Proposal for a Commercial Orbital Transportation System: Preliminary Specifications for an Innovative Approach

Transcription

1 AIAA SPACE 2009 Conference & Exposition September 2009, Pasadena, California AIAA American Institute of Aeronautics and Astronautics Proposal for a Commercial Orbital Transportation System: Preliminary Specifications for an Innovative Approach 08 May, 2009 Team Hippaforalkus Team Leader: Lake Singh lake.singh@gmail.com Team Members: Steven Box Michael Creaven Kris Curtis Ryan Hubbard Andrew Lyford Philip Maloney Chris Smith Advisor: Dr. Kevin Shinpaugh Copyright 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

2 Table of Contents Table of Contents... i List of Tables... iv List of Figures... v List of Acronyms, Symbols, and Units... vi Preface... viii 1 Introduction Problem Definition Document Organization System Architecture Objective Hierarchy Architecture Options and Selection Program Timeline Launch Vehicle System Launch Vehicle Configuration Trade Study Structural Design Propulsion Specifications Launch Vehicle Engine Trade Studies Launch Vehicle Attitude Determination and Control Sensors Actuators Cryogenic Thermal Insulation Communications, Command and Data Handling Communications Command and Data Handling The Spacecraft Structural Design Spacecraft Structure Power, Thermal, Environment Thermal Control Power: Budgets, Sources and Storage Environment: Impact and Radiation Shielding i

3 4.3 Environmental Control and Life Support System (ECLSS) Propulsion Specifications Attitude Determination and Control System Sensors Actuators Control Thrusters ISS Docking Procedure Sensors for Rendezvous Docking Strategy Communications, Command and Data Handling Communication Command, Data Handling, and Computer Systems Redundancy and Distributed Systems Performed Functions Error Detection Launch Abort, Manual Override, and Communications Failure Protocol Entry, Descent, Landing Thermal Protection Systems Reentry Maneuver Testing and Validation ISS Certification Pre-launch Requirements and Cargo Specifications Requirements During Operation Human Rating Ground Support Infrastructure Manufacturing Launch Operation Recovery and Refurbishment Architecture Analyses and Review System Analysis Cost Analysis ii

4 8.1.2 Risk Analysis Mission Analysis Proposed Flight Sequence Launch Manifest Crew Requirements Additional Vehicle Uses Conclusion References iii

5 List of Tables Table 1.1: NASA COTS Capabilities... 2 Table 2.1: Proposed Designs... 5 Table 2.2: Design Evaluation Matrix... 5 Table 3.1: Launch Vehicle Trade Study Summary Table 3.2: Dimensions and Properties of First and Second Stage Fuel Tanks Table 3.3: Properties of All Materials Considered for Launch Vehicle Structure.13 Table 3.4: Dry and Fueled Mass of Each Launch Vehicle Component Table 3.5: Engine Comparison Table 4.1: Mass Budget Table 4.2: Power Budget Overview Table 4.3: Waste Categories Table 4.4: ISS Nominal Atmosphere State Table 4.5: ADCS Overview Table 8.1: Cost Evaluation Table for the Argo Table 8.2: Launch Vehicle Economy Comparison Table 8.3: Cost Evaluation Table for the Manned Lotus iv

6 List of Figures Figure 2.1: Objective Hierarchy... 4 Figure 2.2: Gantt Chart Figure 3.1 Perspective View Figure 3.2 Launch Abort Figure 3.3 Exploded View Figure 3.4 Dimensional View Figure 3.5 Primary Engines Figure 3.6: Hypercone Figure 3.7: Interaction Chart Figure 3.8: Information Flow for the Launch Vehicle Figure 4.1: Generational Development Figure 4.2: Perspective View Figure 4.3: Basic Operations/Maneuvering Figure 4.4: Dimensional View Figure 4.5: System Organization and Integration Figure 4.6: Deployment Structure Figure 4.7: Shell Structure Figure 4.8: Crew and Cargo module Figure 4.9: Launch Abort System Figure 4.10: Cargo Module Figure 4.11: Flight Sequence Figure 4.12: Unmanned Power Distribution Figure 4.13: Manned Power Distribution Figure 4.14: Three-Axis Stabilization Figure 4.15: Final Approach Figure 4.16: The Decision Tree Figure 4.17: The Information Flow for the Spacecraft v

7 List of Acronyms, Symbols, and Units Acronyms ADCS Attitude Determination Control System Al Aluminum APCP Aluminum Perchlorate Composite Propellant ATK Alliant Techsystems C&DH Command and Data Handling CC&DH Communications, Command and Data Handling CCM Crew and Cargo Module CEU Camera Electronics Unit CEV Crew Exploration Vehicle CM Crew Module CO 2 Carbon Dioxide COTS Commercial Orbital Transportation System DC Direct Current ECLSS Environment Control and Life Support System EELV Evolved Expendable Launch Vehicle ETA Estimated Time of Arrival FSC Flexible Star Camera FTINU Fault Tolerant Inertial Navigation Unit GEM Graphite Epoxy Motor GN&C Guidance Navigation and Control GPS Global Positioning System ICCS International Space Station Commercial Cargo Service IMU Inertial Measurement Unit INU Inertial Navigation Unit ISS International Space Station LEO Low Earth Orbit LH 2 Liquid Hydrogen Li Lithium LiOH Lithium Hydroxide LOX Liquid Oxygen MMOD Micrometeoroids and Orbital Debris N 2 H 4 Hydrazine N 2 O 4 Dinitrogen Tetroxide NASA National Aeronautics and Space Administration NGSP Northrop Grumman Space Technology OH Operating Heads PICA Phenolic Impregnated Carbon Ablator RAM Random Access Memory ROM Read Only Memory RLVGS Redundant Launch Vehicle Guidance System SAAs Space Act Agreements SIGI Space Integrated GPS/INS SIRCA Silicone Impregnated Reusable Ceramic Ablator SpaceX Space Exploration Technologies SRB Solid Rocket Booster vi

