Solar Orbit Transfer Vehicle June 1, 2001

Christopher D. Hall, Associate Professor Department of Aerospace and Ocean Engineering Virginia Polytechnic Institute and State University Blacksburg, VA 24061

(540) 231-2314 (540) 231-9632 (FAX) [email protected]

Heat Launch to LEO Mechanism graphite 300 km altitude Deployment block to Wait for 28 ° inclination Inflate 2400 K command collector 1 ½ hours 2 hours 11 ½ hours Continuous

01 234561718 19 Approximate 1 hour Continuous Continuous mission time Establish Perform Check telemetry in hours communication station-keeping data link operations as Ensure proper necessary subsystem Balance operation disturbance torques

SOLAR ORBIT TRANSFER VEHICLE

Submitted By: Nicholas Cummings Kevin Earle Douglas Klingemann Michael Nakles Andrew Pollard Ross Stilling Clinton Stone Eric Van Veldhuizen Richard Winski Emily Woodward

Aerospace and Ocean Engineering Department Virginia Polytechnic Institute and State University Blacksburg, Virginia 3 May 2001

Professor: Dr. Christopher Hall Table of Contents Symbols...... v Abbreviations...... ix List of Figures...... xi List of Tables ...... xiii Chapter 1: Introduction ...... 1 1.1 Request For Proposal ...... 1 1.2 Background...... 2 Chapter 2: Solar Orbit Transfer Vehicle ...... 4 2.1 Overview...... 4 2.2 Modeling...... 5 2.3 Performance ...... 13 2.4 Operations...... 15 2.5 Summary...... 16 Chapter 3: Solar Collector...... 18 3.1 Overview...... 18 3.2 Collector Design Considerations...... 18 3.3 Solar Collector Subsystem Description ...... 19 3.4 Mathematical Modeling ...... 21 3.5 Deployment, Inflation, and Pressurization...... 23 3.6 Summary and Further Research...... 24 Chapter 4: Solar Thermal Engine and Propulsion...... 25 4.1 Introduction...... 25 4.2 Overview...... 26 4.3 Modeling...... 27 4.3.1 Heat Transfer...... 27 4.3.2 Orbital Motion...... 29 4.3.3 Alternative Transfers...... 30 4.3.4 Propellant Feed System...... 30 4.4 Performance ...... 31 4.5 Summary...... 31 4.6 Next Steps ...... 31 Chapter 5: Thermal ...... 33 5.1 Thermal system design...... 33

ii 5.2 Fuel tank...... 33 5.3 Receiver/Cavity Design ...... 36 5.4 Insulation...... 37 5.4.1 Engine performance ...... 37 5.5 Conclusions...... 40 Chapter 6: Fuel Tank/Primary Structure...... 41 6.1 Overview...... 41 6.2 Hardware...... 41 6.3 Modeling...... 41 6.3.1 Monocoque Design ...... 44 6.3.2 Applied and Equivalent Axial Loads ...... 45 6.3.3 Sizing for Tensile Strength...... 46 6.3.4 Sizing for Stability (Compressive Strength) ...... 46 6.3.5 Internal Pressure...... 47 6.3.6 Mass and Volume of Monocoque Tank...... 47 6.3.7 Skin-Stringer Design...... 48 Chapter 7: Attitude Determination and Control...... 51 7.1 Overview...... 51 7.1.1 Background ...... 51 7.1.2 Requirements ...... 54 7.2 Hardware Examples ...... 55 7.2.1 Attitude Control ...... 55 7.2.2 Attitude Determination: ...... 57 7.3 Modeling and Performance...... 58 7.3.1 Modeling ...... 58 7.3.2 Attitude Determination...... 71 7.4 Summary...... 73 7.4.1 Attitude Control ...... 73 7.4.2 Attitude Determination...... 73 Chapter 8: Secondary Structure ...... 74 Chapter 9: Grappling Arms and Mechanical Hand...... 77 9.1 Overview...... 77 9.2 Hardware...... 77 9.3 Results and Summary...... 80 Chapter 10: Communications...... 81

iii 10.1 Telemetry, Tracking and Command Communications ...... 82 10.2 Docking Procedure Communications ...... 83 10.3 Conclusions...... 83 Chapter 11: Command & Data Handling...... 84 11.1 Software Requirements...... 85 11.2 Docking Procedure Requirements...... 86 11.3 Conclusions...... 87 Chapter 12: Power system design ...... 88 12.1 Power requirements...... 88 12.2 Power Conversion...... 89 12.2.1 Thermionic power system...... 89 12.2.2 On orbit power generation ...... 90 12.2.3 Cost and performance considerations ...... 91 12.2.4 Thin film solar array...... 92 12.2.5 Secondary battery system...... 93 12.3 Conclusions...... 93 Chapter 13: Launch Vehicle ...... 94 13.1 Overview...... 94 13.2 Hardware...... 94 13.3 Performance ...... 103 Chapter 14: Conclusions ...... 107 References...... 109 Appendix A : Interplanetary Mission Considerations...... 113 Appendix B : Communications and Data Handling Hardware...... 115 Appendix C : Request For Proposal...... 117 Appendix D : Solar Orbit Transfer Vehicle Design Summary ...... 120

iv Symbols

A Cross-sectional area of the beam

Aarray Area of a primary solar collector

ATOTAL Total collector surface area b Stiffener spacing Cp Specific heat

CpC Specific heat of the graphite block

CpH Specific heat of the hydrogen propellant E Young’s modulus of elasticity e Emmisivity

ESUN Solar radiation energy per square meter F Force

fnat Natural frequency

Ftu Ultimate tensile strength

Fty Yield tensile strength g Gravitational acceleration at Earth’s surface

Hfg Latent heat

I Area momentum of inertia of the beam’s cross-section

ISP Specific impulse

Isp,max Maximum obtainable specific impulse

Ixx Mass moment of inertia about the x-axis

Iyy Mass moment of inertia about the y-axis

Izz Mass moment of inertia about the z-axis k Buckling coefficient

Keff Effective conductivity

kS Spring constant L Beam length

Ld Lifetime degradation m Mass m& Mass flow rate

v m& Flow rate of Hydrogen H2 & Flow rate of propellant m p M Bending moment limit load

mB Mass of beam (uniformly distributed) mblock Mass of graphite block

me Mass of empty tank me Mass of empty tank

ML Payload mass

Mo Mass of vehicle before burn

MP Propellant mass MS Margin of Safety

MSOTV Total fueled SOTV mass mTES Mass of Thermal Energy Storage

mtot Mass of full tank

Mw Molecular weight n Load factor nS Number of stiffeners p Focal length of a parabola

