6 kg Launch Thrust 34.5E6 N Launch Thrust to Weight Ratio 1.26 First Stage - Engines 5 x F-1 Engines Thrust 34.5E6 N at sea level Isp 304s Mass 2.268E6 kg Empty Mass 135E3 kg Burn Time 168s Second Stage 5 x J-2 Engines Thrust 5.166E6 N Isp 421s in vacuum Mass 490E3 kg Empty Mass 39E3 kg Burn Time 390s
A.4.2. Optimum Rocket Launch Stage Recovery Assessment of the Payload Mass Fraction Reduction of Recovering the 1st Stage Saturn V 1st Stage Separation occurs at a height of 67km and speed of 2000m/s at a flight angle of 22 degrees 93.6km downrange from the launch site. From this altitude there are two return options, water recovery where the stage is parachuted into water and is recovered and tail landing where the rocket uses its own thrust to conduct a controlled landing. Landing on water presents significant operation issues. The recovery ships for the stage would be very large and the ingress of sea water would mean the system would have to be overhauled before the next launch. The solution most suited to a reusable system is to recover the stage under powered flight. Once it landed it could be refuelled and reused in a much shorter time.
At 2000m/s a reverse burn would be required to decelerate the craft and then some thrust would be required to minimise atmospheric flight velocity and aerodynamic heating. Additionally small landing system would be required as well as flight systems capable of manoeuvring the craft.
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For return the earth the was calculated as:
⁄ ⁄ ⁄
Using scaled mass data from the Saturn V for the 25,000kg payload and a representative Isp value of 380s:
( ( ) )
( ( ))
An iterative process was then used to calculate the optimum fuel to mass value. The structural mass of the first stage was increased by 10% to account for the landing systems and the total increase in launch mass was 11.1%. Assessment of the Payload Mass Fraction Reduction of Recovering the 2nd Stage Based on the analysis conducted in the re-entry systems section it was found that a re-entry heat shield is approximately 10% of the mass being re-entered into the atmosphere. A.4.3. Mass Predictions of a Spaceplane Table 47: Example Mass Fractions for a SSTO/TSTO Spaceplane (W.Beny & H.Grallert, 1996) Mass Predicted Mass Skylon 25,000 kg 40,000 kg Fractions Payload Payload for Configuration Configuration Single Stage Payload 4.4% Payload Mass [kg] 12,000 25,000 40,000 Fraction Structural 15.6% Structural Mass [kg] 42,545 88,635 141,817 Fraction Fuel 80% Fuel Mass [kg] 218,182 454,546 727,273 Fraction Total Mass [kg] ~283,000 568,180 943,000 Example Length [m] 82 171 273 Example Span [m] 25 52.1 83 A.4.4. Mass Predictions for the Rocket Family Case 1 - 25,000kg Payload, 2 Stage Variant A
Mass Fraction of Fraction of Launch Component [kg] Stage Mass First Stage Structure 4.57E+04 10.42% 7.46% First Stage Fuel 3.41E+05 77.79% 55.70% First Stage Recover Fuel and Systems 51700.14 11.79% 8.44% First Stage Total 4.39E+05 - Second Stage Structure 1.09E+04 7.30% 1.77% Second Stage Fuel 1.36E+05 91.24% 22.19% Second Stage Heat Shield 2173 1.46% 0.35%
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Second Stage Total 1.49E+05 - - 25,000kg Payload to 200km LEO 2.50E+04 4.08% Total 6.12E+05
Variant A Predicted Mass Fractions 100%
90%
25,000kg Payload to 80% 200km LEO Second Stage Heat Shield 70%
Second Stage Fuel 60%
Second Stage Structure 50%
First Stage Recover Fuel 40% and Systems
30% First Stage Fuel
20% First Stage Structure
10%
0%
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Case 2 - 40,000kg Payload, 2 Stage Variant B
Fraction of Fraction of Launch Component Mass [kg] Stage Mass First Stage Structure 4.57E+04 8.49% 7.46% First Stage Fuel 4.57E+05 84.88% 74.64% First Stage Recover Fuel and Systems 3.57E+04 6.64% 5.84% First Stage Total 5.39E+05 - - Second Stage Structure 2.59E+04 8.93% 4.23% Second Stage Fuel 2.59E+05 89.29% 42.29% Second Stage Heat Shield 5.18E+03 1.79% 8.46% Second Stage Total 2.90E+05 - - Third Stage Structure 7.81E+03 8.93% 1.28% Third Stage Fuel 7.81E+04 89.29% 12.75% Third Stage Heat Shield 1.56E+03 1.79% 2.55% Third Stage Total 8.75E+04 - - 40,000kg Payload to 200km LEO 4.00E+04 6.53% Total 9.56E+05
Variant B Predicted Mass Fractions 100% 40,000kg Payload to 90% 200km LEO Third Stage Heat Shield 80%
Third Stage Fuel 70%
Third Stage Structure 60%
Second Stage Heat Shield 50%
40% Second Stage Fuel
30% Second Stage Structure
20% First Stage Recover Fuel and Systems 10% First Stage Fuel
0%
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B. Propulsion Appendix B.1. Concept Analysis B.1.1. TAN Nozzles The proposed thrust augmentation nozzle uses combustion in the nozzle to reduce overexpansion losses. See Figure 34.
Figure 34: Schematic of TAN Nozzle Operation (Scott Forde, 2006) Potential improvements in Thrust to Weight ratio and the minimal impact on ISP, which in this case represents fuel burn rate, is shown in Figure 35.
Figure 35: The potential improvements in T/W and minimal impact on ISP of TAN Systems (Scott Forde, 2006) B.1.2. Aerospike Engines An example of the X-33 Aerospike Engine compared to a traditional bell-engine nozzle is shown in Figure 36. The combustion chambers for the Aerospike engine are arranged along the top edges and they project exhaust onto the wedge assembly.
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Figure 36: X-33 Aerospike Engine (right) Compared to a traditional bell nozzle (left) (NASA, 2000) B.2. Summary of Current Generation Engines For initial calculation a specific impulse was required. Table 48 is a summary of theoretical values for a variety of fuels combusted with liquid oxygen:
Table 48: Summary of LO2 Fuels (Huzel & Huang, 1992) Oxidiser Fuel Isp at Sea Level Isp in a Vacuum Comments LOX H2-Be 49/51 459 540 Includes Beryllium LOX H2 389 455 Common LOX B2H6 342 409
LOX CH4/H2 92.6/7.4 319 379 Used in many first 300 358 LOX RP-1 Stages
B.3. Concept Elimination B.3.1. Solid Rocket Technical Viability: A solid rocket booster is a technically viable solution however although the system cheap to construct the Isp is typically only 70% of an equivalent liquid fuelled rocket making the rocket less efficient. It is also incompatible with a reusable launch stage because a solid rocket motor cannot be reignited efficiently. Solid rockets could be used as a booster system for any of the concepts.
Financial Feasibility: The low development and manufacture costs of a solid rocket are attractive to a low costs launch system, however there is less return from reusing solid rockets. They would tend to require a parachute and sea recovery and this would significantly increase operational costs.
Development Program Risk: Using a solid rocket is a low risk solution. Although the rocket would be larger than a similar liquid fuel rocket it is one of the lower design risk options.
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Impact on Other Subsystems: A solid rocket it quite compatible with most subsystem areas in this report. The difficulties in shutting down a solid rocket booster would reduce system safety levels making it less favourable to a human payload (Day, 2011).
Environmental Impact: The Environmental impact of a solid rocket is comparable to a liquid rocket.
Reusability: A solid rocket would have a low level of reusability due to the requirement to retrieve it from the sea and overhaul it.
Table 49: Numerical Summary of Solid Rockets Area Solid Rocket (0 – 5) where 5 is best Boosters Technical Viability 1.5 Financial Feasibility 1.5 Development Program Risk 3 Environmental Impact 3 Integration 2 Reusability 1 Total (of a possible 30) 12 B.4. Number of Engines Taking the common first stage, for the 40,000kg variant the launch thrust requirement based on average flight acceleration of 2.5g:
Assuming 9 engines: ⁄ Taking the thrust-weight ratio of 150:
( )
Carrying out similar calculations for the other stages: Stage Thrust Required Engines Required Number of Engines Stage 1A 15 MN 5.12 9 Stage 2A 3.41 MN 1.16 3 Stage 1B 23.4 MN 8 9 Stage 2B 2.79 MN 2.8 3 Stage 3B 2.50 MN 0.85 1
There is some miss match between the common first stages as is to be expected. It was feasible to shut down two engines during the launch to prevent too much thrust being supplied to the engines.