8 SSME STGT TDRS TDRSS Ti VSD VSE WDEB Space Shuttle Main Engine Second Tracking and Data Relay Satellite Ground Terminal Tracking and Data Relay Satellite Tracking and Data Relay Satellite System Titanium Value System Design Vision for Space Exploration Wheel Drive Electronics Box Symbols and Units $ Dollar Degrees C Celsius C d Coefficient of Drag cm centimeter g gram GHz Gigahertz hr Hour I sp Specific Impulse kb Kilobyte kg Kilogram km Kilometer kn Kilonewton kpa Kilopascal krad Kilorad kw Kilowatt m Meter mm Millimeter MN Mega Newton MPa Mega Pascal mph Miles Per Hour N Newton Nm Newton-meter p Pressure r Radius σ Yield Strength s Second t Thickness USD United States Dollar V Voltage Vdc Direct Current Voltage W Watt yr Year vii

9 Preface This proposal is a response to the Request for Proposal issued by the AIAA for a reliable and cost effective commercial orbital transportation system to ferry crew and cargo to and from the International Space Station. Detailed in this document are the preliminary design and specification of components for a new commercial launch vehicle, spacecraft, and ground support infrastructure. The proposal has been ordered by mission phase and further by functional division to provide an organized description of the proposed system. In addition to technical design, this proposal considers relevant program drivers such as cost, systems engineering, risk analysis, redundancy, reliability, and marketability. Preliminary CAD drawings of all major systems are included, explained, and dimensioned. Pertinent technical specifications related to mission critical components such as power availability and cargo environment are detailed, as well as a sample flight manifest as part of the mission documentation package. Hippaforalkus was the name of the Lantean warship found by the members of the Atlantis expedition which was later renamed to Orion in the television series Stargate Atlantis. Loosely translated from Greek, it means He who has the strength of a horse. Team Hippaforalkus is comprised of 8 individuals, who are shown below. From Left: Ryan Hubbard, Steven Box, Philip Maloney, Chris Smith, Andrew Lyford, Kris Curtis, Lake Singh, Michael Creaven viii

10 1 Introduction Despite being a pinnacle of aerospace technology, the final flight of the veteran Space Shuttle will be in September The current planned operational date of the Orion Crew Exploration Vehicle (CEV) is September To help fill this gap in American manned spaceflight capability and to reduce the eventual strain on the Orion program from servicing the International Space Station (ISS), the National Aeronautics and Space Administration (NASA) has called for the design and development of a Commercial Orbital Transportation System (COTS). NASA has already awarded two Space Act Agreements (SAAs) in support of this initiative to Space Exploration Technologies (SpaceX) and Orbital Sciences Corporation. However, the award of these initial SAAs signals the emergence of a commercial manned spaceflight market. This proposal is a response to a request for a third COTS system to be fielded by Problem Definition There is an emerging market for COTS systems. This is a result of the world s space agencies collectively turning their sights to the moon at a time when they are just completing the hugely expensive ISS in Low Earth Orbit (LEO). The United States Vision for Space Exploration (VSE) calls for the development and use of commercial services to meet the maintenance needs of the ISS while NASA concentrates on the Constellation program. Additionally, there exists a largely untapped tourism and private venture market which is only beginning to open. A viable COTS system must be able to operate by 2012 and be capable of satisfying each of the four COTS Capabilities as defined by NASA, detailed in Table

11 Table 1.1: NASA COTS Capabilities. Required cargo loads that must be met by the program annually in no more than 8 flights. Mission Capabilities Capability A External (un-pressurized) cargo delivery to ISS & trash disposal (13,000kg/yr up-cargo) Capability B Internal (pressurized) cargo delivery to ISS & trash disposal (11,100kg/yr up-cargo) Capability C Internal (pressurized) cargo delivery to ISS & return to earth (9700kg/yr up-cargo, 4000kg/yr down-cargo) Capability D Crew delivery to ISS & return to earth (1050 kg net, 2 flights) The program must be based around a new cost-effective and reliable launch vehicle and a single spacecraft. The program should also be able to satisfy all requirements in a year within the course of two to eight launches. The spacecraft should be capable of docking with the ISS in its current operational orbit within 180 days of a request for service. Access to cargo aboard the spacecraft should be available as soon as 6 hours before launch. These and several other requirements compose the full problem definition as stated in the Request for Proposal. Further requirements will be discussed throughout the proposal as they are met. 1.2 Document Organization This proposal has been segmented into eight main chapters. The first two chapters introduce the purpose of the design and the overall systems engineering concepts used to conceive it. The next chapter details the design and function of the launch vehicle in relation to each functional division. Chapter 4 details the layout and design of the spacecraft and its primary functions and innovations. Following these sections, Chapters 3 and 4 are organized by vehicle, then functional division, then subsystem design. Chapter 5 discusses the reentry protocols and shielding present on the vehicle. Chapter 6 discusses the requirements, validation, and testing portion of the program, with Chapter 7 discussing the ground support necessary for mission success. The proposal is concluded with Architecture Analyses and Review which assesses the design and makes assertions about its nature and the risks involved in it. 2