Paxial Axial load

Pcr Critical buckling load

Pd Power in daylight

Pe Power in eclipse

Peq Equivalent axial load

Ps Power of solar array Q Heat transfer

 Q&  Rate of heat transfer from the Sun    A  sun r Radius r& Normal velocity r&& Normal acceleration

vi R Orbital radius of spacecraft

rs Stiffener radius T Temperature t Time

T1 Temperature at position 1

T2 Temperature at position 2

Ta Aerodynamic disturbance torque

Ta, worst Worst possible aerodynamic disturbance torque tb Burn time

tc Coast/Reheat time

Tc Cold temperature

Td Time in daylight

Te Time in eclipse

Tg Gravity-gradient disturbance torque

Tg, worst Worst possible gravity-gradient disturbance torque

Th Hot temperature

Tm Magnetic disturbance torque

Tm, worst Worst possible magnetic disturbance torque

Tmax Maximum block temperature

Tmin Minimum block temperature

Tsr Solar radiation disturbance torque

Tsr, worst Worst possible solar radiation disturbance torque Y Change in velocity x Roll ∆x Distance between positions 1 and 2

Xd Efficiency in daylight

Xe Efficiency in eclipse y Pitch z Yaw

δ Deflection

vii γ Reduction factor Total collector reflective efficiency Gravitational constant Poisson’s ratio Angular position & θ Angular velocity && θ Angular acceleration

dev Maximum deviation of the Z-axis from local vertical in radians r Angle of rotation Density ρ Density of liquid hydrogen H 2 ρ Density of aluminum Al σ Elastic buckling stress cr σ Hoop stress h

viii Abbreviations

ACS Attitude Control System ADCS Attitude Determination and Control System ADS Attitude Determination System Al-Li Aluminum-Lithium BECO Booster Engine Cutoff BOL Beginning of life C&DH Command & Data Handling CAD Computer Aided Design CCAS Cape Canaveral Air Station CCB Common Core Booster cg Center of Gravity CIII Centaur Upper Stage CMG Control Moment Gyroscope DEC Dual-Engine Centaur DOD Depth of discharge DSN Deep Space Network EELV Evolved Expendable Launch Vehicle EEPROM Electrically Erasable Programmable Read-Only Memory EOL End of life GEO Geostationary Earth Orbit I/O Input/Output IAE Inflatable Antenna Experiment IAF International Astronautical Federation IN-STEP In-Space Technology Experiments Program KIPS Kilo-Instructions Per Second LEO Low Earth Orbit LV Launch Vehicle LVA Launch Vehicle Adapter Mbps Mega Bits Per Second MECO Main Engine Cutoff MLI Multi-Layer Insulation NASA National Aeronautics and Space Administration

ix PCI Peripheral Component Interconnect PLF Payload Fairing RAX Receiver Absorber Exchanger RFP Request For Proposal RHPPC Radiation-Hardened PowerPC SADA Solar Array Drive Assembly SEC Single-Engine Centaur SIP Standard Interface Plane SOTV Solar Orbit Transfer Vehicle SRAM Synchronous Random Access Memory SRB Solid Rocket Booster SUROM Start-Up Read-Only Memory TDRSS Tracking and Data Relay Satellite System TES Thermal Energy Storage TT&C Telemetry, Tracking, & Command UV Ultraviolet VCS Vapor Cooling Shield

x List of Figures

Figure 2.1: A sample LEO- GEO transfer trajectory ...... 7 Figure 2.2: Propellant usage for a LEO-GEO transfer with no plane change as a function of graphite block size ...... 8 Figure 2.3: Transfer time for a LEO-GEO transfer with no plane change as a function of graphite block size ...... 9 Figure 2.4: Number of thrusting cycles required to complete a LEO – GEO transfer decreases as block mass increases...... 10 Figure 2.5: Transfer time vs. solar collector area for a LEO-GEO transfer...... 11 Figure 2.6: Propellant usage vs. solar collector area for a LEO-GEO transfer...... 12 Figure 2.7: Propellant usage vs. total system mass (including 3000 kg client) for a LEO-GEO transfer...12 Figure 2.8: Final mass vs. total system mass (including 3000 kg client) for a LEO-GEO transfer...... 13 Figure 2.9: LEO – GEO Transfer times for a range of client masses ...... 14 Figure 2.10: LEO – GEO propellant requirements for client masses ranging from 1000 to 4500 kg ...... 14 Figure 2.11: Deployment phase mission timeline...... 17 Figure 3.1: Moog Type 2 Solar Array Drive Assembly0...... 20 Figure 4.1: SOTV illustrating location of Receiver/Absorber/Exchanger and Primary solar collectors ...25 Figure 4.2: Theoretical solar thermal propulsion engine ...... 26 Figure 4.3: Comparison of energy balance model with experimental results...... 29 Figure 5.1: Vapor cooled shield...... 34 Figure 5.2: Effective conductivity vs. MLI thickness...... 35 Figure 5.3: Heat flux vs. mass flow rate ...... 36 Figure 5.4: Burn time to 2000 K...... 38 Figure 5.5: Heat time vs. block mass...... 39 Figure 5.6 Reheating times as a function of collector size...... 40 Figure 6.1. Laterally and Axially Loaded Cantilevered Beam0 ...... 42 Figure 6.2: SOTV fuel tank modeled as a cantilevered beam...... 42 Figure 6.3: Z-Shaped Stiffener (2 Square cm Cross-Sectional Area) ...... 50 Figure 6.4: Cross-Section View of Stiffener Configuration ...... 50 Figure 7.1: Spacecraft Body Reference Frame ...... 52 Figure 7.2: Spacecraft Configuration used for preliminary analysis...... 59 Figure 7.3: Thrust required vs. Slew rate...... 60 Figure 7.4: Mass of propellant vs. slew rate ...... 61 Figure 7.5: Possible slewing maneuvers vs. slew rate ...... 61

xi Figure 7.6: Constellation HR14 reaction wheel28 ...... 62 Figure 7.7: M1700 CMG528 ...... 65 Figure 7.8: Worst case disturbance torques vs. altitude...... 66 Figure 7.9: Mass of hydrogen gas propellant as a function of altitude in a LEO-GEO transfer...... 68 Figure 7.10: Change in mass center of the SOTV as a function of altitude in a LEO-GEO transfer...... 69 Figure 7.11: Positioning of control moment gyroscopes49 ...... 70 Figure 8.1: RAX support system...... 74 Figure 8.2: RAX support structure and solar array system...... 75 Figure 8.3: Spacecraft bus support structure...... 75 Figure 8.4: Bus support structure with components...... 76 Figure 8.5: Launch vehicle adaptor and topside components ...... 76 Figure 9.1: Magnum 7 robotic arm bending and rotation diagram ...... 78 Figure 9.2: Degrees of freedom ...... 79 Figure 12.1: System trade study results ...... 89 Figure 12.2: Receiver mass requirements ...... 91 Figure 12.3: Solar cells mounted on solar concentrators ...... 92 Figure 13.1: Atlas V 500 Launch System2 ...... 95 Figure 13.2: Atlas V Payload Fairing2 ...... 96 Figure 13.3: Atlas V Payload Envelopes2 ...... 97 Figure 13.4: Atlas V Standard Interface Plane (SIP)2 ...... 98 Figure 13.5: Allowable Spacecraft cg Location Above Sep. Plane 2 ...... 99 Figure 13.6: 2.97 m (117 in.) Diameter Truss2...... 100 Figure 13.7: Payload Fairing Using 2.97 m (117 in.) Diameter Truss2...... 100 Figure 13.8: Customized Launch Vehicle Adaptor...... 101 Figure 13.9: Stowed SOTV Configuration ...... 102 Figure 13.10: Stowed Bus Components in Launch Vehicle Adaptor ...... 103 Figure 13.11. Atlas V 552 LEO Capabilities 2 ...... 104 Figure 13.12: Typical Atlas V 552 LEO Ascent 3 ...... 105