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B.5. Development Costs of Reusable Launch System
Figure 37: The Comparative development costs of a Reusable Rocket (Koelle, 2002)
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C. Structures and Materials Appendix C.1. Surface Area to Volume Ratios Shape Surface area to volume ratio Surface area to volume ratio for unit volume Tetrahedron √ 7.21
Cube 6
Octahedron √ 5.72
Dodecahedron √ √ 5.31 √ Sphere 4.836
C.2. Second Moments of Area Second moment of area Cross section of Beam Second moment of area (Ixx) (m4) (Iyy) (m4)
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C.3. Material Properties (A.M Howatson, 1991),(ASM Aerospace Specification Materials Inc, n.d.) Young’s Yield Ultimate Shear Fatigue Density Electrical Thermal Thermal Hardness Shear Melting modulus strength Tensile Strength Strength (g/cm3) resistivity Conductivity expansion (Vickers modulus point Material (GPa) (MPa) Strength (MPa) (MPa) (10- (Wm-1k-1) (µm.m-1.k- MPa) (GPa) (Kelvin) (MPa) 6ohm- 1) cm) Aluminium 70 15-20 40-50 - - 2.70 2.82 237 23.1 167 26 933.47 Aluminium 69 276 310 207 96.5 2.7 3.99 167 23.6 107 26 925 Al6061-T6 Aluminium 68.9 276 434 152 68.9 2.7 3.32 200 23.4 83 25.8 927 Al6063-T6 Aluminium Al5052- 70.3 386 538 165 138 2.68 4.99 138 23.8 87 25.9 922 H38 Aluminium 72.4 345 427 283 124 2.78 4.49 151 23.2 142 27 911 Al2024-T6 Aluminium 71.7 503 572 331 159 2.81 5.15 130 23.6 175 26.9 908 Al7075-T6 Aluminium Al2219- 73.1 290 414 255 103 2.84 5.7 120 22.3 130 27 916 T62 Titanium 105 275-410 344 - - 4.5 55 16.4 8.6 209 45 1938 grade 7 Titanium Grade 7 103 340 430 380 240 4.5 55 16.4 8.6 209 39 1938 annealed Titanium Ti-6Al-6V- 1900- 110.3 1350 2100 760 690 4.54 157 6.6 9 367 45 2SN 1922 Annealed Titanium 120 910 937 620 725 4.37 197 6 8.5 349 46 1813 Ti-8Al-
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1Mo-1V Titanium Ti-5Al- 110-125 827 861 520 530 4.48 160 7.8 9.4 349 48 1863 2.5Sn Titanium Ti-5Al- 1877- 113.8 790 860 520 410 4.48 178 6.7 8.6 341 44 2.5Sn 1933 Annealed Aluminium 76 210 320 190 276 2.59 2.8 88 21.4 97 28 833-923 2090-T3 Aluminium 76 520 550 320 220 2.49 2.8 88 21.4 176 28 833-923 2090-T83 Aluminium 77 210 340 200 276 2.54 2.8 95.3 21.4 102 28 873-928 8090-T3
(Performance Composites Ltd, n.d.) Ultimate Ultimate Ultimate Ultimate Coefficient Coefficient Youngs Youngs Shear Shear Density tensile tensile compressive compressive of thermal of thermal Material Modulus Modulus Modulus strength (g/cm3) strength strength strength strength expansion expansion (Gpa) 0o (Gpa) 90o (Gpa) (MPa) (MPa) 0o (MPa) 90o (MPa) 0o (MPa) 90o 0o 90o Standard Carbon 70 70 5 1.6 600 600 570 570 90 2.1 2.1 Fibre fabric High Modulus Carbon 85 85 5 1.6 350 350 150 150 35 1.1 1.1 Fibre fabric Glass 25 25 4 1.9 440 440 425 425 40 11.6 11.6 fabric Kevlar 30 30 5 1.4 480 480 190 190 50 7.4 7.4 fabric Standard 135 10 5 1.6 1500 50 1200 250 70 -0.3 28 Carbon
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Fibre Uni directional High modulus Carbon 175 8 5 1.6 1000 40 850 200 60 -0.3 25 Fibre Uni directional Glass fibres Uni 40 8 4 1.9 1000 30 600 110 40 6 35 directional Kevlar fibres Uni 75 6 2 1.4 1300 30 280 140 60 4 40 directional Boron fibres Uni 200 15 5 2 1400 90 2800 280 140 18 40 directional
C.4. Foam Core Properties Range of Shear Strength (MPa) Shear Modulus (MPa) Compressive Strength Max operating Foam density (g/cm3) (MPa) temperature (K) Polyurethane 0.021-0.4 0.15-3.1 1.55-104 0.2-0.35 408 Polystyrene 0.03-0.06 0.25-0.6 4.5-20 0.3-0.9 373 Polyvinylchloride 0.03-0.4 0.35-4.5 8.3-108 0.3-5.8 328-393 Polymethacrylimide 0.03-0.3 0.8-7.5 19-290 0.8-16 413 Polyetherimide 0.06-0.11 0.8-1.4 18-30 0.7-1.4 453-463 Styrene acrylonitrile copolymer 0.048-0.16 1.3-3.5 13.8-41.4 0.35-10.3 408 Epoxy 0.08-0.32 0.45-5.2 - 0.62-7.4 450 Phenolic 0.005-0.16 0.01-1.45 - 0.014-2.07 418-473 Carbon/graphite 0.03-0.56 0.05-3.9 - 0.2-60 ~2773
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C.5. Foam Core Characteristics Foam Cost (relative) Characteristics of foam Polyurethane Low Moderately resistant to fire and good resistance to solvents. Good capability with temperature Polystyrene Low Lowest operating temperature, fragile but cheapest with moderate mechanical properties Polyvinylchloride Great strength and stiffness, very easy to bond, great impact resistance but only a moderate capability Low with temperature Polymethacrylimide Very high performing in terms of mechanical properties, dimensionally stable at elevated temperature, High low thermal conductivity and resistant to solvents. Polyetherimide High Thermally stable with good dielectric properties, good strength Styrene acrylonitrile copolymer Moderate High stiffness, impact and fatigue strength. Environmentally friendly Epoxy Good high temperature performance, high strength and stiffness, highly compatible with laminated Moderate systems Phenolic Moderate Good properties at high temperatures, a good insulator but brittle Carbon/graphite High Best temperature properties by far, high stiffness but very high cost
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C.6. Honeycomb Properties (HexWeb, 2000)
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C.7. Advantages and Disadvantages of Honeycomb Cores (Beckwith,2009)
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D. In Orbit Operations and Electronics D.1. Graphs to show the mass that can be lifted to different orbital altitudes
The graphs show that as the altitude that the mission requires above the earth increases the amount of fuel that the mission requires increases. As the amount of fuel required increases the mass of payload that can be carried has to be reduced. The green line shows the payload that can be carried if the payload needs to return to earth. The mass of the heat shield and additional fuel required to recover the vehicle by returning to a LEO orbit and re-enteringtheearth’satmosphere requires a large amount of the spacecrafts mass, therefore reducing mass available for the payload. Therefore missions that travel to MEO and GEO orbits will not be able to return the entire craft/payload to the earth.
30000
25000 The maximum mass that can be lifted to an orbital altitude (equatorial plane) 20000 Payload mass to alltitude (allowing for fuel and 15000 250kg boost weight) kg
disposable launch boost 10000 stage minimum mass
5000 If payload requires re- Masskg in entry kg 0 200 4200 82001220016200202002420028200322003620040200 Altitude in km -5000
-10000
-15000 Figure 38 Graph to show the mass that can be lifted to an orbital altitude by the variant A rocket (25000kg capacity)
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50000 The maximum mass that can be lifted to an Payload mass to alltitude (allowing for fuel and 40000 orbital altitude (equatorial plane) 250kg boost weight) kg
30000
20000
10000Mass kg in
0 200 4200 8200 12200 16200 20200 24200 28200 32200 36200 40200 Altitude in km -10000
-20000 Figure 39 Graph to show the mass that can be lifted to an orbit altitude by the variant B rocket (40000kg capacity)
D.2. Sources of disturbances D.2.1. Aerodynamic torques Aerodynamictorquesoccurifthespacecraft’scentreofpressureisdifferentfromthecentreof mass. The aerodynamic drag that affects all orbits below 600km creates a force on the spacecraft body this causes a rotation. This effect is amplified on craft with a large surface area such as orbiters with wings or arrays of solar panels.
D.2.2. Solar Radiation Pressure Electromagnetic radiation carries momentum, when the radiation collides with the spacecraft, the momentum is transferred from the radiation to the spacecraft causing the momentum of the spacecraft to be changed, this causes the spacecraft orbit to change (this is called a perturbation) (Fortescue, et al., 2009). The force caused by collision with electromagnetic radiation is called solar radiation pressure.
D.2.3. Magnetic disturbances Magnetic disturbances are causedwhen the earth’s magnetic fieldinteracts withthe spacecraft (Sellers, et al., 2005). Spacecraft electronics create magnetic fields and also if ferris materials are used in the structure that also creates magnetic fields, when these magnetic field interact with the earth’s magnetic field there will be a torque generated between the fields, this will alter the spacecraft’sattitude and orbit.
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D.3. Addition Sensor Systems D.3.1. Sun sensors- [EE05-05] Sun sensors can be used to work out the orientation of the spacecraft with respect to the sun. A sensor works on having an optical sensor in a box with a small aperture. Depending on the angle between the sensor and the sun the light shines on a different part of the sensor. A series of sun sensors can be combined to form a digital sun sensor which will give a binary output depending on the angle to the sun. To improve reliability the sensors should be arranged in pattern that creates a Gray code output (Fortescue, et al., 2003).
D.3.2. Magnetometers- [EE05-06] Magnetometers can be used to calculate the spacecraft attitude relative tothe earth’s magnetic field. A flux gate magnetometer contains a ferromagnetic core with two coils of wire around it. A triangular waveform is applied to the primary coil. The magnetic field generated by the primary core generates a current in the second coil. However the current in the second coil is also dependant on the magnetic field of the earth, this causes it to be asymmetrical. This phase difference allows the strength of the field in that direction to be measured. If 3 flux gate magnetometers are used orthogonal to one another then the field vector of the earth can be calculated, this allows the spacecrafts attitude to be determined. This method of measuring the spacecraft’s attitude is most suited to low altitude orbits, however there are many cases in which the measurements of the earth magnetic field have been inaccurate. Inaccuracies are caused by other magnetic fields interfering with the earth magnetic field (other fields are created by electronic equipment). Therefore magnetometers are often mounted on booms away from the main body of the space craft.
D.3.3. Doppler ranging If the spacecraft is able to measure the Doppler effect of a transmission from the earth, the spacecraft should be able to calculate the distance between the spacecraft and earth station. This could be done periodically to make sure that the spacecraft is in the correct position. D.4. Redundancy and Error Checking
Hamming code can be used to check for errors in data, for every data byte stored in memory a 4 bit code is created, when the data is later read from memory this code is generated again from the data this allows the two codes to be compared, this allows any bit flips to be corrected. Triple modular redundancy (TMR) allows data to be corrected before it is used by the processor, the data is read from the memory and the data compared using TMR allowing the correct data to be found through the majority answer, this data is then written back to the memory units. This is done constantly to reduce errors in the memory. TMR corrects errors in the data when it is used
Robert Taylor D - In Orbit Operations and Electronics Page D3 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 however it does not correct the original error in the data stored in memory. This is done by a process called memory scrubbing. Memory scrubbing is a process that removes errors from the memory of the system. A section of code is repeatedly executed; this piece of code reads the memory and corrects the errors using triple redundancy, hamming coding or a combination of the two systems. Once the error has been corrected, the correct data is written back to the memory removing the error in memory. D.5. Type of Modulation
3
2
1 standard freqency 0
1 amplitude
38 75
297 112 149 186 223 260 334 371 408 445 482 519 556 593 630 667 704 phase -1
-2
-3
Figure 40 graph to show the different kinds of modulation that can be used to send data over a carrier signal The change in frequency, amplitude or phase signifies the change between a digital 1 and a digital 0. However the size of phase changes can be varied and if quadrature phase shift key was used, the changes in phase could signify the change between 00, 01, 10 or 11 allowing more data to be transmitted. D.6. OSI model
The link between the physical and software layers within a modern electronics system are often arranged using the OSI model which specifies the processes that are carried out by each layer of the system and how the layers interface. This allows pieces of software and hardware to be combined easily as they use a standard interface. An outline of the OSI model is shown in the diagram below:
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Figure 41 Diagram of the OSI layer model from (ESCOTAL.COM, 2012)
D.7. Avionics
The easiest way to control a set of avionics that might be required for flight or gliding back to earth would be to use traditional avionics components; these would require a standard avionics bus such as MIL-STD- 1553, as a series of standard modules would be able to be used to control the spacecraft in the atmosphere. However as the majority of the spacecraft control system would not be avionics, an interface between the two types of bus would be required that would allow the flow of data between the two buses. This interface could also manage the avionics bus which would make this into the secondary bus with the spacecraft bus being the primary bus. The MIL- STD-1773 storage management system would be used with MIL-STD-1553 systems. The system will need to control a variety of control surfaces including ailerons, rudder, elevators and airbrakes that allow the flight of the aircraft to be controlled. These requirements are very similar to a traditional plane. Traditional avionics includes several modules such as, an international navigation system this module is used to navigate the aircraft through the atmosphere. Avionics also requires a GPS unit, the spacecraft only requires one GPS therefore a choice will need to be made between a standard aviation module (that may not be suitable for space) or a spacecraft version that might not be accurate enough for normal flight. In the worst case a new GPS module may have to be created that is suitable for both applications.