12 2 System Architecture System architecture defines the criteria necessary to develop a viable solution to the proposed COTS mission. Established within this architecture, is the evaluation process for conceived design proposals, a timeline to completion, and a defined hierarchy of system characteristics for which the proposed design s conception and construction can be founded. System architecture is further fundamental to the design process, and dictates the organizational structure in reaching a finalized launch system. The following sections describe this process in detail and outline the COTS system architecture. 2.1 Objective Hierarchy Based upon the established needs for a commercial launch system, an iterative design process is employed to encompass the needs, alterables, and constraints necessary to achieve mission success and optimum functionality. Conception of a space delivery vehicle begins with identifying the criteria that will govern its success as an implemented design, as well as its commercial applicability. The goal of designing an optimal launch system can be broken down into a set of smaller goals that focus on key design points. Once an appropriate objective hierarchy is determined, a system can be designed based on applying a weighted value of each objective and sub-objective. This Value System of Design (VSD) classifies what the proposed launch vehicle needs to achieve, as well as provides a measure for how well those mission requirements are accomplished. The objective hierarchy for the VSD of the conceived launch system is shown in Figure

13 Optimal Launch System Selection Maximize Performance Capability Maximize Safety Maximize Applicability Minimize Cost Maximize Lifetime Lifetime in Years Maximize Versatility # of mission types Maximize Abort Capability Mission Duration in % Minimize Launch Vehicle Failure Minimize Sources of Failure Minimize Spacecraft Failure Maximize Crew and Cargo Integrity Survivability in % Maximize Availability Availability in % Maximize Marketability Market in $ Minimize Manufacturing and Operating Cost Cost in $ Failure in % Failure in % Figure 2.1: Objective Hierarchy. The VSD for the conceived launch system Conception of the spacecraft and launch vehicle is also highly dependent upon the primary mission objectives. Mission architecture culminates several capabilities that the conceived launch system must be able to achieve; as previously discussed and illustrated within Table 1.1. Thus, the final proposed launch delivery system will comprise all of the mentioned capabilities within is operational functionality, as well as illustrate a high compatibility with the VSD established for the overall system. 2.2 Architecture Options and Selection From a structural perspective, multiple initial design considerations are generated to analyze how the design problem will be approached. Through a diverse number of design proposals it is intended that an optimal solution will present itself; and further, that the design will also prove to exceed the initial mission and design requirements. A cast of suggested designs is produced within Table

14 Hybrid Engine Launch Air Launch Platform Launch Standard Launch Table 2.1: Proposed Designs. A list of suggested launch systems. Proposed Designs A completely reusable launch vehicle (no staging) that incorporates a hybrid engine to facilitate atmospheric travel and escape; conceptually similar to a space plane. A launch vehicle that uses modern aviation technology to elevate the spacecraft to a certain altitude from which it can separate and proceed boosting into space; conceptually similar to Spaceship One. A launch system that derives most of its structure and system functionality via a single platform; conceptually similar to a space elevator. A launch system similar to modern standards and technology; conceptually founded on rocket technology. Each of proposed designs within Table 2.1 illustrates a type of launch system (which focuses primarily on the structural approach) that could potentially be applicable to the mission and satisfy all the mentioned requirements. The proposed launch systems are evaluated based upon the VSD and its established measurable criteria. This is evaluation process is comprised in Table 2.2. Launch System Table 2.2: Design Evaluation Matrix. Evaluation of the proposed launch systems. Design Evaluation Matrix Criteria Reliability Launch Versatility Launch Manufacturing Capability Total (0.2) Rate (0.05) (0.15) Cost (0.25) Cost (0.25) (0.1) (1) Hybrid Engine Air Platform Standard =Poor 2=Moderate 3=Average 4=Good 5=Excellent Given the strong emphasis on minimizing cost and the timeframe for development and implementation, a single design concept presented itself as the optimal choice; that is one based upon modern launch systems and available technologies. Therefore, the proposed structural design incorporates modern launch system approaches, while optimizing design factors to achieve a cost effective and commercially applicable vehicle. All subsequent subsystems and inclusive technologies will go through their own evaluation process to eventually culminate and produce a viable transportation system. 5

15 One of the first steps in defining the architecture of the launch vehicle was determining the number of stages to be used. This was determined based on an analysis of similar scale modern launch vehicles and a determination of the advantages and disadvantages of additional stages. It was determined from the analysis of other launch vehicles that no modern system employs more than three stages; consequently only single, double, and triple stage architectures were analyzed. As stages are added, benefits are seen in increases of specific impulse at each separation point because a different nozzle can be used which is optimized for the separation altitude. However, due to the inherent large mass of engines, the addition of stages decreases the overall mass/fuel ratio of the launch vehicle. The loss of efficiency at high altitudes from carrying empty tankage and inefficient nozzle design made a single stage to orbit cost ineffective. Similarly, the added mass of a third engine to the launch vehicle as well as the added complexity made the three-stage-to-orbit method inferior to a two-stage-to-orbit approach. The two stage approach provides an ideal combination of simplicity and efficiency given technology which can be made ready by Program Timeline By following a design schedule, the COTS project will be completed in a timely and efficient way. The Gantt chart in Figure 2.2 shows the timeline for completion of the COTS project. Following the Gantt chart ensures that each system will be integrated into the overall vehicle in an efficient and logical manner. In this program, the Gantt chart sets a rapid pace for design integration, fabrication, and physical integration of the prototype in order to meet the operational date requirement set forth in the RFP. 6