xii List of Tables

Table 3.1: Moog Type 2 SADA Characteristics0 ...... 20 Table 5.1: Specific heat for parahydrogen ...... 35 Table 6.1: Fuel Tank Characteristics and Requirements...... 43 Table 6.2: Limit loads ...... 46 Table 6.3: Optimal skin-stringer tank configuration...... 49 Table 7.1: Attitude Control Systems49 ...... 55 Table 7.2: Examples of Cold-Gas Propulsion49...... 57 Table 7.3: Attitude Determination Systems49 ...... 58 Table 10.1: SOTV communications hardware overview...... 82 Table 11.1: SOTV itemized computer software requirements...... 84 Table 11.2: SOTV itemized ADCS computer software requirements...... 86 Table 12.1: Power requirements ...... 88 Table 12.2: Thermionic converter specifications...... 90 Table 13.1: Atlas V 500-series maximum load factors...... 106 Table A.1: Interplanetary mission probe mass transfer capability...... 114

xiii Chapter 1: Introduction

This report outlines the conceptual design of the Solar Orbit Transfer Vehicle (SOTV). Many of the performance issues related to the design of such a vehicle are explored, and the subsystems are designed to the level of detail such that the performance of the SOTV design can be predicted and analyzed. The report first gives an introduction to the SOTV concept by outlining the request for proposal on which the design is based. It also describes the conceptual design alternative chosen for the SOTV. Next, the report takes on a mission analysis perspective and describes the performance capabilities of the overall SOTV spacecraft and how these capabilities are influenced by the design tradeoffs between the individual subsystems. The report then focuses on the individual subsystems, discussing them in detail each in their own individual chapter. First, an overview of the subsystem is presented, including performance requirements and design considerations. Next, hardware alternatives are discussed, compared, and modeled. Finally, the performance of the subsystem is presented and the design is summarized. Some chapters also state recommendations for further study. Finally, the paper concludes by summarizing the capabilities of the SOTV and comparing them to the requirements set forth in the request for proposal. Several recommendations for further study are made that suggest how the SOTV design can be improved. Appendices at the end of the report give the overall SOTV design specifications organized by subsystems, an analysis of the SOTV’s ability to perform interplanetary missions, and other supplemental information.

1.1 Request For Proposal The design of the SOTV is based on the constraints and requirements outlined in a Request For Proposal (RFP) issued by the U.S. Air Force. The following section summarizes the main design guide- lines and constraints outlined in the RFP. The actual RFP is included in Appendix C. The main mission objective specified by the RFP for SOTV is to transfer a 3000 kg to 5000 kg client satellite from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO). In addition, the SOTV must be able to perform at least one of the following alternate missions: • Recovery and transfer of spacecraft that have been launched into incorrect orbits • Cleanup of nonfunctional satellites in GEO • Delivery of payloads to and from the International Space Station Additionally, a trade study of interplanetary probe transfers to an appropriate escape trajectory must be completed. The LEO – GEO transfer must take no longer than 90 days, including the time required for rendezvous and capture of the client satellite. However, a goal of the SOTV design is to complete this

1 transfer in 60 days or less. Hot propellant gas may not be used in the direction of the client satellite. Additionally, the RFP restricts client spacecraft loading to 0.1 g. The SOTV must operate semi-autonomously. For example, the rendezvous and capture phase of the missions requires pilot-in-the-loop architecture; however, the orbit transfer phase can be completed autonomously. An on-sun pointing accuracy of ± 1° is a system requirement for gross pointing and tracking. A fine pointing accuracy of ± 0.1° is required for the primary concentrator. SOTV must use a solar thermal engine as its primary propulsion system. The solar thermal engine must produce at least 67 N (15 lbs) of thrust, and the maximum allowable bulk average temperature of the solar receiver cavity is 2400K (± 25 K). A Boeing Delta-IV Medium (5,4 or smaller) or a Lockheed-Martin Atlas V (522 or smaller) must be used to launch the SOTV. These rockets are comparable in launch capabilities and reliability. The minimum lifetime requirement for the SOTV is five years, but a ten-year lifetime is prefer- able. The liquid hydrogen storage tank would need to be refilled at some time, especially if multiple LEO-GEO transfers are to be accomplished over the lifetime of the SOTV. The RFP specifies that such refueling operations are to be assumed feasible. Finally, SOTV must function at a reliability rate of 95% throughout the ten-year lifetime.

1.2 Background Two possible design alternatives satisfy the criteria set forth by the RFP. The first is a low cost alternative whereby the SOTV would launch with the client satellite and transfer the client from LEO to GEO similar to a conventional upper stage. The SOTV would remain with the client in GEO throughout the life of the client, supplying power, and eventually transferring the client to supersynchronous orbit. The second alternative offers a maximum performance approach to the RFP. This design alternative consists of a stand-alone vehicle equipped with a grappling arm. Such a design supports all of the missions proposed by the RFP, but would require refueling to complete more than one LEO to GEO transfer. The technology behind the solar thermal propulsion system has been tested terrestrially, but has yet to be applied to space applications. The first SOTV will therefore function as a technology demon- strator. Although both design alternatives have the potential to satisfy the RFP, the low cost alternative is not a sensible option from a programmatic standpoint. Customers would have a difficult time selecting an unproven SOTV as an upper stage when proven upper stages already exist and are quite reliable. The recovery of a misplaced satellite is a much better candidate for a first SOTV mission. Nu- merous vehicles have been placed in incorrect orbits by malfunctioning launch vehicles and guidance systems. Many of the misplaced vehicles would be completely functional if placed in the correct orbit.

2 In the case of a misplaced satellite, the client has nothing to lose. Even if the vehicle is damaged on recovery, or if the operation fails completely, the customer is no worse off than before the recovery effort if the contract states that the customer does not pay unless the mission is completed. However, if the rescue is successful and the client is recovered, the customer will have a fully operational satellite at, in many cases, a fraction of the replacement cost. The high performance alternative is the only design that can perform such a mission. High performance missions are necessary to prove the value of the SOTV in a space recovery market where no other unique alternatives exist. With a functional, high performance SOTV in operation, other SOTV-type spacecraft could be produced with varying levels of functionality, based on the needs of the customer. One possible option is the low cost/upper stage alternative discussed in this report. A high performance SOTV in operation not only proves the reliability of the technology, but also allows for rescue of a potential customer in the case of failure of a solar thermal upper stage.