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D.8. Bus Topologies
Star Ring Bus
Figure 42 Diagram to show the different kinds of bus topologies D.9. MIL-STD-1553B Bus
Figure 43 Diagram of how a MIL-STD-1553 bus system can be implemented (ESA, 2012). D.10. Antennas
Helix antennas [EE04-02] are suited to frequencies below 4 GHz however they are only capable of providing circular polarisation and a relatively wide beam (Fortescue, et al., 2003). They are often used with S band systems Horn antennas are ideal for communication with the earth due to its relatively small aperture, in its simplest form a wave guide is spread into a rectangular or circular aperture. Horn antennas are capable of transmitting frequencies of 4 GHz or higher (Fortescue, et al., 2003). Dish antennas [EE04-04] are a type of reflector antenna; these are normally illuminated by a feed horn. The position of the feed horn can be varied to produce a variety of different antennas with different characteristics. A front feed is most common as this is very simple to produce, however this is not ideal as the feed and the supporting structure block the path of the beam (Fortescue, et al., 2003).
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D.11. Active Temperature control Systems D.11.1. Variable conducting heat pipes- [EE09-07] Variable conducting heat pipes are a variation on a normal heat pipe in which the pressure of the vapour inside the pipe is kept at constant to maximise the heat transfer. This is done using a piston and a second gas on the other side which changes the volume that the heat transferring vapour occupies.
D.11.2. Louvers and shutters- [EE09-09] Louvers and shutters are mechanical moving planes on the outside surface of the spacecraft, changing the angle of the louver changes the emitance of the radiator beneath the louver. When the louver is closed the radiator cannot radiate energy into space, however when the louver is open the radiator is free to radiate into space. Although they provide a variable system they are very heavy and mechanically unreliable. D.12. System monitoring
Radiation monitoring could be used to monitor the total dose of radiation to the electronic subsystem, this would allow the module to be replaced before the modules receive more than the dose that they are designed to withstand. Radiation monitoring would also allow information to be collected about the conditions under which single events are most likely, this would allow subsystems to be turned off when high levels of radiation are passing through the spacecraft reducing the chance of errors. The spacecraft will need an array of temperature sensors to monitor different sections of the system. This data will be used with the temperature control system to keep the system within its optimum operating range. Flight data from other missions in the past often shows a correlation between a temperature change and failures on the spacecraft, (changes in temperature cause failures and failures can cause changes in temperature.) The outer surface of the spacecraft will need to be monitored during re-entry to make sure that the temperature of the heat shield or the surface of the spacecraft beneath the heat shield do not reach a critical temperature. The forces on the spacecraft and the effect they have on the structure could be monitored by using a variety of sensors. G- force sensors and accelerometers could be used to calculate the magnitude of the force. The space frame could have stress and strain gauges placed around the system to monitor the effect these forces have on the structure. Radar could be used around the spacecraft to detect space junk and micro meteorites that may hit the spacecraft. Fuel sensors will be required to monitor the amount of fuel in the fuel tanks. The human payload module will require atmospheric monitoring to make sure that the air quality within the spacecraft is to a sufficient standard.
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The system monitoring system will be able to detect failures and changes in conditions within the spacecraft this will improve the safety of the system. If modules have faults or are exposed to harsh conditions outside their normal operating ranges, this information will be logged so that when the module next returns to earth it can receive maintenance and damaged modules can be removed. For example if the system monitoring system does detect modules and systems outside their specified operating temperature range it will ask the heating system to change the temperature so that it is within the desired range. However the change in temperature could be caused by a fault in the temperature control system or the module. If the module experiences harsh conditions due to a system failure, this will be logged in the system monitoring system. Then one if not all of the modules should continue to function allowing the mission to be completed, however when the module returns to the ground, the system monitoring data can help to assesses which modules have been exposed to environments that would cause their future performance to be compromised, in which case these modules would be replace removing any risk to a future mission. D.13. Solar Cells- [EE06-01]
The efficiency of solar cells can be improved by creating a triple junction solar cell, these involve having three p-n junctions made of different materials stacked on top of one another, this allows different energies of electromagnetic waves to excite electrons creating more electron hole pairs. Triple junction cells are about 36% efficient which is a large improvement over traditional solar cells. Solar cells are grouped together into solar panels these can either be place on the outer surface of the spacecraft which does offer a degree of thermal protection. Having solar panels on all faces of the spacecraft requires the spacecraft to have more solar panels than necessary. A more efficient method is to have the solar panels in arrays outside the space craft that rotate so that they can track the sun, optimising the power that can be generated. D.14. Radioisotope Thermoelectric Generator- [EE06- 06]
High power and long duration missions, such as planetary exploration into deep space often use Radioisotope Thermoelectric Generator (RTG) power generation. This has an advantage of being able to produce a lot of power for a long time and it is not dependant on conditions outside the spacecraft such as the sun. However this kind of power generation does have several disadvantages, firstly the RTG creates a lot of radiation and a magnetic field that interferes with other electronic systems on space craft. To solve this problem RTG are often put on long arms (for example the voyager spacecraft). This radiation also means that they are not suited to
Robert Taylor D - In Orbit Operations and Electronics Page D8 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 missions that involve humans, although a RGT using Americium would not produce neutron radiation. Finally the high chance of a failure during launch creates a significant risk of spreading radioactive material across a large area of the earth, therefore RTG are built to survive a launch failure making them very heavy, therefore RTG’s are only used when other options cannot be used. D.15. Power Regulation D.15.1. Regulated power- [EE06-07] A regulated power bus in which the power system regulates the bus to a set voltage and all of the other subsystems connect directly into this regulated power supply; this system reduces the mass of the power system as all of the power regulation is done in the control system. However this system in vulnerable to failure, if any module fails the whole spacecraft is likely to lose power.
D.15.2. Unregulated power supply- [EE06-08] An unregulated power supply lets the bus voltage vary and each subsystem regulates its own voltage supply from the unregulated bus, this provides more protection to the subsystems but reduces control over the batteries charging and discharging, again making the system vulnerable to failure.
D.15.3. Hybrid system- [EE06-09] The third option is to go for a hybrid system that incorporates areas from both systems. An implementation that could be used on the proposed spacecraft would be to regulate the bus voltage in the power subsystem, allowing careful control over the power produced. Each subsystem could then have power switches that would regulate the power into that module, this adds protection, as if a module fails the power to that module is cut, protecting the power bus and the other subsystems. D.16. Payload deployment D.16.1. Robotic arm To release payloads from the spacecraft with a payload bay, a robotic arm maybe required this would need to be able to operate without destabilising the spacecraft, therefore the movements made by the arm will need to be countered with reaction wheels to conserve momentum. The robotic arm would need to have fine movement control to be able to carefully remove the payload from the payload bay.
D.16.2. Explosive bolts and springs If explosive bolts are used there will need to be a controlled detonation system that is linked to the rest of the space craft, this would require a series of interlocks to make sure that the payload is only released if the correct conditions are met. Once the explosive bolts have been fired the
Robert Taylor D - In Orbit Operations and Electronics Page D9 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 payload is normally pushed away from the spacecraft using a large spring, as this provides a reliable known force. D.17. Launch Error Detection
The path that the rocket takes will be determined using data from the IMU, if the launch control system detects that the system has flown off course, it could do one of two things depending on design. It could be set up to carry out an automatic abort in which case the system would try to return the payload to the ground safety. Another option would be that the system could give an indication to the mission control centre where a human could make a decision whether the mission should continue or be aborted. Although introducing humans into the abort system could introduce delays or human error that would increase the risk to the payload in emergency situations, it reduces the chance of false alarms which carry financial and time penalties for the company. A possible solution would be to have a hybrid system in which the system has a series of defined zones which the launch profile can be compared to. If the rocket is within zone A then the launch profile is deemed acceptable. If the launch profile is on zone B it is not ideal and a human could abort if deemed necessary. If the rocket drifts into zone C then a major failure is likely to occur and the system would automatically abort and try to land the payload safely. Unfortunately there is always the possibility of a failure on the launch pad before the rocket lifts off, this is the worst case as there is likely to be a large fire/explosion and it would be very difficult to remove the payload from the danger zone. A sensible solution would be to add extra escape options into a human payload. The loss of a human crew would be likely to ground the project for several years if not ending it completely therefore this must be avoided. D.18. Calculating the system cost
The cost for both the 25000kg and the 40000kg systems were estimated the estimation are included below. The first estimation is for the 25000kg system.