16 Figure 2.2: Gantt Chart. Timeline for project completion. The first major milestone is a preliminary design review in May The program review has been completed and culminated in this proposal. Critical design reviews for the launch vehicle and spacecraft, the first unmanned test flight, and first manned flights are other important milestones for the proposed program. Critical design reviews will present a design which has undergone a detailed analysis by each functional division. The analysis will ensure that failures at each phase of flight are known and mitigated. 3 Launch Vehicle System A primary facet of the creation of a viable commercial manned spaceflight architecture is the design and operation of a cost effective and reliable launch vehicle. Towards that end, Hippaforalkus has laid out a preliminary design for the Argo Launch System. The Argo is a two and a half stage Evolved Expendable Launch Vehicle (EELV) class rocket employing the use of modern materials and commercialoff-the-shelf components. Preliminary structure, components, and performance values have all been 7

17 specified based on idealized analysis and assumptions based on the specifications of other currently operational launch vehicles. 3.1 Launch Vehicle Configuration Trade Study An essential part of the preliminary design was the comparison of several launch vehicle options. The first of these was the configuration of the launch vehicle and whether it would be a multi engine first stage or a single engine first stage. The heavy lift option consisted of 3 first stage main engines, coupled with a single second stage engine to produce enough thrust to lift approximately 23,000 kg of payload with the additional weight of the Lotus itself. The standard option consisted of a single first stage engine with 4 solid boosters and a single second stage engine to produce the thrust to lift 13,000 kg of payload into orbit. The heavy lift version was advantageous in that it required fewer launches, maximized use of the cost effective re-useable first stage engines, and provided a platform for other orbits should the client request them. The disadvantages of the design were the inability to meet rapidly changing needs, and the high unit cost undermined the versatility of the program. The standard version of the launch vehicle offered greater flexibility in the type and amount of cargo transported at a time, and allowed for a more rapid response to changing mission needs. The downside is that it requires more launches to transport the same cargo and a greater percentage of the parts are non-reusable. The standard launch variant was selected as it offered greater versatility and was more in line with the spirit of a transportation system. The heavy lift variant would require a complete mission profile re-design and the promise offered by it was not deemed sufficient for further study. If this mission were to be continued past the preliminary design stage it could conceivably be offered a mission variant for select mission requirements. 8

18 Configuration Advantages Table 3.1: Launch vehicle trade study summary. Disadvantages SLV Versatility High Cost Per kg of Cargo Optimum for Manned Missions Minimal Ability for Larger Mission Requirements Rapid Response Larger Percentage of the Vehicle is Disposable HLV Low Cost Per kg of Cargo No Versatility Room for Cargo Increases Optimal for Cargo Only Larger Percentage of Vehicle is Reusable Slow Response Time 3.2 Structural Design The preliminary launch vehicle design is a two-stage vehicle employing the use of four solid rocket boosters. The solid rocket boosters that will be used will be Alliant Techsystems (ATK) GEM 60 s. The launch vehicle is similar in both size and performance to the Lockheed Martin Atlas V and the Boeing Delta IV launch vehicles. The first and second stages of the vehicle will implement monocoque shells to simplify preliminary structural analysis. Stringers will be added in subsequent designs to bolster the vehicles ability to resist bending and axial loads. This addition will decrease the necessary thickness of the skin because it will only be required to overcome shearing and torsional loads. A dimensioned diagram of the launch vehicle is provided in Figure 3.4. Each stage consists of a hollow monocoque cylinder with an engine mounted at the base. The second stage of the launch vehicle has a diameter of 5.9 m to accommodate the equally wide Lotus spacecraft. The launch vehicle first stage has a 3.8 m diameter. An inter stage adapter ring attaches the first and second stages together. The first stage is equipped with an RD-180 engine and therefore has both a cryogenic liquid oxygen tank unpressurized RP-1 kerosene tank. The second stage has an RL-10B-2 engine mounted at the base which requires one tank of pressurized liquid hydrogen and another of pressurized liquid oxygen. shows the specifications of the fuel tanks. 9

19 The Argo Figure 3.1: Perspective View A perspective view of the Argo launch vehicle with some of its primary features identified. Figure 3.3: Exploded View An exploded view of the launch vehicle with its primary components illustrated. Lotus Spacecraft Figure 3.2: Launch Abort An illustration of a launch abort scenario and its structural orientation. Launch Abort System used to eject the ascent compartment of the CCM Excelsior Stage Petals and engines serve as brakes RL-10B-2 Engine Interstage Expendable Portion during abort Core Stage Two Stage Rocket: to achieve the desired orbit, the Argo is designed as a two stage lift vehicle with solid rocket boosters to provide further lift assistance. Payload Versatility: Although the Argo is designed to lift the Lotus, it still maintains the capability to carry other payloads and spacecraft. Launch Site Versatility: Because of its common design and relative size, the Argo has the ability to be launched from several launch sites. Primary Airlock: The Argo has a primary airlock to facilitate easy crew and cargo access. Figure 3.4: Dimensional View The dimensions of the launch vehicle, as well as some of internal components illustrated. Liquid Oxygen Tank m RP-1 Tank 1 m m Figure 3.5: Primary Engines A view from the bottom of the launch vehicle where primary thrust originates. GEM-60 (4) m 3.8 m 5.9 m 16.2 m Liquid Hydrogen Tank Liquid Oxygen Tank 12 m RD-180 Engine 10