3 Chapter 2: Solar Orbit Transfer Vehicle

2.1 Overview Appendix D shows a design summary of the SOTV, including subsystem specifications and basic performance estimates. The primary structural component of the SOTV is a hydrogen propellant storage tank. A solar thermal propulsion system is attached to one end of the tank, with an inflatable solar collector mounted so that the collector can be rotated to track the sun about one axis independent of the rest of the spacecraft. The spacecraft bus, which houses the onboard computer and communications equipment, is mounted at the opposite end of the tank from the propulsion unit, along with the low gain and high gain antennas. Two robotic arms for client spacecraft grappling are mounted on the sides of the spacecraft bus structure. SOTV uses solar thermal propulsion, whereby a solar collector focuses the Sun’s energy on a graphite block, heating it to a maximum temperature of 2400 K. Hydrogen at relatively low temperature (approximately 100 K) is injected into the block, where it flows through passages carved in the graphite. As the hydrogen flows through the passages, heat transfers from the block to the hydrogen, raising the temperature of the hydrogen to approximately the temperature of the block. The hydrogen is then expanded through a converging-diverging rocket nozzle to produce thrust. As heat transfers from the block to the hydrogen propellant, the temperature of the block decreases. Once the temperature of the block decreases by a specified amount, the propellant flow ceases and the solar collector reheats the block. This “burn-coast-burn” cycle is repeated numerous times during an orbit transfer. The propulsion system is discussed in detail in Chapter 4. The SOTV employs two mechanical arms to capture client satellites by gripping the launch vehicle adapter of the client. Each arm is equipped with two video cameras to facilitate the docking process. A technician at a ground station uses the video feed to aid in the operation of the arm. The mechanical arms are discussed in Chapter 9. The video signals from the cameras are transmitted to the ground station via a high gain parabolic antenna, which is discussed in Chapter 10. Operations personnel can also use the cameras as a visual tool to locate any problems that may occur during operation of the SOTV. A helix antenna transmits normal telemetry data are transmitted to the ground. An omni-directional antenna provides redundancy in case the helix antenna malfunctions. Four Control Moment Gyroscopes (CMGs) are used for rotational maneuvers and to balance disturbance torques. Slewing maneuvers are accomplished by changing the gimbal angles of the control moment gyros, each of which rotates at a constant angular velocity. The CMGs selected have the ability to rotate through a 360° gimbal angle, which eliminates the need for momentum dumping. A system of

4 star sensors determines the attitude of the SOTV. Chapter 7 discusses the attitude determination and control system. Electrical power for the SOTV is provided by 25 kg of rechargeable Nickel-Hydrogen batteries. Flexible thin film solar cells mounted around the edge of the primary collector recharge the batteries. The hydrogen propellant is stored in a large cryogenic tank that also functions as the primary structural member of the SOTV. The tank stores the hydrogen at a pressure of 1atm (100 kPa) in liquid form. Graphite fiber insulates the tank, effectively minimizing the amount of propellant “boil-off.” The tank design and other structural considerations are discussed in Chapter 6. The SOTV is designed to launch on a Lockheed Martin Atlas V-522. The payload fairing size of the Atlas V restricts the volume (and thus storage capacity) of the propellant tank. The SOTV stores just enough propellant to transfer a 3,000-5,000 kg client from LEO to GEO and then, if fuel permits, return to LEO to await a refueling operation. For this reason, minimizing propellant usage is an issue critical to the design of the SOTV.

2.2 Modeling Various Mathematica and MatLab codes model the transfer of client satellites. Although the methods used by the programs are slightly different, both codes use numerical algorithms to solve the differential equations of motion which determine propellant usage, transfer trajectory, and transfer time for a given orbit transfer. These programs can be modified in various ways to predict the influence of input variables (such as client mass) on the performance of the SOTV. The SOTV orbit transfer is modeled similarly to a traditional low thrust spiral maneuver. How- ever, instead of constant thrust, the engine is fired in cycles. The amount of time that the solar thermal engine can produce thrust depends on the mass flow rate of the propellant, the mass of the block, and the amount of heat that radiates from the block during a thrusting cycle. As the engine produces thrust, the temperature of the graphite core drops and the propellant efficiency of the engine decreases. Chapter 4 discusses the engine in more detail. Once the temperature of the graphite block drops to the point where the propellant efficiency (as indicated by specific impulse, Isp) is below some predetermined value, thrusting stops and the graphite block must be reheated until the core temperature is restored to its maximum value. This heating process is completed using a combination of two primary inflatable solar collectors and two secondary concentrators. The primary collectors are further explored in Chapter 3. Once the temperature of the core reaches its maximum operational temperature, thrusting resumes and the cycle of burning and reheating contin- ues. The mass flow rate of the propellant, the specific heat capacities of the propellant and the block, and the minimum and maximum core temperatures determine the burn time of the engine. The approximate relationship used in the transfer model to determine thrust time is given below:

5 t = Cp (T − T )m /(m& T Cp ) b C max min block max H (2.1)

where tb designates burn time, CpC and CpH are the specific heat capacities of the graphite block and the hydrogen propellant, respectively, Tmax is the maximum temperature of the graphite block, Tmin is the & minimum temperature of the block, m represents the mass flow rate of the propellant, and mblock is the mass of the graphite block. The size of the primary and secondary concentrators and the mass and specific heat capacity of the graphite block determine the amount of time required to reheat the block. This relationship is given below: m = − block (2.2) tc CpC (Tmax Tmin ) Aarray

where tc is reheat time and Aarray is the area of the primary collector.

Throughout this chapter, the mass flow rate of the propellant is taken as 0.01 kg/s and Tmin and Tmax are set at 1,680 K and 2,400 K, respectively. For this temperature range an average specific heat capacity of 15.1 kJ/kg is used for the graphite block, as well as an average specific heat of 15.1 for hydrogen. The methods for choosing these parameters are based on the performance of the propulsion system and are discussed in detail in Chapters 4 and 5. Although these relationships represent approximations of the heating and cooling processes, they are sufficient for the orbital modeling algorithms. The orbital modeling algorithms integrate the equations of motion over the burn and coast times based on a given system mass, beginning orbital radius, and final orbital radius. No thrust vectoring is used and eclipse effects are neglected. Furthermore, the model assumes that the solar collector focuses perfectly on the block during the reheating portion of the burn- coast cycle and is off-pointed during the thrusting portion of the cycle. Although the collector would ideally focus on the block whenever possible, this assumption serves to “balance out” the neglect of eclipse and the assumption that the collector would be able to point all of the incident solar energy directly at the secondary receiver. Chapeter 4 discusses in more detail the assumptions and derivations that these relationships are based on. A sample LEO - GEO orbit transfer trajectory, using a system mass (including 3,000 kg payload mass) of 17,500kg, a collector size of 150 m2 and a block mass of 150 kg, is shown in Figure 2.1.