Human module $ $ power B 200000 1 200000 Additional communications module (higher speed) 150000 2 300000
OBC 201600 5 1008000 TT&C 100000 3 300000 ADCS controller 100000 3 300000 ADCS sensors 100000 3 300000 ADCS actuator 100000 3 300000 escape system controller 120000 3 360000 landing gear controller 40000 3 120000
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Data bus 20000 3 60000 temperature control system 100000 1 100000 System monitoring 80000 1 80000 heat shield deployment electronics 60000 3 180000 Total 3608000
Stage 2 $ TMR $
power 100000 1 100000 OBC 201600 5 1008000 TT&C 100000 3 300000 ADCS controller 100000 3 300000 ADCS sensors 100000 3 300000 ADCS actuator 100000 3 300000 escape system controller 120000 3 360000 landing gear controller 20000 3 60000 Data bus 10000 3 30000 System monitoring 80000 1 80000 heat shield deployment electronics 60000 3 180000 Total 3018000
stage 1 $ TMR $ power 100000 1 100000 OBC 201600 5 1008000 TT&C 100000 3 300000 ADCS controller 100000 3 300000 ADCS sensors 100000 3 300000 ADCS actuator 100000 3 300000 escape system controller 120000 3 360000 landing gear controller 20000 3 60000 System monitoring 80000 1 80000 Data bus 10000 3 30000 Total 2838000
Total 9464000 total + 10% margin 10410400
The second estimation below is for the 40000kg system
Cargo module $ TMR $
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power 100000 1 100000 OBC 201600 3 604800 TT&C 100000 3 300000 ADCS controller 100000 3 300000 ADCS sensors 100000 3 300000 ADCS actuator 100000 3 300000 escape system controller 120000 3 360000 landing gear controller 20000 3 60000 Data bus 10000 3 30000
temperature control system 20000 1 20000 heat shield deployment electronics 60000 3 180000 Total 2554800
stage 3 $ TMR $ power 100000 1 100000 OBC 201600 3 604800 TT&C 100000 3 300000 ADCS controller 100000 3 300000 ADCS sensors 100000 3 300000
ADCS actuator 100000 3 300000 escape system controller 120000 3 360000 landing gear controller 20000 3 60000 Data bus 10000 3 30000 System monitoring 80000 1 80000 heat shield deployment electronics 60000 3 180000 Total 2614800
stage 2 $ TMR $ power 100000 1 100000 OBC 201600 3 604800 TT&C 100000 3 300000 ADCS controller 100000 3 300000 ADCS sensors 100000 3 300000
ADCS actuator 100000 3 300000 escape system controller 120000 3 360000 landing gear controller 20000 3 60000 Data bus 10000 3 30000 System monitoring 80000 1 80000
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heat shield deployment electronics 60000 3 180000 Total 2614800
stage 1 $ TMR $ power 100000 1 100000 OBC 201600 5 1008000 TT&C 100000 3 300000 ADCS controller 100000 3 300000 ADCS sensors 100000 3 300000 ADCS actuator 100000 3 300000 escape system controller 120000 3 360000 landing gear controller 20000 3 60000 System monitoring 80000 1 80000 Data bus 10000 3 30000 Total 2838000
Total 10622400 total + 10% margin 11684640
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E. Re-entry Appendix Technical Financial Development Impact on other Environmental Concept Reusability Total Viability Feasibility Risk subsystems Impact RE01-01 4 5 4 2 5 3 23 RE01-02 1 2 4 2 1 5 15 RE02-01 5 5 5 4 5 5 29 RE02-02 4 5 3 2 5 5 24 RE03-01 2 5 4 5 5 5 26 RE03-02 5 4 5 4 5 4 27 RE03-03 2 1 4 4 1 1 13 RE03-04 3 4 1 1 4 5 18 Table 50- Re-entry Down Selection 1
Figure 44 - Diagram of a re-entry corridor (Sellers, 2005) E.1. Ceramic Thermal Protection system Due to the brittle nature of the ceramics, they should not carry structural load. They also should be used in tile form, as a large piece of ceramic material would be very susceptible to cracking. The space shuttle required frequent inspections to check for damaged tiles. A damaged tile was responsible for the loss of one of the shuttles, Columbia, during re-entry. Despite this problem, the STS TPS system was effective, lightweight and reusable, although expensive. The TPS system had a total weight of 8578.7 Kg, covering a surface area of 1105.0 square metres. The orbiter had a weight of roughly 100000kg on re-entry, meaning roughly 10% of the re-entry weight was the TPS (NASA, n.d.). Figure 16 shows how the surfaces are protected. The distribution is matched with where the highest temperatures are seen. The top surface is the coolest, with temperatures lower than 315°C. Advanced Flexible Reusable Surface Insulation (AFRSI) blankets are used to reduce weight and improve aerodynamic performance. This blanket is based mainly on silica fibres and is layered.
The sides of the shuttle are at an intermittent temperature. White low-temperature reusable surface insulation (LRSI) tiles are used here. The bottom surface reaches the highest temperatures of up to 1370°C, so black high temperature reusable surface insulation (HRSI) tiles are used. Both HRSI and LRSI are made of LI-900, a high performance ceramic containing 94% air (by volume) and
James Dobberson E - Re-entry Appendix Page E1 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 high purity silica glass fibres. The difference in colour affects how the material absorbs heat, with the lower surface absorbing rather than reflecting like the top surface.
The most demanding areas of the spacecraft are the wing leading edges and nose cone. Here, the highest temperatures and loads are experienced, as the air is slowed down drastically and passes through the bow shock wave. Aerodynamic forces can reach 38KN/m2 and temperatures can reach over 1600°C. Reinforced Carbon-Carbon (RCC) is used, as it is sufficiently strong and also has excellent high temperature resistance. RCC is a carbon matrix with carbon fibres embedded. These components are moulded in a single piece, varying between a quarter and a half an inch thick. (NASA, n.d.) RCC can resist high temperatures but is not a great insulator, so the back of the RCC is also very hot. This means that additional silica blankets or tiles are needed behind it.
Thermal Location Operating Area Total Weight Weight/area Protection temperature (°C) (m2) (Kg) Kg/m2 FRSI Top surface <315 332.7 532.1 1.6 LRSI Inbetween FRSI 315-650 254.6 1014.2 4.0 and HRSI, sides HRSI Bottom Surface 650-1370 peak 479.7 4412.6 9.2 RCC Nose and Wing tips 1650 38.0 1697.3 44.7 Misc Joints N/A 918.5 N/A Table 51 - Summary of Space Shuttle Thermal Protection System
Figure 45 Orbiter 10 Thermal Protection System (NASA, n.d.)
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Figure 46 Geometric altitude vs. temperature, pressure, density, and the speed of sound derived from the 1962 U.S. Standard Atmosphere (Anon., n.d.)
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E.2. Deployable heat shield Patent To estimate the weight of the panels, a spread sheet was created. Figure 47 shows how the panel surfaceareawasfound,withs,tandθbeingtheknowndimensions.Thedimensionsisequalto the radius of the craft body. The dimension t is the determined radius that the panel needs to protrude to protect the craft adequately.
θisthepanelangle.Thesurfaceareacanbefoundby: Cone width: r = s + t
Cone/nose cap height: / g =
Cone/nose cap edge length: √ / √
Surface area of a cone/nose cap: /
Surface area of capped Cone:
Figure 47 - Heat shield cone dimensions
The weight of the panel can then be found by finding the panel weight per area. The re-entry capsule has a radius of 2.5 metres. The protrusion and panel angle cannot be accurately determined, due to lack of information on the panel CD and craft length/shock angle. An estimated angle of 70° and a protrusion of 1m will be used. This gives the total radius once deployed to be 3.5 metres. The panel edge length (l-m) will equal 1.06 metres. This is also the length of craft body required by the panel when folded. The surface area of these panels is 6.86 m2
The weight of a panel with this surface area can be estimated based on the material and thickness. The patent suggests using a Carbon fibre reinforced ceramic such as SiC, which is lighter than aluminium, has roughly the same strength, but also has very good elevated temperature properties. The density is roughly 1900 Kg/m3.
The only function of the panel is to deflect the airflow and create the bow shock wave. Due to this, an infinitely thin panel would be optimal. However, the panel must be thick enough to importantly withstand the panel loading. Using a 5cm thickness, the panels have a weight of 652 kg.
This system would also require the nose cone of the craft to be heavily protected. Using Reinforced Carbon-Carbon (RCC) insulation would adequately protect the craft. The surface area depends on how flat the nose is. Using a nose which blends into the heat shield will reduce heat and load concentrations. This means the nose cone must have a 70° angle at the edges.
If a circular nose cone is used, the loading will be completely even. To have a circular shape, a 70° edge angle and a radius of 2.5 metres, the nose cone would have a height of 0.44m. The area ofadomeis2πrh,whichgivesanareaof20m2. Using the RCC weight per area of 44.7 kg/m2 found in Section D.1, this gives a nose cone weight of 905 kg. The steel structure weight can be estimated based on the panel surface area. Using a 40% ratio of structure weight/panel weight, this gives a structural weight of 261 kg.
For the rest of the craft, minor protection is required for when it is exposed during launch and also as insulation in orbit. As aerodynamics will be important, the FRSI Blankets described in
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Section E.1 would be appropriate. They are also very light weight. The craft radius is 2.5 metres, and the length is 5 metres, giving a total surface area of 78.5 m2. With a weight per surface area of 1.6 kg/m2, this gives a weight of 126 Kg for the remaining craft protection. The total TPS weight is therefore 1.94 Tonnes E.3. Parachute size Thecraftmassandearth’sgravitationalfieldcanbeconsideredconstant,meaningthatthecrafts weight is fixed. The velocity will be decreasing as the deceleration phase goes on.
For Steady velocity:
The chute area, S is given by:
∑ √
Substituting in and rearranging for the chute radius r:
,
This then allows the highest radius parachute(s) that will be required to achieve an acceleration at a given altitude and height to be found, as the craft will also produce drag. The parachute material and shape determines the coefficient of drag of the parachute.
Parachute Type Diameter (m) Weight CD (Kg) (Average) Drogue Variable Porosity Conical Ribbon 2.99 - 0.58 Pilot Ringslot 7.01 - 0.59 Main Quarter Spherical Ringsail 35.36 136 0.94 Table 52 -Summary of Parachutes to be used for Orion spacecraft (R. Olmstead, 2009) (K. Bledsoe, 2009) As the craft approaches the ground, the maximum descent speed much be satisfied. A speed of 10m/s is reasonable. As the craft can be considered to be at sea level, the radius of the parachute(s) can be found. Assuming the parachutes weight is proportional to the area (more tethers will be needed and thicker tethers will be needed for larger parachutes), the weight of each parachutes was also found by scaling the 136Kg 35.36 m diameter main parachute developed for Orion
Number of parachutes Area of each parachute Radius of each Estimated weight (m) parachute (m) of parachutes (Kg) 1 1,703 23.2 178 2 851 16.4 252 (126 X 2) 3 567 13.4 309 (103 X 3) 4 426 11.6 356 (89 X 4)
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Figure 48 - Deployable heat shield patent cover page (U. Trabandt, 2010)
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F. Landing Appendix Technical Financial Development Impact on other Environmental Concept Reusability Total Viability Feasibility Risk subsystems Impact LA01-01 4 5 4 1 5 5 24 LA01-02 4 5 5 5 5 3 27 LA01-03 3 3 3 4 3 3 19 LA01-04 1 2 2 3 1 1 10 LA01-05 3 4 5 1 5 4 22 LA01-06 5 3 5 5 5 4 27 LA01-07 1 2 2 3 3 4 15 LA02-01 5 5 4 1 5 5 25 LA02-02 5 3 3 1 4 4 21 LA02-03 5 3 5 5 4 3 25 LA02-04 4 5 4 4 5 4 26 LA02-05 4 4 3 4 3 3 21 Table 53 - Landing Down Selection 1 F.1. Wing Size The lift required by a wing for steady flight equals the weight, so:
Where L = Lift, W = Weight, CL =Coefficientofliftforthewing,ρ=Airdensity,V=Airspeed and S = wing area. CL is a function of the aerofoil shape and angle of attack. A given wing will have a defined lift curve slope, determining the angle of attack for zero lift, stall and the lift at zero incidence. This can be modified by trailing edge flaps, which will allow a craft to slow down before a landing. As the weight, air density and velocity at landing are the only two known values, the formula can be rearranged to give:
From here, the wing area can be found for a given aerofoil. The aerofoil used can be relatively thick. This is due to the large wing chord, meaning that the aerofoil can have a thick section with a relatively gentle slope. The leading edge needs to be blunt for the same reason as the front of the vehicle. As the aerofoil will have a high angle of attack, it can also be near symmetrical, reducing induced drag. A typical CL of a wing with these characteristics for landing (high angle of attack) is 0.25. This suggests a wing area of 27m2. This is 13.5m2 per wing.