20 Tank Table 3.2: Dimensions and Properties of First and Second Stage Fuel Tanks Thickness (mm) Material Volume (m 3 ) Mass of Fuel (kg) Mass of Tank (kg) Total Mass (kg) 1 st stage LO x 2 Al st stage RP-1 2 Al nd stage LO x 2 Al nd stage LH 2 2 Al The liquid oxygen and hydrogen tanks will be constructed out of Aluminum 2090 to meet the cryogenic storage requirements. The RP-1 tank will be constructed with Aluminum 7075 as the RP-1 tank is unpressurized, and this material is less expensive than Al The following equation is used to calculate the thickness of each tank. It is derived from the magnitude of the hoop and longitudinal stresses in the tank. 5 where p is the pressure in the tank, r is the radius and σ is the yield strength of the material. When the thickness was solved for, the necessary value for each tank was found to be less than one millimeter. Due to manufacturing limitations and the addition of a margin of safety, the thickness was increased to 2 mm for all of the tanks. Figure 3.4 provides a diagram of the launch vehicle with the tanks visible. In order to reduce mission costs, the RD-180 engine in the first stage will be recovered for reuse using a revolutionary retrieval system. The engine will be jettisoned from the first stage cylinder and a device called a hypercone will deploy on the topside of the engine section providing protection from the aerodynamic loads experienced by the engine in its descent. This cone is a high drag device that will begin to reduce the velocity of the engine and is illustrated in Figure

21 Figure 3.6: Hypercone. Innovative device for high speed braking. When the engine speed is nearly subsonic, the cone will be released and a round ring slot parachute will be deployed to reduce the engine s speed to a minimal value. This chute will then be jettisoned and a parafoil chute with a drogue line will be deployed. This chute allows the engine to control itself through the air at slow speeds and has a built in GPS allowing the chute to autonomously guide itself towards a desired location. When the engine reaches the correct location, a helicopter with an underside grappling hook will secure the drogue line attached to the engine chute and will transport it back to a ship. The implementation of this method eliminates forces from ground impact and keeps the engine out of the corrosive salt water where its internal parts could be damaged. Preventing this damage will both reduce the cost to rebuild the engine and increase its longevity. Carbon composites and various types of aluminum were candidates to be the materials used to manufacture the components in the launch vehicle. While the composite materials have excellent tensile strength and fatigue properties, their relative costs are too high for this low cost mission where no structural parts will be reused. Therefore, various types of aluminum will be used to produce the launch vehicle. The cylinders, adapter pieces, and the unpressurized RP-1 tank will be composed of Aluminum This alloy contains 87.2% aluminum, 5.6% zinc, 2.5% magnesium and trace amounts of copper and chromium. It has a density of 2.8 g/cm 3 and when heat treated has a yield strength of 505 MPa making it an ideal low cost material for the mission. The liquid hydrogen and liquid oxygen tanks will be made from Aluminum 2090 because of their cryogenic storage requirement. This metal is a 12

22 dominant material in the cryogenic tank field because it retains its material properties at low temperatures. It has a tensile strength of 455 MPa and a density of 2.59 g/cm 3. shows the analysis on the materials that were considered for the structure of the launch vehicle. The deciding factors were the overall cost and overall weight as these translate to a higher launch cost. Material Table 3.3: Properties of All Materials Considered for Launch Vehicle Structure Cost per kg. (USD) Density (g/cm 3 ) Yield Strength (MPa) Relative Yield Strength Required Mass 10-3 (kg) Total Cost (Thousand USD) Al Al Ti-6Al-4V Carbon fiber high modulus The carbon fiber material is the lightest; however, it is triple the cost of the Al The titanium alloy Ti-6Al-4V is strong but is six times more expensive than Al The Al-5052 is the cheapest per kg, however due to its low tensile strength the required amount of material needed makes it twice as heavy as Al-7075 and more expensive. Table 3.4: Dry and Fueled Mass of each Launch Vehicle Component Component Dry Mass (kg) Fueled Mass (kg) Stage 2 Cylinder Stage 2 Liquid Hydrogen Stage 2 Liquid Oxygen RL-10B-2 Engine Stage 1 to 2 Adapter Stage 1 Cylinder Stage 1 Liquid Oxygen Stage 1 RP RD GEM Total Mass The core stage of the Argo launch vehicle is designed to provide a little over half of the necessary kinetic energy for the spacecraft to reach orbit. This stage makes use of veteran practices to 13

23 provide a low cost first stage that is part of getting the Lotus to orbit. The first stage also contains a revolutionary way to recover and re-use its engine and has an option for solid rocket booster additions. The core stage provides the basis for the mission and is the precursor to Lotus function. 3.3 Propulsion Specifications The RD-180 engine chosen for the first stage provides a sea level thrust of 4.15 MN at an I sp of 311 to 338 seconds. It is a dual chamber engine and is half of an original Russian design for a 4 chamber main engine of the Zenit class of rockets. Dual chambering creates an increased risk of failure propagation but is fairly minor when only 2 chambers are present. Failure propagation rates increase in an exponential form so that increasing the chambers by two is only a small increase over one chamber. The use of RP-1, its ability to be produced by an American company, its human rating, and its thrust profile are all factors in determining that it was the best engine for this program s first stage. In addition, 4 GEM-60 solid rocket motors, produced by ATK, will be attached to the first stage. These provide 850 kn of thrust at an I sp of 270 seconds. These engines will be used to augment the lift capacity of the Argo. The second stage is powered by a RL-10B-2 which is a newer variant of a veteran design that has been widely used by NASA. The current design provides a thrust of 110 kn at an I sp of 462 seconds. The RL-10B-2 is currently in use as the second stage engine of the Delta IV series of launch vehicles. The RL- 10B-2 is the second stage engine present on the Delta IV family of rockets Launch Vehicle Engine Trade Studies Each option discussed in the following trade studies has an ideal position on a launch vehicle. Each has tradeoffs in performance, cost and safety which must be understood before a recommendation can be made. The following is a brief background and analysis of the use of eight engines selected that exemplify the principles of the COTS design of reliability and cost effectiveness. 14