6 Orbital radius, km 40000

20000

-40000 -20000 20000 40000

-20000

-40000

Figure 2.1: A sample LEO- GEO transfer trajectory

Fundamental performance requirements such as propellant usage, transfer trajectory, and transfer time are among the outputs of the orbital model. The orbital model can therefore be used in numerous ways to predict the performance of the SOTV as input values vary. One of the most important design considerations for the SOTV involves the impact of the solar collector size and graphite block mass on propellant usage and transfer time for a given orbit transfer. Increasing the block mass allows the engine to thrust for longer periods of time during the course of a burn-coast cycle. Figure 2.2 shows a plot of propellant usage vs. block mass for a LEO-GEO transfer using 150 m2 solar collector, a 17500 kg total system mass (which includes the mass of the client satellite).

7 Propellant Mass,kg

10000

8000

6000

4000

2000

Block Mass, kg 100 150 200 250 300

Figure 2.2: Propellant usage for a LEO-GEO transfer with no plane change as a function of graphite block size

Figure 2.2 shows that increasing block mass (thus, increasing the burn time and coast time) has very little effect on the propellant usage of the system. The model allows for a tolerance of 100 km in the final orbital radius of the SOTV when solving the differential equations of motions. The propellant usage varies with block size by only a few kilograms. This change is far outside the precision range of the model, given the assumptions on which the model is based. Increasing the block mass increases the total mass of the system and must therefore affect propellant usage; however, the total system mass increase due to a 100 kg block mass increase is less than 1 % and is therefore undetectable at this level of precision. The total propellant usage for a given orbit transfer directly relates to the total amount of time the engine is fired and the mass of the system. System mass remains nearly constant as the block mass is increased (again neglecting the small increase in the mass of the system as the block mass increases). Figure 2.3 shows transfer time plotted vs. block size.

8 Transfer Time, s 35

34.5

33.5

32.5

Block Mass, kg 100 150 200 250 300

Figure 2.3: Transfer time for a LEO-GEO transfer with no plane change as a function of graphite block size

Once again, the plot shows a bit of erratic behavior due to the tolerance of the final orbit. How- ever, generally, the transfer time is relatively constant as the mass of the block is increased. As the mass of the block increases, the burn time and coast time both increase. Figure 2.3 shows the relationship between block mass and burn time and reheating time: Although increasing the amount of time spent thrusting should affect transfer time, increasing the coasting time should increase the transfer time. However, this increase is not reflected in Figure 2.3, where transfer time stays constant as the block mass (and thus the coast time) increases. The lack of dependence of transfer time on block mass can be explained by the fact that although the coast time increases with block mass for a given cycle, the total number of cycles necessary for a LEO-GEO transfer decreases, as shown in Figure 2.4.

9 Number of thrusting cycles 3000

2500

2000

1500

1000 Block mass, kg 100 150 200 250 300

Figure 2.4: Number of thrusting cycles required to complete a LEO – GEO transfer decreases as block mass increases

Thus, as the mass of the graphite block increases, the time required to reheat the block increases but the total number of reheating cycles decreases. The net effect is that transfer time remains relatively constant as the block mass increases. The effect of the reheating time on the total transfer time is illustrated more clearly when the block mass is held constant and the solar collector size is varied. Increasing the area of the solar collector decreases the amount of time required to reheat the graphite block. However, increasing the area has no effect on the time that the engine can be fired or the amount of thrust/reheat cycles required to complete the LEO - GEO transfer. Note the assumption that the block is on-pointed only during the reheating cycle, as discussed previously. This effect is illustrated for a 17500 kg SOTV, which includes the mass of a 3000 kg client, in Figure 2.5.

10 Transfer Time, days 55

Array Size, m2 100 150 200 250 300

Figure 2.5: Transfer time vs. solar collector area for a LEO-GEO transfer

Increasing the solar collector size decreases the total transfer time significantly. By taking advantage of inflatable structures technology, the SOTV is able to achieve the large solar collector areas that allow the SOTV to far surpass the transfer times required by the RFP. Figure 2.6 confirms the assertion that increasing the collector size does not affect propellant usage. Propellant usage remains constant as collector area increases.

11 Propellant Mass,kg

10000

8000

6000

4000

2000

Array Size, m2 100 150 200 250 300

Figure 2.6: Propellant usage vs. solar collector area for a LEO-GEO transfer

Propellant usage is strongly affected by the mass of the SOTV-client system, although it is not ef- fected by relatively small increases in block mass or increases in collector size. Figure 2.7shows a linear relationship between the mass of the system and the propellant mass required to transfer a 3000 kg client from LEO-GEO and return to LEO.

Propellant usage, kg

12000

11000

10000

9000

8000

Initial mass, kg 12000 13000 14000 15000 16000 17000

Figure 2.7: Propellant usage vs. total system mass (including 3000 kg client) for a LEO-GEO transfer

12 Figure 2.8 utilizes the propellant usage calculations to illustrate how much mass remains after completing a LEO – GEO transfer, depositing the client, and returning to LEO (note that the initial mass includes the mass of the client but the final mass does not): Final mass, kg

1600

1400

1200

1000

800

Initial mass, kg 12000 13000 14000 15000 16000 17000

Figure 2.8: Final mass vs. total system mass (including 3000 kg client) for a LEO-GEO transfer

Figure 2.7 is used with Figure 2.8 to determine the structural mass ceiling for the SOTV, which depends on the mass of propellant carried in the tank. For example, according to Figure 2.7, a 14,000 kg SOTV requires approximately 12,000 kg of propellant to transfer a 3,000 kg client and return to a 300 km LEO. Assuming no propellant remains after the return trip, this leaves approximately 2,000 kg for the structure of the SOTV. The structural mass and propellant mass estimates are used to determine the mass ceiling, propellant usage, and payload mass. The propellant storage estimation is limited by the tank size, which is in turn limited by the volume of the launch vehicle payload fairing. Based on the volume and mass limitations of the launch vehicle, the SOTV hydrogen storage tank is able to hold only 11,000 kg of propellant, and the total mass of the SOTV is about 14,000 kg with propellant. According to Figure 2.7, this means that SOTV will not be able to complete a LEO – GEO transfer with a 3,000 kg client and return to a 300 km orbit. Depending on the mass of the payload, the SOTV will have to return to some orbit in between LEO and GEO to await a refueling mission.

2.3 Performance The SOTV is capable of accomplishing many types of missions. The primary mission for the SOTV as mandated by the RFP is a LEO – GEO transfer of a 3,000 kg client. Figure 2.9 shows the transfer times predicted for such a mission as a function of client mass.