ThewingareaisafunctionofthewinglengthandsweepangleΔ.Thisistheanglebetweenthe fuselage and the point at which the leading edge attaches. To prevent the wings protruding too far, a large sweep angle is best, as the area will be distributed down the length of the craft. Figure 21 shows the required wing lengths for this surface area for various wing sweep angles. The required wing root chord and lengths for each sweep angle are shown in Figure 22. The most important dimension is the wing length, as this has a direct impact on the heat shield size required. However, it is also not desirable to have a high maximum chord, as the craft will need to be longer. A 50° wing sweep gives both a short wing and a short maximum chord.
This 50 degree wing sweep gives the dimensions shown in Figure 49. The dimensions of the body must be large enough to allow the wings to fit. The weight distribution must also place the centre
James Dobberson F - Landing Appendix Page F1 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 of mass near to the centre of aerodynamic lift. This means that the body must not be too long, otherwise the mass will be too far forwards for stability.
40 Tonne
25 Tonne
3.36 50° 2.67 50°
Undefined 4.01 Undefined 3.18
Figure 49 - Diagram of wing dimensions (metres) for a 40 tonne and 25 tonne Craft
Figure 50 - Variation in delta wing surface area with sweep angle and wing maximum length
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Figure 51 - Wing dimensions required for various sweep angles for a 40 tonne and 25 tonne craft To estimate the mass of the wings, the following formula can be used. This has a 2% error when used to estimate the space shuttle wing weight (Rohrschneider, 2002)
Where: Mwing = Wing mass (to be found) Nz = Ultimate load factor (3.75 - safety factor combined with limit load) Mland = Landed mass of vehicle (40 or 25 tonnes) η=Wing/bodyefficiencyfactor(0.15foracontrolconfiguredvehicle) m2 Sbody = Plantform area of the body (~48 for a body 8 m wide and 6 m long) 2 Sexp = Exposed wing plantform area (wing surface area – 27m ) troot = Wing root thickness (1 m reasonable) Kwing = Exposed wing material/configuration constant (0.286 for an aluminium skin/stringer with stored propellant wing) 2 2 bstr =wingstructuralspanalonghalfchordline(8.70:2x√((3.36/2) + 4.01 )) Kct = wing carry through constant (0.12 for a wet carry through) bbody = maximum width of body = 8m
[ ] ( ) [
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A vertical stabiliser is also required. These are much lighter than the wings, as the load is very low, so they are much thinner and much smaller. This weight can be neglected for the purposes of estimating the wing weight. As the craft will not be powered and will glide, the lift to drag ratio is very important. This will define how far the vehicle will be able to travel once in flight.
F.2. Ground Impact shock. The shock delivered on impact is a factor of the final impact speed and the length of time the craft takes to come to a rest after touching the ground.
The acceleration can then be divided by the acceleration due to gravity g (9.81) to give the amount of G’s experienced. Impacting at 40 m/s with an impact time of 1 second, gives a deceleration for 40m/s which is 4G for example. This is a gentle impact – impacts usually have a much shorter impact time without damping, so produce much higher Gs. The maximum shock allowable is determined by the most sensitive components on board. For humans laying upside down, this is roughly 35G (Wehrly, 1987). However, this is very uncomfortable, so should be avoided. This is the most sensitive component of the craft, so gives the maximum G force. The ultimate deceleration rate reached by the main parachutes is 10m/s. an impact length of 0.1s (typical undamped impact) produces 10.91G. This would be uncomfortable for the passengers, but tolerable. The point of impact may also be damaged. Providing a damping which is increases the impact time to 1 second would produce an impact G of 1.02 G. This is very gentle, as it is equivalent to the loading experienced by the forces of gravity.
To decelerate 10,000 kg 10 m/s over 1 second requires a certain force:
However, this force needs to be applied constantly over the 1 second. Many devices, such as springs, deliver varying load based on displacement. A compensating hydraulic shock absorber achieves this linear deceleration. The deceleration force could also be split over multiple absorbers (multiple legs). Using 4 legs gives good stability and a reasonable load of 25kN per leg. Using Integration gives the distance travelled over this deceleration, which is 5 metres. This means that to achieve this deceleration, a 5 metre travel is required on the absorbers. These may be able to slide along the sides of the capsule. However, as these absorbers would be formed of impact resistant steel and would need to be considerably thick, the weight would be very large.
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G. Payloads and Markets Appendix G.1. Appendix A
Table 54 Average Flight rate, Launch price, Average capability (kg), Average Cost ($/kg) Average Launch Average Average Flight Price ($ Capability Cost 2% Mission rate million) (kg) ($/kg) Growth
People to space stations 8 637.5 15100 7868.75 157.375 GEO Sats(& MEO Sats) 10 133 4144 2138.5 42.77 LEO Sats 8 165 11135.5 5650.25 113.005 Space Station supply 8 201 14350 7275.5 145.51 Large Space Station supply 4 321.6 7175 3637.75 72.755
Space Infrastructure 8 149 14350 7249.5 144.99
Human module Payload mass fractions Human Payload fractions will be calculated for different types of vehicles namely: A multipurpose crew vehicle (MPCV), lunar lander or Martial Lander, Space exploration vehicle. Saturn V, Command module Length of commute from earth to the moon: 3days ECLSS hardware total = 2511 kg Supplies required for 12 days = 1102 kg Volume per person 5.3/3 = 1.77 cubic metres System Mass (kg) % Structure 1833 32 ECLSS hardware 2511 43 Supplies 1102 19 Crew 360 6 Total mass 5806 100
Note: Propulsion and avionics are contained in the service module Space Shuttle Volume per person = 69.7/7 = 9.96 cubic metres L 1 Lunar Lander Mass Targets Ascent Module Descent Stage System % Mass Limit % Mass Structures 20 483.20 22 1011.12 Propulsion 21 507.36 27 1240.92 Power 15 362.40 13 597.48 Avionics 13 314.08 1 45.96 ECLSS 19 579.84 24 643.44
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Crew 6 Thermal control 7 169.12 7 321.72 System (TCS) Science 0 - 16 735.36 Equipment Total Mass 100 2416 100 4596
Other space vehicles
DELTA- Advanced Reusable Transport (DART) An alternative manned Spacecraft System Mass (kg) % Structures (inc. Escape 10%) 1107 23 Systems (interface) 7 1 Propulsion 1477 31 Avionics 573 12 ECLSS hardware 636 14 Supplies 574 12 Crew (x5) 400 8 Total 4772 100
The Mass fraction of the DART spacecraft will be utilised It takes two days to approach the ISS. So it would take 4 days to make a return journey Mission length is 4 days to 200km. The fuel required for a 25000kg to travel in the lower atmosphere (up to 1000km) would require a propellant mass of 5732kg leaving 19267kg for the restofthespacecraft.5732/25000*100=23%(FromRobert’svalues) Supplies for 5 days = 30.60 x 5 x 5 =765kg
Crew each weighing an average of 80 kg, Supplies for 4 days = 25 x 30.60 x 6 = 4590 kg (including 2 days back up supplies) Configuration for 25 people Table 55 Mass fraction for 25 man configuration System Mass (kg) % Structures 6000 24 Propulsion 5732 23 Avionics 3000 12 ECLSS + Supplies 3500 14 Supplies 4590 19 Crew (x25) 2000 8 Total 250000 100
Configuration for 10 people This configuration will initially be used in order for the vehicle to comply with safety standards. The remaining mass fraction will be added to the cargo mass fraction. Table 56 - Mass fractions for 10 person configuration System Mass (kg) % Structures 1940 7.76 Escape System 1560 6.24 Heat Shield 2500 10 Propulsion 7750 31 Avionics 3000 12 ECLSS Hardware 3500 14 Supplies 1836 7.3
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Crew (x10) 800 3.2 Cargo 2125 8.5 Total 25000 100
The space hotel is assumed to be in close proximity 200km, so fraction of propulsion is subject to decrease compared to that of the DELTA to 500km
Figure 52 Payload Fractions for Manned Vehicle carrying 10 people
Figure 53 Payload Fractions for Manned Vehicle with 25 people
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G.2. Appendix C
Figure C-1: Space shuttle Waste Collection System(WCS) (Keith, 1998)
Figure C-1:Soyuz’sWCS(Malik, 2012)
Mukudzei Muchengeti G - Payloads and Markets Appendix Page G4 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
H. Mass Estimating Relationships H.1. Mass Estimating Equations
The main body of the launch system can be described by the following:
The oxidizer and fuel tank are estimated through the following equations:
Where is the density of fuel and is the volume of the tank.
The anti-vortex's mass in the tank can be described by the following:
̇ ̇
Where ̇ and ̇ is the fuel mass flow rate and oxidiser fuel mass flow rate respectively. and is the density of the fuel and density of the oxidiser respectively.
The slash baffle's mass in the tank can be described by the following:
Where is the diameter of the body and is the volume of the oxidiser in the tank.
The inter-tank's mass can be described by the following for stage 1:
Where is the inter-tank wetted area.
The inter-tank's mass can be described by the following for stage 2:
The forward skirt's mass can be described by the following for stage 1:
Where is the forward skirt wetted area.
The engine compartment's mass can be described by the following for stage 1:
William Wou H - Mass Estimating Relationships Page H1 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Where is the engine compartment wetted area.
The engine compartment's mass can be described by the following for stage 2:
The aft skirt's mass can be described by the following:
[
]
Where is the aft skirt wetted area, is the maximum dynamic pressure on the body
The thrust structure's mass can be described by the following:
( )
Where is the vacuum thrust of the engines and is the number of engines
The engine's mass can be described by the following:
( )
The engine installation's mass can be described by the following:
The engine subsystem's mass can be described by the following:
The thrust vector control's mass can be described by the following:
The purge system's mass can be described by the following:
Where is the volume of the main tank
The feed system's mass can be described by the following:
̇ Where is the mass flow rate of the propellant.
The primary power group's mass can be described by the following:
( )
Where is the mass of the propellant.