24 Table 3.5: Engine Comparison, chart dealing with the advantages and disadvantages of each engine proposed Engine Advantages Disadvantages Pratt & Whitney RS-68 Pratt & Whitney RD-180 Inexpensive Acquisition Cost Thrust Profile within Requirements Well Tested and Man Rated Simple Design Man Rated Heavily Tested(Atlas 5) Fuel is Low Cost Fuel Requires Less Support Structure Snecma Moteurs Vulcain High I sp High Performance Fuel Reasonable Reliability Solid Rocket Booster Inexpensive Easily Stored Reliable Easily Re-useable Merlin 1C Man Rated (will have to be to perform on Falcon 9) Fuel is Low Cost Fuel Requires Less Support Structure Re-useable Sea Recovery J-2 X High Performance Fuel Will be man Rated for Ares Program Original Tested Heavily on the Apollo Program Expensive Fuel Extensive Cooling System Required for LH 2 Dangerous if Failure Occurs High Containment Demand Expensive Engine Manufacturing Cost Dual Chamber Leads to Higher Failure Propagation Risk Lost Thrust Expensive Fuel High Containment Demand Extensive Cooling System Required for LH 2 Failure means Loss of the Vehicle Highly Explosive Low I sp Never Tested Not in-use, Completely New Design Reliability in Question Lower I sp due to Engine Design Low Individual Thrust Expensive Fuel New Design Readiness Level Unknown HM7-B High I sp Expensive Fuel Low Thrust High Containment Demand RL-10B-2 High I sp High Performance Fuel Legacy US Design Man Rated Produced in Country Well Tested Expensive Fuel Containment Requirements Option 1: RS-68 The RS-68 was developed as part of the Evolved Expendable Launch Vehicle (EELV) program to power the Delta IV family of lift vehicles. The RS-68 has been rated to be capable of generating 3.3 MN of thrust in vacuum with a specific impulse (I sp ) of 409 seconds and a thrust to weight ratio of 50. The RS-68 uses cryogenic LH 2 /LO x fuel. The primary advantage of the RS-68 over other engines of similar 15

25 class is its simple design. The RS-68 is best suited to first stage operation due to its high thrust performance. One or two engines of this type would be sufficient for a first stage. Option 2: RD-180 The RD-180 is a Russian design licensed to Pratt & Whitney for production in the United States. It is the primary engine used on Atlas III and V launch vehicles. The RD-180 produces 4.15 MN of vacuum thrust with a specific impulse of 338 seconds and has a thrust to weight ratio of 77. It uses a RP- 1/LO X fuel mixture. Its dual chamber, dual nozzle configuration makes it a suitable choice for a first stage only. Its thrust is the highest of any liquid option. The RD-180 is a veteran design with great reliability, despite an inferior specific impulse when compared to other liquid options. Option 3: Vulcain The Vulcain is the primary engine for the first stage of the Ariane 5 launch vehicle. It only produces 1.34 MN of thrust in a vacuum. However, the Vulcain makes up for low thrust by having a comparably high specific impulse of 431 seconds and a high thrust to weight ratio of 81. The Vulcain uses LH 2 /LO x fuel and has reasonable reliability. The Vulcain could be used as a second stage engine. Option 4: Solid Engines While liquid propulsion is the dominant form used in modern launch vehicles, there are still many applications and advantages to using a solid engine. While solid engines have inferior specific impulse and thrust to weight ratio, their fuel is inexpensive and easily stored, and their function is far more reliable. They have also been shown to be easily reused on several missions. One such example is the Space Shuttle SRB. Each booster provides 13.8 MN of sea-level thrust at a specific impulse of 269 seconds. The thrust to weight ratio of the booster is 15.5, but the Ammonium Perchlorate Composite Propellant (APCP) is far less costly than cryogenics and easier to store. The solid engine is applicable as only a first stage or addition to a first stage due to comparatively low I sp and inability to throttle. 16