13 Transfer time, days 36

Client mass, kg 1500 2000 2500 3000 3500 4000 4500

Figure 2.9: LEO – GEO Transfer times for a range of client masses

Similarly, Figure 2.10 shows propellant usage as a function of client mass. Propellant usage,kg

11000

10500

10000

9500

Client mass, kg 1500 2000 2500 3000 3500 4000 4500

Figure 2.10: LEO – GEO propellant requirements for client masses ranging from 1000 to 4500 kg

The transfer times predicted are well with in the 60 – 90 day range stipulated by the RFP. Addi- tionally, the 30 day fast track mission could be accomplished for a small client by dumping all propellant except that needed to rendezvous with the client, transfer the client, and return to an orbit reachable by a refueling mission. Based on Figure 2.10 and the propellant storage capability of the SOTV, the largest client trans- ferable by the SOTV is approximately 4500 kg. Such a transfer would expend all of the propellant available to the SOTV and would not allow for a return trip. The transfer would be completed in approximately 35 days. The structural mass of the SOTV would have to be reduced in order to complete the

14 LEO-GEO transfer for a larger client. The propellant storage tank (see Chapter 6) is designed to withstand internal pressures of 5 atm, despite the fact that the hydrogen is stored at a pressure of 1 atm. Additionally, the tank is designed for a factor of safety of 2. If shell thickness of the tank is reduced, to allow for a smaller design margin, the mass of the SOTV can be reduced by the 500 kg necessary to transfer the 5000 kg client from LEO to GEO.

2.4 Operations The SOTV mission profile can be separated into several segments: deployment, rendezvous, transfer, and return. The deployment mission segment begins with launch, where a Lockheed Martin Atlas V launches the SOTV to a 300 km circular orbit. Once communication with the ground station is accomplished with the low gain antenna, the SOTV inflates the solar collectors and deploys the high gain antenna. During this phase, the operations personnel at the ground station test all of the subsystems of the SOTV and check all of the telemetry data to ensure the SOTV is operating properly. Once inflated, the solar collectors begin focusing solar energy on the graphite block, heating the block to a temperature of 2400 K. When the block reaches the maximum temperature of 2400 K, the main engine can be fired periodically for station keeping. The engine can be fired before the core reaches 2400 K if necessary, but the lower the core temperature, the less efficient the engine. Throughout the mission, the control moment gyroscopes will be gimbaled to balance disturbance torques. A timeline representing the deployment mission profile is shown in Figure 2.11. The rendezvous mission segment begins with the SOTV receiving a command from the ground station to rendezvous with a client. The operations crew at the ground station sends the SOTV orbital elements or equivalent information so that the SOTV can compute a trajectory to reach the client. The SOTV then completes a combined orbit transfer and plane change, if necessary, to rendezvous with the client. Once the SOTV is close enough to the client to obtain visual contact, the SOTV activates the high gain antenna and establishes the video feed from the two cameras mounted on the grappling arms. At this point in the mission, control of the SOTV switches from autonomous mode to ground control mode, and the pilot on the ground maneuvers the SOTV to grappling position using the CMGs and solar thermal engine. Before attempting to grapple the client satellite, the pilot first obtains visual confirmation of the predefined grappling location. The grappling location will usually be the launch vehicle adapter of the client. Once the grappling target has been visually confirmed, the pilot maneuvers the grappling arms and clamps each arm to the target. The pilot then uses the arms to align the center of gravities of the two spacecraft with the thrust vector as closely as possible. Next, the SOTV returns to autonomous operation and prepares to transfer the client. The transfer mission segment consists of the SOTV completing the desired orbit transfer of the client using several thrusting cycles of the solar thermal engine. Each cycle consists of the SOTV firing

15 the solar thermal engine until the temperature of the graphite block decreases to a minimum temperature of 1,680 K, where the efficiency of the system drops dramatically (see Chapter 4 for a discussion of the performance plateau). During the thrusting portion of the cycle, the two inflatable solar collectors are on- pointed whenever possible to prolong the amount of time that the engine can be fired. Once the block reaches the minimum temperature, the propellant flow is terminated and the collectors reheat the block to 2,400 K. The cycle repeats until the SOTV reaches the desired orbit. The control moment gyroscopes rotate the entire SOTV/client system to vector the thrust of the engine whenever necessary. Throughout the transfer, the mechanical arms can adjust the position of the client relative to the SOTV to keep the centers of mass of the two spacecrafts aligned with the thrust vector. The powerful onboard computer of the SOTV can complete most of these operations autonomously; however, the ground station personnel can send commands and corrections to the SOTV periodically. During the transfer segment, the high bandwidth data connection is not necessary; the SOTV communicates with the ground station using the low gain antenna. When the transfer is complete, the SOTV deploys the client satellite by first releasing the grip of the grappling arms on the client. Next, the SOTV rotates in a manner such as to not interfere with the client when the main propulsion system is fired, and the SOTV departs to a safe distance from the client to await the command for the next mission. The SOTV will generally need to refuel after a LEO-GEO transfer mission. The return portion of the mission is less complicated than the transfer segment since there is no client satellite involved. Al- though the SOTV will usually not have enough propellant to return to the initial 300 km orbit, it is desirable for the SOTV to return to the lowest possible altitude to reduce the complexity of a refueling mission.

2.5 Summary The development of the SOTV system consists of many design tradeoffs between subsystems to optimize the performance of the vehicle. For instance, the size of the primary collector directly affects the transfer time, while the initial mass of the system affects the propellant usage. The following chapters develop the subsystems of the SOTV, beginning with the solar collector system.

Figure 2.11: Deployment phase mission timeline

Mechanism De- Heat graphite Launch to LEO ployment block to Wait for • • 300 km altitude Inflate collector 2400 K command • • 28 ° inclination Deploy high gain antenna 11 ½ hours Continuous 1 ½ hours 2 hours

0 1 2 3 4 5 6 17 18 19 Approximate mis- 1 hour Continuous Continuous sion time in hours Establish Perform station- communication Check telemetry data keeping operations link • Ensure proper as necessary subsystem opera- • Balance distur- tion bance torques with CMGs

17 Chapter 3: Solar Collector

3.1 Overview The Solar Orbit Transfer Vehicle requires accurate collection and focus of light into the graphite engine in order to operate efficiently. This requirement is met by using two inflatable off-axis paraboloid mirrors that rotate about an off-focal axis direction. In comparison to rigid refelctors of similar performance, inflatable reflective surfaces save on cost, mass, and on volume when they are stored in the deflated launch configuration.