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The hydraulic systems group's mass can be described by the following:
H.1.1. Mass Estimates for each Variant
Table 57 shows the mass estimates for Variant-A stage 2
Sub-Section Mass (kg) Tank Mass 230 Anti-vortex 14 Slash Baffles 22 Inter-tank 1,708 Engine Compartment 327 Thrust Structure 233 Engines 5,459 Engine Install 1,210 Engine Thrust vectoring 1,272 Engine Sub System 742 Propellant Purge 9.5 Feed Sub Systems 1,239 Avionics 303 Primary Power Group 451 Hydraulic System 360 Heat Shield 2,173 Payload 25,000
Table 57 showing the mass estimates for Variant-A stage 2 Table 58 shows the mass estimates for the Variant-B stage 2
Sub-Section Mass (kg) Tank Mass 580
Anti-vortex 28
Slash Baffles 45
Inter-tank 3,417
Engine Compartment 656
Thrust Structure 468
Engines 13,649
Engine Install 4,036
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Engine Thrust vectoring 2,546
Engine Sub System 1,485
Propellant Purge 36
Feed Sub Systems 2,478
Avionics 607
Primary Power Group 1,255
Hydraulic System 721
Heat Shield 5,179
Table 58 showing the mass estimates for Variant-B stage 2 Table 59 shows the mass estimates for the Variant-B stage 3
Sub-Section Mass (kg) Tank Mass 203
Anti-vortex 15
Slash Baffles 45
Inter-tank 1,863
Engine Compartment 656
Thrust Structure 145
Engines 4,899
Engine Install 1,345
Engine Thrust vectoring 849
Engine Sub System 495
Propellant Purge 11
Feed Sub Systems 1,406
Avionics 320
Primary Power Group 995
Hydraulic System 721
Heat Shield 1,562
Table 59 showing the mass estimates for Variant-B stage 3. H.1.2. Turnaround Modelling
Modelling the turnaround for the reusable launch vehicle was developed by using formulas that were derived using operational statistical data. Each major system of the launch system has its own empirical formula that estimates the length of time needed to fully service the sub-system. Once the total amount of hours for each stage was added up, a number of ground staff was
William Wou H - Mass Estimating Relationships Page H4 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3 assigned to maintenance. This calculated how much the service would cost from man power alone and how long it would take depending the ground staff's work shifts. Assumptions were made for the of standard pay and duration of shifts for the ground staff to determine the direct operational cost. The ground staff worked 8 hours shifts, 5 days a week, 46 weeks a year. Furthermore, the sub-system's maintenance requirements were taken as constants from (Pearson, et al., 2000) due to the lack of data on these sub-systems for this report.
Once the calculated maintenance time for each sub-section was determined, this maintenance time had to convert into how long the maintenance line would take to service each variant. First, the engines from each stage were taken to another maintenance line to receive servicing which allowed the stage engine to be served in parallel streams. Moreover, each stage from both RLS variants allowed for parallel work to occur due to the proposed three stage maintenance line.
Furthermore, by using the automated monitoring system which reduces the overall inspection time required and by taking 50 ground staff at each maintenance line, the amount of days required to maintain a give variant can be calculated. These estimated times are conservative estimates assuming an efficient maintenance team and process activities.
The equations used are shown below:
The amount of hours required for the vehicle's structure is shown below (Scanlan, 2004).
( ( ) ( ))
( ( ) )
Where is the mass of the structure.
The amount of hours required for the vehicle's engines is shown below (Gstattenbauer, 2006).
[ ( ) ]
To calculate the amount of hours to service the thermal protection system, the wetted area of the vehicle had to be calculated first (Gstattenbauer, 2006):
The Second stage thermal protection system maintenance time is calculated below:
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H.1.3. Finance equations
Development Cost for an expendable launch system and reusable launch system derived by (Koelle, 2002) are shown below.
Where is the mass of each stage. Both these equations results in man-years which is equivalent to $300,000 in the year 2012.
The production cost for an expendable launch system and reusable launch system derived by (Koelle, 2002) are shown below.
The engine development cost for an expendable launch system and reusable launch system derived by (Koelle, 2002) are shown below.
The engine production cost for an expendable launch system and reusable launch system derived by (Koelle, 2002) are shown below.
The pre-launch cost for a launch vehicle derived by (Keith, n.d.) describes the cost for assembly, Integration and fuelling of the vehicle before launch.
Where: is the mass of the vehicle before fuelling is the number of launches per year is the number of stages on the vehicle is the vehicle launch type, for multistage ELV (Expendable Launch Vehicle) fv is 1, for a RLV (Reusable Launch Vehicle) fv is 0.7 fo is the assembly and Integration mode which is set to 0.5 for horizontal assembly and pad erection.
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H.1.4. Further Finance graphs
Cost/Kg for various rocket systems
15,000
10,000
5,000 Cost/Kg Cost/Kg (USDollars/Kg) 0 Atlas V Ariane 5 Launch Systems Space Shuttle Expendable Falcon 9 Reusable Variant A-Crewed Falcon Heavy
Figure 54 showing the RLS crewed vehicle against leading competitors in terms of cost/kg H.1.5. Finance Calculation Results for all variants
The below table shows the calculated operating costs for the RLS Variant-A with crewed missions
Reusable Missions Fleet Operation Launch Cost Total Cost per Cost per Year per year Required Cost (USD) (USD) annum (USD) mission (USD) 2,030 4 1 880,190,000 5,230,970,000 6,111,160,000 1,527,790,000 2,031 4 1 404,190,000 5,230,970,000 5,635,160,000 1,408,790,000 2,032 4 1 404,190,000 5,230,970,000 5,635,160,000 1,408,790,000 2,033 4 1 404,190,000 5,230,970,000 5,635,160,000 1,408,790,000 2,034 4 1 404,190,000 5,230,970,000 5,635,160,000 1,408,790,000 2,035 10 2 896,770,000 2,293,170,000 3,682,180,000 368,220,000 2,036 10 2 420,780,000 2,293,170,000 3,206,190,000 320,620,000 2,037 13 2 429,070,000 1,810,870,000 2,732,180,000 210,170,000 2,038 13 2 429,070,000 1,810,870,000 2,732,180,000 210,170,000 2,039 13 2 429,070,000 1,810,870,000 2,732,180,000 210,170,000 2,040 13 2 429,070,000 1,810,870,000 2,732,180,000 210,170,000 2,041 14 2 431,830,000 1,694,030,000 2,618,100,000 187,010,000 2,042 14 2 431,830,000 1,694,030,000 2,618,100,000 187,010,000 2,043 14 2 431,830,000 1,694,030,000 2,618,100,000 187,010,000 2,044 14 2 431,830,000 1,694,030,000 2,618,100,000 187,010,000 2,045 22 3 929,950,000 1,127,860,000 2,550,050,000 115,910,000 2,046 22 3 453,950,000 1,127,860,000 2,074,050,000 94,280,000 2,047 23 3 456,720,000 1,083,630,000 2,032,590,000 88,370,000 2,048 23 3 456,720,000 1,083,630,000 2,032,590,000 88,370,000 2,049 24 3 459,480,000 1,042,910,000 1,994,630,000 83,110,000 2,050 20 3 448,420,000 1,228,880,000 2,169,540,000 108,480,000
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2,051 21 3 451,190,000 1,176,090,000 2,119,510,000 100,930,000 2,052 21 3 451,190,000 1,176,090,000 2,119,510,000 100,930,000 2,053 21 3 451,190,000 1,176,090,000 2,119,510,000 100,930,000 2,054 22 3 453,950,000 1,127,860,000 2,074,050,000 94,280,000 Table 60 showing the financial results from the financial analysis for a RLS Variant-A crewed missions
Table 61 shows the calculated operating costs for an equivalent ELS Variant-A with crewed missions Expendable Missions Operation Launch Cost Total Cost per Cost per Year Fleet Required per year Cost (USD) (USD) annum (USD) mission (USD) 2,030 4 3,736,410,000 5,040,550,000 1,260,140,000 2,550,000,000 2,550,000,000 2,031 4 3,736,410,000 5,040,550,000 1,260,140,000 2,550,000,000 2,550,000,000 2,032 4 3,736,410,000 5,040,550,000 1,260,140,000 2,550,000,000 2,550,000,000 2,033 4 3,736,410,000 5,040,550,000 1,260,140,000 2,550,000,000 2,550,000,000 2,034 4 3,736,410,000 5,040,550,000 1,260,140,000 2,550,000,000 2,550,000,000 2,035 10 1,637,980,000 4,158,360,000 415,840,000 6,375,000,000 6,375,000,000 2,036 10 1,637,980,000 4,158,360,000 415,840,000 6,375,000,000 6,375,000,000 2,037 13 1,293,480,000 4,421,990,000 340,150,000 8,287,500,000 8,287,500,000 2,038 13 1,293,480,000 4,421,990,000 340,150,000 8,287,500,000 8,287,500,000 2,039 13 1,293,480,000 4,421,990,000 340,150,000 8,287,500,000 8,287,500,000 2,040 13 1,293,480,000 4,421,990,000 340,150,000 8,287,500,000 8,287,500,000 2,041 14 1,210,020,000 4,541,240,000 324,370,000 8,925,000,000 8,925,000,000 2,042 14 1,210,020,000 4,541,240,000 324,370,000 8,925,000,000 8,925,000,000 2,043 14 1,210,020,000 4,541,240,000 324,370,000 8,925,000,000 8,925,000,000 2,044 14 1,210,020,000 4,541,240,000 324,370,000 8,925,000,000 8,925,000,000 2,045 22 805,620,000 5,758,490,000 261,750,000 14,025,000,000 14,025,000,000 2,046 22 805,620,000 5,758,490,000 261,750,000 14,025,000,000 14,025,000,000 2,047 23 774,020,000 5,929,610,000 257,810,000 14,662,500,000 14,662,500,000 2,048 23 774,020,000 5,929,610,000 257,810,000 14,662,500,000 14,662,500,000 2,049 24 744,930,000 6,103,230,000 254,300,000 15,300,000,000 15,300,000,000 2,050 20 877,770,000 5,425,230,000 271,260,000 12,750,000,000 12,750,000,000 2,051 21 840,060,000 5,590,230,000 266,200,000 13,387,500,000 13,387,500,000 2,052 21 840,060,000 5,590,230,000 266,200,000 13,387,500,000 13,387,500,000 2,053 21 840,060,000 5,590,230,000 266,200,000 13,387,500,000 13,387,500,000 2,054 22 805,620,000 5,758,490,000 261,750,000 14,025,000,000 14,025,000,000 Table 61 showing the financial results from the financial analysis for a ELS Variant-A crewed missions