26 Option 5: Merlin 1C The Merlin 1C is a rocket engine developed by SpaceX to be used on the Falcon 9 series of launch vehicles. The Merlin engine uses RP-1 and liquid oxygen as propellants in a gas-generator power cycle. It is designed to be re-usable and for sea recovery. This version uses a regeneratively cooled nozzle and combustion chamber and has been tested at over 27 minutes, which is about ten total flights. It is currently undergoing qualification but is stated to be ready for the 2009 Falcon 9 launch date which is before this program is set to begin. While this design is applicable at all stages of launch, it seems best suited to be augmented on the first stage or as a standalone on the second stage. 1 Option 6: Kestrel 2 The Kestrel Engine is a RP-1/LO x pressure-fed rocket engine. The Kestrel was developed by SpaceX for the upper stage of the Falcon 1. As these are meant for upper stage and space use a pyrophoric system gives them multiple restart capability. This allows for highly accurate altitude and orbital insertion. This system is an example of a small kerosene pintle engine. This thruster has very low maximum thrust making it useful on the upper stages of a launch vehicle or a spacecraft. Its multiple firings allow for its use in attitude and orbital control after launch; this could be integrated into the spacecraft design. 2 Option 7: J-2X The J-2 engine is the largest production liquid hydrogen rocket beside the Space Shuttle Main Engines (SSME). The J-2X is a new variant of the engine being designed to support the Constellation program and its replacement of the Space Shuttle. It uses a gas generator power cycle. This is an example of a heavy LH 2 /LO x engine and its performance. The high thrust and specific impulse of this rocket make it an excellent choice for launch of a spacecraft. The drawbacks include a low thrust to weight ratio meaning that this is a very heavy engine and less attractive for stages beyond the ground stage as more energy must be expended to launch it. It has superb safety and is one of the better liquid 17

27 options for ground launch. This engine should not be used beyond the first stage however, if cost savings are to be received. This fuel type is very expensive and can cause large cost to the program. 3 Option 8: RL-10B-2 The RL-10B-2 is an engine used as the second stage engine for the Delta IV family of rockets. It provides a thrust of kn. It has a specific impulse of 462 seconds using the LH 2 /LO x mix common in aerospace application. The engine has a weight of 301 kg and is produced by Pratt & Whitney. This engine is an excellent engine for the second stage of a launch system. The drawbacks are the use of a thermally demanding fuel which is also high cost. 4 The conclusion of these trade studies was the use of the RD-180 engine for the first stage and the use of the RL-10B-2 engine for the second stage of the launch vehicle. These were chosen due to their capability to be made in the United States and their adherence to the mission profile thrust requirements. Two stages were chosen with the design in mind to allow the partial recovery and re-use of the first stage and the expendability of the second stage. 3.4 Launch Vehicle Attitude Determination and Control The Attitude Determination Control System (ADCS) consists of a system of sensors and actuators. This system determines the position and orientation of the Argo and compares it to a known trajectory. The sensors will then create an error function, which is sent to the onboard flight computer to calculate the torques and forces required to reach a desired orientation. Finally, the flight computer will send commands to the actuators to execute the required torques. All sensors and actuators on the Argo are reliable products, which are commercially available. When choosing the components for the ADCS, it is important to know how the ADCS will interact with other subsystems. Figure 3.7 illustrates how the ADCS interacts with subsystems. 5 18

28 Figure 3.7: Interaction Chart. Outline of the subsystem interactions with ADCS Mission operations outline the sensor pointing and special maneuvers to be performed by the ADCS. The ADCS on the Argo must ensure that antenna are oriented in a way which communications are effective. Guidance Navigation and Control (GN&C) subsystems determine the desired position. The power subsystem determines the amount of power that can be allocated to sensors and actuators. The ADCS on the Argo determines the following elements which have an impact on the structure of the launch vehicle: sensor placement, actuator placement, center of mass, and moments of inertia. Since the ADCS works together with other subsystems, it is important to choose commercial components which are veterans of past launch vehicles Sensors Sensors determine the launch vehicle s position and orientation during its launch by performing static and dynamic measurements. The BF Goodrich Data Acquisition System stores the measurements and relays them to the onboard flight computer. 6 Static and dynamic measurements are determined by Honeywell s Space Integrated GPS/INS (SIGI). The SIGI is an integrated global positioning/inertial 19

29 navigation system. Honeywell s SIGI is qualified for launch vehicle environments. The SIGI provides the Argo with the following options to navigate: pure inertial, GPS-only, and blended GPS/INS. Since Honeywell s SIGI is used on the space shuttle and the ISS, it is a reliable component. 7 To ensure double fault tolerance on unmanned missions, a Honeywell Inertial Navigation Unit (INU) will be utilized as well as Honeywell s SIGI. The INU is a common navigation unit currently used in several modern launch vehicles. Honeywell s INU is a self-contained package capable of inertial measurements and flight control. Both the INU and the Inertial Measurement Unit are a three-axis strap-down system containing accelerometers and ring laser gyros. The ring laser gyros allow for high accuracy in inertial measurements when compared to a mechanical gyro, because laser gyros have a smaller chance of mechanical failure. 8 In order to provide a triple fault tolerance for manned missions, Honeywell s Redundant Launch Vehicle Guidance System (RLVGS) will be used. The RLVGS is a next-generation, high accuracy launch vehicle navigation system. Since the RLVGS is capable of performing inertial measurements, there will be a triple fault tolerance in inertial measurements. The avionics on the RLVGS will serve as a redundant system to the avionics in the SIGI. Honeywell s RLVGS also has a robust fault tolerance in all functional areas. 9 Therefore, it is clear that when there is a manned mission, the Argo has an ADCS system with a low chance for catastrophic failure Actuators Once the flight computers have determined the necessary torque to apply to the Argo, burn commands are sent to actuators. The actuators will consist of four 27 N rotational hydrazine thrusters and eight 40 N lateral hydrazine thrusters. The combination of lateral and rotational thrusters provides the Argo with three-axis stability. 6 20