3.2 Collector Design Considerations The solar collector used on the SOTV is an inflatable off-axis paraboloid reflector. The paraboloid reflector features design elements similar to the NASA In-Space Technology Experiments Program (IN-STEP) Inflatable Antenna Experiment (IAE) built by the L’Garde Corporation. The IAE was flown on STS-77, and was successfully deployed and inflated from a Spartan spacecraft on 20 May 19960. This experiment proved that a large inflatable solar reflector could be inflated to a sufficient degree of accuracy in the space environment. A sophisticated surface measurement system developed by L’Garde was able to verify the desired reflective properties of the paraboloid reflector. The reflector was found to have a surface root-mean-squared accuracy of approximately 1 mm when compared to an ideal paraboloid section, even though the reflector itself has a diameter of approximately 14 m.16 The choice of an inflatable system may seem like an unnecessary risk. The deployment, manu- facture, and dynamic response of rigid structures are thoroughly understood and documented; much more so in comparison to inflatable structures. If the SOTV encounters space debris, such as a bolt, a nut, or even a paint chip, a gas-inflated collector will have a leaking hole, as opposed to just a small hole or dent in a rigid collector. The rigid collector is also a more efficient and dependable optical reflector than the inflatable collector. Rigid space-based optical systems already have an outstanding flight heritage, as proven by their successful use onboard the Hubble Space Telescope. Although a rigid collector is a design possibility, the benefits of using an inflatable collector outweigh the risks. The volume, mass, cost, and projected reliability of an inflatable collector give it a clear technical advantage over rigid collector systems. A L’Garde IAE concentrator derivative is the design of choice for several reasons. First, the design is the ideal size for use in the SOTV. The coasting and thrusting models developed for the SOTV reveal that the size of the collector does not have any effect on SOTV propellant usage, only on transfer time. Modeling results in Section 2.2 show that an equivalent reflector area of 150 m2 can be successfully used in a thrust/coast transfer scheme. Coincidentally, the IAE has a diameter of 14 m, which creates a reflector with a surface area of approximately 150 m2. However, a larger reflector is needed when the

18 material reflective index, wrinkling, holing, and other material and environmental factors are taken into consideration. The SOTV uses two IAE-scale size reflectors, each 154 m2, mounted on opposite sides of the RAX (Receiver-Absorber-Exchanger) engine to provide the necessary energy to reheat the block at a rate fast enough to meet mission timeframe constraints. The graphite RAX engine designed by BWX Technologies features two secondary collectors on opposite sides of the graphite block. Twin collectors that operate with independent rotation and inflation systems will also provide a measure of redundancy to the energy collection and focusing system. The established flight history of the IAE is also a key factor in its selection as the solar collector design for the SOTV. Although it has only been deployed on one mission in a space environment, it represents a significant step ahead of other solar collector designs and technologies that still exist only on paper or in ground laboratory test configurations. Lessons learned about the inflation and deployment dynamics of the first IAE allow engineers to refine and improve the design for SOTV application using data from the IAE space experiment. For example, the deployment process triggered an unpredicted dynamic response in the inflated array. The cause was later traced to pockets of gas that had not escaped from the deflated collector while it was stored in the ejection canister. The oscillations were eventually damped out, and the design proved to be reliable and able to withstand unforeseen minor loads such as these deployment vibrations0. Although the SOTV collector is similar in scale to the IAE, there are many aspects of it that are different. The IAE design represents a successful system design with space flight heritage that can be applied to the needs of the SOTV. The SOTV collector does not have the exact same physical dimensions or properties as the IAE, because the IAE is an on-axis paraboloid mirror that re- ceives the light that it transmits. The production and inflation techniques are very similar for the SOTV and the IAE because of their similar sizes.

3.3 Solar Collector Subsystem Description The solar collector can be divided into three major parts: the deployment/inflation system, the torus and support struts, and the lenticular reflector. The deployment and inflation system is contained in an ejecton canister with a volume of 1.2 m2. This volume is the same as that of the IN-STEP IAE sys- tem0. However, the SOTV’s ejection canisters are in the shape of a truncated cone, with the small diameter facing the engine and the large diameter opening away from the engine. This prevents the focused solar energy from damaging any support structure, should a slight focusing error occur. The inside of each ejection canister is coated with a thin layer of reflective aluminum to scatter any light that hits it. Two of these deployment systems fit next to the RAX engine during the launch phase. After the launch phase, the thermal engine is extended away from the deployment boxes to prevent damage to the rotation systems. Each deployment box will rotate about an axis parallel to the y-axis of the SOTV.

19 Two independent SADAs (Solar Array Drive Assemblies) are used to rotate each collector and its inflation hardware in order to accurately maintain the focal point within the secondary collectors of the RAX engine. The collector used is a Moog Model 2-2 SADA, shown below in Figure 3.1. For certain orbital inclinations and sun-angles, rotating arrays allow for continuous thrust missions to be executed when transfer time is a concern. For burn-coast-burn missions, a rotating collector allows the SOTV to maintain a slow rotation about the orbit normal axis while the primary collectors keep a steady focus of light into the secondary collectors of the graphite block. Although slight adjustments with the ACS systems are necessary to compensate for the rotating primary collectors, the power expended to accomplish these maneuvers is far less than the power expended to rotate the entire spacecraft.

Figure 3.1: Moog Type 2 Solar Array Drive Assembly0

Table 3.1: Moog Type 2 SADA Characteristics0

Specification Units Type 2 Output Step Angle Degrees 0.02 Harmonic Drive Ratio - 100:1 Motor Step Angle Degrees 2 Output Torque Angle N-m 11 Holding Torque Powered N-m 9 Holding Torque Unpowered N-m 4.5 # Power # 22 Amps 2 Slip Ring # Signal # 12 Amps 0.5 Voltage Volts 135 Position Sensor Accuracy Degrees 0.08 Type Optical Encoder Total Assembly Mass kg 3

20 In reference to Table 3.1 above, the Model 2-2 is chosen because it has the highest position accuracy (± 0.08 degrees) of any type of SADA listed in Moog's online document. This accuracy is necessary to guarantee the pointing accuracy of the entire solar collector on the secondary collector of the RAX engine. The Moog SADA is also chosen because of its proven spaceflight heritage on spacecraft systems such as Deep Space 1, SSTI, Indostar, Kompsat, Rocsat, and others0. The three inflatable struts and the structural torus of the solar collector are made of neoprene- coated Kevlar. Kevlar is readily available, strong, bondable, flexible, and reasonably priced. The neoprene coating makes the material leak-proof, although extra nitrogen inflation gas is carried to ensure consistent inflation pressures0. The total thickness of the neoprene/Kevlar structural members is 28 P0. The SOTV goes one step farther than the IAE design by using an impregnated resin in the Kevlar material in order to improve the rigidization of the solar collector as a whole. This resin deposits a rigidizing substance; once exposed to the space environment, this resin can be hardened by UV radiation or out- gassing of the resin solvent. This extra rigidizing step will limit the elastic deformation of the beams when the collectors are rotated about the spacecraft, as well as mitigating the effects of holing and other problems encountered in the space environment. The tube diameter of the torus and struts respectively is 610 mm and 460 mm0. Although the reflective mirror has a focal length of only 14 m, it will be necessary to have two struts that are 15.75 m long, and a single strut of length 8.75 m. The resulting conic section should be an off-axis paraboloid section whose focal length between the center of the primary collector and the secondary collector is exactly 14 m. This distance from the vehicle should minimize the effects of shadowing by the SOTV spacecraft, as well as reducing any communications interference caused by a large reflector.