Table 62shows the calculated operating costs for the RLS Variant-A with unmanned missions
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Reusable Missions Fleet Operation Cost Launch Cost Total Cost per Cost per mission Year per year Required (USD) (USD) annum (USD) (USD) 2,024 4 1 424,710,000.00 1,729,650,000.00 2,154,360,000.00 538,590,000.00 2,025 4 1 10,930,000.00 1,729,650,000.00 1,740,580,000.00 435,150,000.00 2,026 4 1 10,930,000.00 1,729,650,000.00 1,740,580,000.00 435,150,000.00 2,027 5 1 13,670,000.00 1,414,940,000.00 1,428,610,000.00 285,720,000.00 2,028 5 1 13,670,000.00 1,414,940,000.00 1,428,610,000.00 285,720,000.00 2,029 5 1 13,670,000.00 1,414,940,000.00 1,428,610,000.00 285,720,000.00 2,030 9 2 568,090,000.00 833,670,000.00 1,401,760,000.00 155,750,000.00 2,031 9 2 154,310,000.00 833,670,000.00 1,416,020,000.00 157,340,000.00 2,032 9 2 154,310,000.00 833,670,000.00 1,416,020,000.00 157,340,000.00 2,033 9 2 154,310,000.00 833,670,000.00 1,416,020,000.00 157,340,000.00 2,034 9 2 154,310,000.00 833,670,000.00 1,416,020,000.00 157,340,000.00 2,035 9 2 154,310,000.00 833,670,000.00 1,416,020,000.00 157,340,000.00 2,036 9 2 154,310,000.00 833,670,000.00 1,416,020,000.00 157,340,000.00 2,037 11 2 159,780,000.00 695,920,000.00 1,283,730,000.00 116,700,000.00 2,038 11 2 159,780,000.00 695,920,000.00 1,283,730,000.00 116,700,000.00 2,039 11 2 159,780,000.00 695,920,000.00 1,283,730,000.00 116,700,000.00 2,040 12 2 162,510,000.00 643,500,000.00 1,234,050,000.00 102,840,000.00 2,041 12 2 162,510,000.00 643,500,000.00 1,234,050,000.00 102,840,000.00 2,042 13 2 579,030,000.00 598,770,000.00 1,605,830,000.00 123,530,000.00 2,043 13 2 165,250,000.00 598,770,000.00 1,192,060,000.00 91,700,000.00 2,044 13 2 165,250,000.00 598,770,000.00 1,192,060,000.00 91,700,000.00 2,045 13 2 165,250,000.00 598,770,000.00 1,192,060,000.00 91,700,000.00 2,046 13 2 165,250,000.00 598,770,000.00 1,192,060,000.00 91,700,000.00 2,047 14 2 167,980,000.00 560,140,000.00 1,156,150,000.00 82,580,000.00 2,048 14 2 167,980,000.00 560,140,000.00 1,156,150,000.00 82,580,000.00 2,049 14 2 167,980,000.00 560,140,000.00 1,156,150,000.00 82,580,000.00 2,050 12 2 162,510,000.00 643,500,000.00 1,234,050,000.00 102,840,000.00 2,051 12 2 162,510,000.00 643,500,000.00 1,234,050,000.00 102,840,000.00 2,052 13 2 165,250,000.00 598,770,000.00 1,192,060,000.00 91,700,000.00 2,053 13 2 165,250,000.00 598,770,000.00 1,192,060,000.00 91,700,000.00 2,054 13 2 165,250,000.00 598,770,000.00 1,192,060,000.00 91,700,000.00 Table 62 showing the financial results from the financial analysis for a RLS Variant-A crewed missions
Table 63 shows the calculated operating costs for the RLS Variant-B with unmanned missions Reusable Missions Fleet Total Cost per Year Operation Cost Launch Cost Cost per mission per year Required annum 2,024 3 1 518,050,000.00 3,993,390,000.00 4,511,440,000.00 1,503,810,000.00 2,025 3 1 9,010,000.00 3,993,390,000.00 4,002,400,000.00 1,334,130,000.00 2,026 3 1 9,010,000.00 3,993,390,000.00 4,002,400,000.00 1,334,130,000.00 2,027 3 1 9,010,000.00 3,993,390,000.00 4,002,400,000.00 1,334,130,000.00 2,028 3 1 9,010,000.00 3,993,390,000.00 4,002,400,000.00 1,334,130,000.00
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2,029 3 1 9,010,000.00 3,993,390,000.00 4,002,400,000.00 1,334,130,000.00 2,030 10 2 539,070,000.00 1,351,300,000.00 1,890,370,000.00 189,040,000.00 2,031 10 2 30,030,000.00 1,351,300,000.00 1,381,330,000.00 138,130,000.00 2,032 10 2 30,030,000.00 1,351,300,000.00 1,381,330,000.00 138,130,000.00 2,033 12 2 36,040,000.00 1,146,800,000.00 1,182,840,000.00 98,570,000.00 2,034 12 2 36,040,000.00 1,146,800,000.00 1,182,840,000.00 98,570,000.00 2,035 13 2 39,040,000.00 1,067,090,000.00 1,106,130,000.00 85,090,000.00 2,036 14 2 998,350,000.00 998,240,000.00 2,543,520,000.00 181,680,000.00 2,037 17 3 1,079,730,000.00 838,200,000.00 2,464,870,000.00 144,990,000.00 2,038 17 3 1,007,360,000.00 838,200,000.00 2,392,490,000.00 140,730,000.00 2,039 18 3 1,010,360,000.00 796,170,000.00 2,353,460,000.00 130,750,000.00 2,040 18 3 1,010,360,000.00 796,170,000.00 2,353,460,000.00 130,750,000.00 2,041 19 3 1,522,400,000.00 758,360,000.00 2,827,690,000.00 148,830,000.00 2,042 19 3 1,013,360,000.00 758,360,000.00 2,318,650,000.00 122,030,000.00 2,043 20 3 1,016,370,000.00 724,140,000.00 2,287,440,000.00 114,370,000.00 2,044 20 3 1,016,370,000.00 724,140,000.00 2,287,440,000.00 114,370,000.00 2,045 25 4 1,540,420,000.00 592,390,000.00 2,679,740,000.00 107,190,000.00 2,046 25 4 1,031,380,000.00 592,390,000.00 2,170,700,000.00 86,830,000.00 2,047 26 4 1,034,380,000.00 571,840,000.00 2,153,160,000.00 82,810,000.00 2,048 26 4 1,034,380,000.00 571,840,000.00 2,153,160,000.00 82,810,000.00 2,049 27 4 1,037,390,000.00 552,740,000.00 2,137,060,000.00 79,150,000.00 2,050 23 3 1,025,380,000.00 638,550,000.00 2,210,860,000.00 96,120,000.00 2,051 24 3 1,028,380,000.00 614,550,000.00 2,189,870,000.00 91,240,000.00 2,052 24 3 1,028,380,000.00 614,550,000.00 2,189,870,000.00 91,240,000.00 2,053 25 4 1,031,380,000.00 592,390,000.00 2,170,700,000.00 86,830,000.00 2,054 26 4 1,034,380,000.00 571,840,000.00 2,153,160,000.00 82,810,000.00 Table 63 showing the financial results from the financial analysis for a RLS Variant-B crewed missions
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Table 64 shows the total revenue and payment for the proposed RLS system showing the income generated and revenue produced for all variants
Missions per Year year Total Income Revenue 2,024 7 1,131,000,000.00 -5,534,800,000.00 2,025 7 1,131,000,000.00 -4,611,980,000.00 2,026 7 1,131,000,000.00 -4,611,980,000.00 2,027 8 1,296,000,000.00 -4,135,010,000.00 2,028 8 1,296,000,000.00 -4,135,010,000.00 2,029 8 1,296,000,000.00 -4,135,010,000.00 2,030 23 5,636,000,000.00 -3,767,280,000.00 2,031 23 5,636,000,000.00 -2,796,500,000.00 2,032 23 5,636,000,000.00 -2,796,500,000.00 2,033 25 6,279,200,000.00 -1,954,810,000.00 2,034 25 6,279,200,000.00 -1,954,810,000.00 2,035 32 10,425,800,000.00 4,221,470,000.00 2,036 33 10,425,800,000.00 3,260,070,000.00 2,037 41 13,526,900,000.00 7,046,120,000.00 2,038 41 13,526,900,000.00 7,118,500,000.00 2,039 42 13,526,900,000.00 7,157,520,000.00 2,040 43 13,691,900,000.00 7,372,210,000.00 2,041 45 14,478,400,000.00 7,798,560,000.00 2,042 46 14,812,400,000.00 8,269,810,000.00 2,043 47 14,812,400,000.00 8,714,800,000.00 2,044 47 14,961,400,000.00 8,863,800,000.00 2,045 60 21,026,200,000.00 14,604,350,000.00 2,046 60 21,175,200,000.00 15,738,390,000.00 2,047 63 21,977,700,000.00 16,635,800,000.00 2,048 63 22,259,700,000.00 16,917,800,000.00 2,049 65 22,897,200,000.00 17,609,350,000.00 2,050 55 19,244,600,000.00 13,630,150,000.00 2,051 57 20,015,100,000.00 14,471,670,000.00 2,052 58 20,365,100,000.00 14,863,660,000.00 2,053 59 20,365,100,000.00 14,882,830,000.00 2,054 61 21,151,600,000.00 15,732,330,000.00 Table 64 showing the payment and revenue for the RLS for all variants over the project lifetime
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H.1.6. Sensitivity analysis
Sensitivity Analysis on reocurring operating costs 7,000
6,000
5,000
4,000
3,000
2,000 US Dollars US (Millions) 1,000
0 2030 2032 2034 2036 2038 2040 2042 2044 2046 2048 2050 2052 2054 Year
High Estimates Low Estimates Expected Costs
Figure 55 showing a sensitivity analysis for operational costs for the RLS
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I. Project Management I.1. Gantt Chart Group 2 - Project Plan
Week 1 Week 2 Week 3 Week 4 Week 5 Week 6 Week 7 Week 8 Week 9 Week 10 Week 11 Week 12 Week 13 Week 14 Week 15
M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su M T W T F S Su Market Research I I Research F n n Electronics and i Payload Options c c Orbital Systems a e e Propulsion n Launch Vehicle and p p Reentry l t t Structures and Fuel Propulsion i i Payload R o o Integration e Risk n n p Concept Development o Interpretation of R R Phase 1 r e e Electronics and Orbital t Others p p Systems o o Propulsion D r r Reentry u P t t Structures and Fuel e o Payload s Assemble t D Integration t o u e e r P Concept Development O
r Phase 2 r P i Electronics and Orbital a r n Systems l e t Propulsion s Reentry E e Structures and Fuel x n Integration a t Concept Optimisation m a Electronics and Orbital i t Systems n i Propulsion a o Reentry t n Structures and Fuel i
Integration o E System Integration and Launch System Concept n v e n i n (Human) Payload Research, Assessment and Integration g
Infrastructure, Concept of Operations and Performance
Individual Financial Assessment
Whole Life Cost
Write up and Compile Review and Print
Figure 56 Project Final Gantt Chart
Tim West Section 23 - Appendices Page 1 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