30 3.5 Cryogenic Thermal Insulation The launch vehicle cryogenic fuel tanks must be maintained at very low temperatures in order to minimize boil off of fuel. Insulation is necessary to help maintain these temperatures. A combination of spray on foam insulation and phenolic thermal insulators will be used to perform this function. This insulation scheme is similar to that used by the External Tank on the Space Shuttle and the tanks on the Delta IV family of expendable launch vehicles. 10 A projected mass contribution of 420 kg is expected for the first stage and 130 kg for the second stage for a combined total of 550 kg. 3.6 Communications, Command and Data Handling Communication, command and data handling is the figurative nerve center of any space vehicle. It is responsible for interfacing with other systems in the generalized system architecture and for automated control Communications The antenna package onboard the Argo will consist of three antennas placed azimuthally along the fuselage of the second stage. The antennas will be placed 120 apart from each other, which will allow for 360 coverage Command and Data Handling Since the first stage will be jettisoned shortly after launch, the main computer components for the launch vehicle will be placed in the second stage. There will be a specific type of avionics system for the launch vehicle and all other subsystems onboard will be dealt with via a RAD750 distributed system. The avionics system aboard the launch vehicle will be a Honeywell Fault Tolerant Inertial Navigation Unit (FTINU). The FTINU will use the avionics equipment onboard the launch vehicle, such as the ring laser gyros and the space integrated GPS, and will relay to the ground station the position and orientation of the launch vehicle during flight

31 To ensure that all systems are functioning properly, sensors for temperature, pressure, and stress will be located in critical areas on the launch vehicle and will send a constant stream of telemetry to the ground station. This system will allow the ground station to monitor the status of the Argo during flight. Systems will be in place to control the jettison of the boosters near the end of their burn indicated by a timer in the Argo flight computer. Once the exact time has been reached, the main computer onboard the launch vehicle will send a high level discrete pulse to the exploding bolts connecting the booster to the main phase. The exploding bolts will then detonate in such a way to cause the boosters to peel off from the launch vehicle. Sensors will indicate when the first stage is out of fuel. This condition will then trigger the separation sequence, in which a set of exploding bolts located between the first stage and inter stage will be activated. To help distance the two stages even further, retro rockets will be ignited via a high analog signal. A similar mechanisms is used to remove the inter stage. After the first stage is sufficiently distanced from the second stage, the onboard computer system will activate the igniter on the second stage s engine to propel the spacecraft to its desired orbit. Once the second stage is depleted of fuel and the spacecraft has been placed in its proper orbit, the second stage will be jettisoned and the computer system onboard the spacecraft will be in charge of further activities. 22

32 Figure 3.8: Information Flow for the Launch Vehicle. This shows the inflow of commands to the avionics system and main computer and the out flow of data collected by the onboard data handling system. 4 The Spacecraft After detailed functional analysis and system synthesis, a proprietary structural design was chosen for the spacecraft based upon modern rocket launch systems. The conceptual design supporting this vehicle is intended to be integrated within the launch vehicle and mimic modern systems such as the Orion or Apollo capsules. Innovation was applied in the optimization of the overall reusability of the 23

33 spacecraft (and launch system), which further supported the commercial aspect of the mission requirements. 4.1 Structural Design Preliminary considerations for the design of the spacecraft centralized around a cylindrical housing that incorporated all necessary subsystems (similar to Orion capsule). However, after further consideration it was decided that reusability can be achieved by integrating the payload housing into the spacecraft structure. This gave conception to the Lotus spacecraft. The Lotus derives some of its primary functionality through utilizing the fairing as structural members for both attitude control and subsystem stem integration. The Lotus spacecraft officially went through three generations of design where the structure was optimized and desired functionality was achieved. This generational development is illustrated in Figure 4.1. Generation 1 Generation 2 Generation 3 Figure 4.1: Generational Development. Three generations of the Lotus design From Figure 4.1 it is evident that the lotus spacecraft will serve as the nose cone structure of our launch vehicle. Through three generations of design the spacecraft was able to incorporate aspects of modularity, launch abort capability, easy crew and cargo access, and overall optimization of size and weight. The proposed design is depicted in the Lotus foldout. 24

34 1.25 m 5.85 m Housings (4) (4): Internal structure used for module access or subsystem integration. Figure 4.4: Dimensional View The dimensions of the spacecraft R=25.4 m Airlock: Primary access to the spacecraft at any point during operation. The Lotus The Hippaforalkus Program Module: The manned module is shown here; other modules may be utilized. 10 m 0.75 m 1.5 m Primary Engine (4): (4) The main boost engine used for orbit changes, maneuvering, and reentry. 3m 2m 2.5 m Airlock: An access hatch used for docking with the ISS and entering/exiting the vehicle during space operation. Solar Panels (8): the primary source of power for the spacecraft; culminates 24 square meters of solar panel surface area. Petals (4): A quarter section of the included fairing, used for subsystem integration (i.e. solar panels, thrusters, etc.). Figure 4.2: Perspective View A perspective view of the Lotus spacecraft with some of its primary components illustrated Reentry Boost Configuration Figure 4.3: Basic Operations/Maneuvering An illustration of the basic maneuvers and operations that the spacecraft will perform during flight. The primary direction of travel The direction of primary thrust Turning direction capable with vector thrust Orbital Configuraton Figure 4.5 5: System Organization and Integration Green:: Crew and Cargo Housing/Structure Blue:: CC&DH, Power, Thermal Control, Storage Systems Red:: Primary thrust devices and location of structural integration Olive Green: Green Entry, landing, and descent systems location Pink:: Communication system integration Yellow:: Structural location for other systems; sy such as attitude thrusters and sensors Splash-down Petals Deploy 25