I.2. Group Actions List Action Action Week Number Date Action Week # on Date Due Due e.g. 4-1 Issued Issued Action (initial) By By Completed All to make sure that minutes are reviewed at begin of every formal Y 3.1 25/10/2012 Wk4 meeting (1/11/2012) ALL O.G - 3.2 25/10/2012 Wk4 Actions to be numbered –Robert- 26/10/2012 RT 26/10/2012 Wk4 Y 3.3 25/10/2012 Wk4 Tim to also keep a list of action for the entire project - 31/10/2012 TW 31/10/2012 Wk5 Y Mukudzei to do more in-depth payload analysis from research with Y the aim of defining the volume required – ongoing task (review 3.4 25/10/2012 Wk4 progress 1/11/2012) MM O.G - Robert to share electronics information with the group Tuesday Y 3.5 25/10/2012 Wk4 30/10/2012 RT 31/10/2012 Wk5 William to start work on producing a budget -ongoing task (review Y 3.6 25/10/2012 Wk4 progress 1/11/2012) WW 01/11/2012 Wk5 3.7 25/10/2012 Wk4 Supervisors to give feedback next week 1/11/2012 Su 01/11/2012 Wk5 Y Tim to create spread sheet of critical path– ongoing task (review Y 3.8 25/10/2012 Wk4 progress 30/10/2012) TW 30/10/2012 Wk5 Robert to create agenda for the next formal meeting to be emailed out Y 3.9 25/10/2012 Wk4 30/10/2012 RT 30/10/2012 Wk5 3.10 25/10/2012 Wk4 To have an integration slot in the meeting on 6/11/2012 - Robert RT 06/11/2012 Wk6 Y 4.1 01/11/2012 Wk5 Prepare for the first down selection session on Tuesday ALL 06/11/2012 Wk6 Y Prepare a formal down selection criteria and record and agree a Y 4.2 01/11/2012 Wk5 method for recording it in the final report. ALL 06/11/2012 Wk6 5.1 To finalised the definition on how the down selection processes is to be Y 08/11/2012 Wk6 continued to a point where the group is satisfied it. All 15/11/2012 wk7 5.2 08/11/2012 Wk6 To ask the supervisors about the tense of the final report All 15/11/2012 wk7 Y
Tim West Section 23 - Appendices Page 1 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
5.3 08/11/2012 Wk6 Determine the pressure of the payload module for the crew MM 15/11/2012 Wk7 Y Generate a Delta-V and mass model for the rocket and space-plane TW, Y 5.4 08/11/2012 Wk6 system WW,MC 22/11/2012 wk7 5.5 08/11/2012 Wk6 Generate a cost and operation model WW 22/11/2012 wk7 Y All Y minus 6.1 15/11/2012 Wk7 To produce 20 page sections for report, ready for integration MM 22/11/2012 Wk8 6.2 15/11/2012 Wk7 To fix orbit transfer model RT 20/11/201 Wk8 Y 6.3 15/11/2012 Wk7 To look at saftey during the intergration phase All 27/11/2012 Wk10 Y 7.1 22/11/2012 Wk8 Safety through integration. All 06/12/2012 Wk10 Y 7.2 22/11/2012 Wk8 Baked bean v Space plane down selection JD 29/11/2012 Wk9 Y 7.3 22/11/2012 Wk8 To discussion about the various operations at the launch pad WW 27/11/2012 Wk9 Y Assess the number of missions of various types per year, throughout 30 Y 8.1 29/11/2012 Wk9 year service MM 06/12/2012 Wk10 8.2 29/11/2012 Wk9 Cost analysis of electronics RT 06/11/2012 Wk10 Y 8.3 29/11/2012 Wk9 Life time estimates of components ALL 05/11/2012 Wk10 Y 8.4 29/11/2012 Wk9 To calculate mass and shape of electronics bay for rocket stages MC, RT 06/11/2012 Wk10 Y TW, RT, Y 8.5 Record launch, inflight, landing operation plan 04/11/2012 29/11/2012 Wk9 JD Wk10 9.1 06/12/2012 WK10 Look into the effects of sensitivity analysis when costs are not certain. WW O.G. Wk11 Y 9.2 06/12/2012 WK10 Finalise the cost of electronics RT 13/12/2012 Wk11 Y 9.3 06/12/2012 WK10 Re-check the speeds associated with parachutes JD 13/12/2012 Wk11 Y 9.4 06/12/2012 WK10 Specifically distribute the 30 pages to individuals ALL 11/12/2012 Wk11 Y To review the diameter of the human module with a view to tapering JD, Y 10.1 13/12/2012 WK11 the rocket above or at Stage 2A for Human Payloads. MC,MM 08/01/2012 Wk12
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I.3. Concept Summary Table 65: Group 2 Concept Summary Concept ID Concept Name 1 2 3 4 5 6 7 8 9 Launch Philosophy LP00-00 Staged Expendable (Benchmark) LP01 Vertical Launch Rockets LP01-01 Staged Reusable LP01-01a Staged Reusable First Stage LP01-01b Staged All Stages Reused LP01-02 Staged Reusable with Returning Orbiter LP02 Captive Launch LP02-01 Captive Air Launch Subsonic LP02-02 Captive Air Launch Supersonic LP02-03 Captive Ground Launch LP03 Space Plane LP03-01 Jettisonable Atmospheric Flight System LP03-02 Ground Assisted Runway Accelerator LP03-03 Boosted Ground Launch LP03-04 Unassisted Ground Launch LP04 Exotic Concepts LP04-01 Electromagnetic Rail Accelerator Propulsion PR00 Current Generation Rocket Technology PR01 Optimisation of Rocket Technology PR01-01 Alternative Fuels PR01-02 Thrust Augmentation Nozzles PR01-03 Aerospike Engines PR01-04 Dual -Expander/Throat Nozzles PR01-05 Solid Fuel Rockets PR01-06 Hybrid Rockets PR02 Atmospheric Oxygen Propulsion Systems PR02-01 Air-Breathing Rockets PR02-01 SC/RAM Assisted Engines PR03 Non-Rocket Propulsion PR03-01 Pulse Detonation Engines PR03-02 Nuclear Thermal Rockets Re-entry RE01 High altitude deceleration RE01-01 Aerodynamic drag RE01-02 Reverse thrust RE02 Low altitude deceleration RE02-01 Parachutes RE02-02 Flight RE01 Thermal Protection System RE03-01 High performance alloys RE03-02 Ceramics RE03-03 Ablative heat shielding RE03-04 Deployable heat shield
Tim West Section 23 - Appendices Page 1 Group 2 Re-usable Launch and Payload Delivery System MDDP 2012/3
Landing LA01 Landing method LA01-01 Runway with landing gear LA01-02 Vertical landing with landing gear LA01-03 Vertical landing with thrusters LA01-04 Vertical landing with Sky Crane LA01-05 Horizontal landing on water LA01-06 Vertical landing on water LA01-07 Mid-air retrieval LA02 Landing location LA02-01 Runway on Site LA02-02 Runway LA02-03 Water Landing LA02-04 Vertical landing on Site LA02-05 Vertical landing OP01 Orbit Insertion OP01-01 Direct insertion OP02 Transfer space craft from a lower orbit OP03 transfer payload from a lower orbit Electronics EE01 On boards data Handling EE01-01 Cots components EE01-02 Rad hard components EE01-03 Mil standard components EE01-04 FPGA EE01-05 Processor EE01-06 Computer EE01-07 Analogue computer EE01-08 Distributed EE01-09 Central processing EE01-10 Clock speed (slow, fast, non) EE02 Spacecraft Bus EE02-01 Centralised (Star) EE02-02 Distributed (ring) EE02-03 bus EE02-04 wireless (like to fail) EE02-05 CAN EE02-06 1553 + 1760 (Aircraft) EE02-07 spacewire EE02-08 I2C EE02-09 Dual Bus system?(Space/Aircraft) EE04 Communications EE04-01 Laser up link 1 meter receiver? EE04-02 Helical antenna, EE04-03 phased array EE04-04 disk EE04-05 Dipole EE04-06 Am modulation EE04-07 FM modulation
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EE04-08 Phase modulation EE04-09 Multiple access EE04-10 single ground station EE04-11 C-band EE04-12 S-bans EE03 Telemetry EE03-01 same as main communication system Separate from the main communication EE03-02 system EE05 ADCS EE05-01 Re-entry control? EE05-02 Gravity gradient EE05-04 Spin stabilised EE05-03 3 axis EE05-05 Sun sensors EE05-06 Magnetometers EE05-07 GPS EE05-08 Star camera EE05-09 Horizon sensors EE05-10 Magnetorquers EE05-11 Reaction wheels EE05-12 control moment Gyros EE05-13 Thrusters EE05-14 Re-entry flaps EE05-15 Deploying parachutes EE05-16 Electrostatic thrusters (PPT) EE05-17 IMU EE06 Power EE06-01 Solar cells (multi junction cells) EE06-02 Hydrogen fuel cells EE06-03 Primary Batteries EE06-05 Batteries (NiCd, NiMH, NiH2, Li-ion) EE06-06 RGT EE06-07 regulated bus EE06-08 unregulated bus EE06-09 Hybrid distribution EE07 System Monitoring (integrations task) EE07-01 temperature sensors EE07-02 stress and strain gauges EE07-03 G- force sensors EE07-04 Accelerometers EE07-05 Fuel sensors EE07-06 Atmospheric monitoring (human payload) EE07-07 Radiation monitoring EE07-08 Radar to monitor outside EE08 Payload Drop off - (integration task) EE08-01 Robotic arms EE08-02 Spring EE08-03 Explosive bolts EE08-04 release catch
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EE08-05 Rockets/jets EE09 Temperature Control EE09-01 run hot EE09-02 run cold EE09-04 heat shielding - Active control EE09-06 Pumped cooling loops EE09-07 Variable conductive heat pipes EE09-08 Heaters EE09-09 Louvers and shutters EE09-10 Refrigeration and heat pumps EE09-11 Radiators Passive control EE09-11 heat pipes EE09-12 coatings EE09-13 controlled conductive paths EE09-14 MLI Electronics for payload/life support EE10 (integration tasks) EE10-01 power connection EE10-02 data connection EE10-03 communication connection EE10-04 own communication system EE11 Aeronautics (if required) EE11-01 control surface control EE11-02 airflow monitoring EE11-03 navigation EE11-04 flight control (auto pilot) EE11-05 Auto landing Structures ST01 Fuselage Structure ST01-01 Monocoque ST01-02 Semi monocoque ST01-03 Truss ST02 Tank Structure ST02-01 Artificial gravity induced by spin ST02-02 Positive Expulsion Tank ST02-03 Surface Tension Device ST03 Internal Stiffeners and Support Structures ST03-1 I beam ST03-2 T beam ST03-3 Circular beam ST03-4 Square beam ST03-5 Hollow Sections ST03-6 Sandwich Panels IF01 Launch Platforms LF01-01 Sea Launch LF01-02 Land Launch LF01-03 Submarine Launch LF01-04 Air Launch
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LF01-05 Missile Silo LF02 Transportation to launch pad LF02-01 Mobile Crawler (Vertical) LF02-02 Mobile Crawler (horizontal) LF02-04 Silo LF02-04 Modules to be transported and assembled LF03 Infrastructure LF03-01 Built new infrastructure LF03-02 Use existing infrastructure
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