3.4 Mathematical Modeling Each collector is a 14 m-diameter-cylinder and off-axis paraboloid section, meaning that the actual focal point is offset from the normal to the plane of the reflector. The equation for the paraboloid section is:

y2 z2 (3.1) x = + 28 28 which gives a paraboloid with a focal length of 7 m along the x-axis. Note that these axes are not the same axes as the body-centered reference frame of the entire SOTV. The coordinate system for this equation is the same as that of the SOTV system, meaning that this equation represents an initial state where the paraboloid opens toward the x-axis, or the sunlight rays are parallel to the x-axis in the negative-x direction. Three equations are used to determine the paraboloid equation. The first equation is the general equation for a parabola:

21 y 2 x = (3.2) 4 p

where x is the axis of symmetry and p is the focal length of the parabola. Taking the derivative of Equa- tion 3.2, the slope can be set for a line tangent to a specified point (x,y). This equation is: y = 1 (3.3) 2 p

The focal length of the mirror is 14 m, so the Pythagorean equation provides the third equation for the three unknown variables: − 2 + 2 = (x p) y 196 (3.4)

These three equations are now solved for the focal length, p, and the focal coordinate, (x,y). The result is p = 7 and (x,y) = (7,14). To make the parabola a paraboloid, Equation 3.1 contains the additional term z2/28 that creates a 3-D paraboloid surface extending into the x-z plane. Now that the paraboloid is defined, it is necessary to define an intersecting cylinder section to which the struts can attach. The basis for this intersection is a cylinder parallel to the y-axis and centered on the 3-D focal coordinate, (x,y,z) = (7,14,0). The equation for this cylinder is: − 2 + 2 = ( y 14) z 49 (3.5)

This gives a cylinder of radius 7, which results in the proper reflective area that is orthogonal to the sun’s rays. The resulting paraboloid/cylinder intersection yields a set of points to which the support structure can be attached. The optimal location for attachment is at the two points that are coplanar to the center of the cylinder in the x-z plane. These coordinates are (8.75, 14, 7) and (8.75, 14, -7). When the struts are attached to these points, they provide the greatest triangular stability without casting shadows on the reflective surface. The third strut is attached to a point (1.75, 7, 0) on the edge of the collector that is closest to the focal point. In 3-D coordinates, the focal point is at (7,0,0). Unfortunately, the intersection of a cylinder and paraboloid does not yield a shape around which the torus can be modeled. Instead, the intersection of the paraboloid and the plane that is defined by the endpoints of the three support struts serves as the plane of the torus. By taking the cross product of two vectors in the plane (defined by the three strut endpoints), the equation of the plane is found: y − x = 5.25 (3.6)

22 When the plane defined by Equation 3.6 is combined with the paraboloid in Equation 3.1, the result is the conic section torus for use on the SOTV solar collector. The reflective surface area of each reflector can be approximated accurately as a planar circle, with a surface area of 153.9 m2. Modeling the incident light area as a planar circle is an appropriate solution, as the amount of incident light will be limited by the planar cross sectional area of the torus

2 under any circumstances. With two reflectors, the total surface area (ATOTAL) is 307.9 m . The reflector material is made from a single aluminized Mylar gore that is 13 PWKLFNZKLOHWKHFOHDUFDQRS\WKDW forms the other side of the lenticular collector is constructed of 62 gores of transparent 13 P0ylar0. (A gore is one production layer of a thin film material.) Current industry advances limit the reflective efficiency, , of each array to 50-60 %, which creates the need for two solar collectors to heat the block in a timely manner0. The total energy that the solar collector will be able to provide is: E ⋅ A ⋅η = 208 kW SUN TOTAL (3.7)

2 Where ESUN is the solar radiation energy per square meter at Earth’s orbit, 1350 W/m . Finite Element Analysis and experimentation by L'Garde reveal that an inflation pressure of 5.52 3DFDXVHVDVWUHVVRI 3DZLWKLQWKHFDQRS\DQGUHIOHFWRUZKLFKLVVXIILFLHQWWRHOLPLQDWHRUVLJQLIi- cantly reduce most of the wrinkles introduced by the deployment process0. CP1 polyimide films manufactured by SRS Technologies of Huntsville, Alabama may provide more desirable physical properties than the aluminized mylar film, but they lack the spaceflight heritage of the materials used to build L'Garde inflatable structures0. Ground experimentation also shows that they are prone to permanent deformation when wrinkled.

3.5 Deployment, Inflation, and Pressurization Inflation and pressure maintenance of the solar collector are accomplished through the use of compressed Nitrogen gas, stored in tanks at 2.1 MPa within the deployment box of each collector. Each nitrogen tank is about the size of a scuba tank0, having a volume of approximately 0.0024 m2. The inflation scheme is as follows: The lid of the ejection canister springs open, and nitrogen gas begins to inflate the three primary support struts. The support struts and torus are inflated to a pressure of 6890 Pa, and then control valves begin to inflate the solar collector to a pressure of 5.52 Pa. The entire deployment and inflation process takes 5 minutes to complete0. The inflation tanks contain extra nitrogen gas in the event that a small puncture or tear is encountered in the solar collector at some point in its service life. Al- though a resin-impregnated or UV (Ultraviolet)-hardened reflector and canopy was not tested on the IN- STEP IAE mission, this extra rigidization step will prove useful to maintaining the desired optical properties of the collector in the event of a puncture or tear. A rigidized collector requires less inflation pressure to keep the desired surface shape should a puncture occur.

23 3.6 Summary and Further Research The solar concentrator system consists of two inflatable solar collectors that feature materials and deployment systems common to the IN-STEP IAE program. The collectors are 14 m paraboloid off-axis mirrors that rotate about a vector parallel to the y (pitch) axis of the SOTV. The focal points are the secondary collectors located on opposite sides of the SOTV. On-sun pointing is accomplished with the SADA system to an accuracy level that exceeds the RFP standards. The dynamic response of the rigidized inflatable structure is assumed to have an insignificant impact on the exact focal point location relative to the secondary collector. The IAE inflatable structure experienced a harmonic response at 4 Hz, so the SOTV solar collectors of similar size, rigidity, and construction materials should behave in roughly the same way. The slow rotation of the solar collectors should not cause these types of dynamic re- sponses on the SOTV. However, this assumption requires further study of the rigidized struts and torus using finite element analysis.

24 Chapter 4: Solar Thermal Engine and Propulsion

4.1 Introduction The solar thermal propulsion system of the Solar Orbit Transfer Vehicle is unique. The concept of thermal propulsion is simple: heat is transferred to a gas, typically hydrogen, and the temperature and pressure are increased. The gas is then released to create thrust. Solar thermal propulsion means that heat transfer results from solar energy. Currently, the most effective solar themal engine is the Re- ceiver/Absorber/Exchanger (RAX) developed by BWX Technologies. The heating and propulsion of the RAX is a simple process; to heat it, solar radiation is reflected from the primary concentrator and distributed to the graphite cylinder by a secondary concentrator. The locations of the RAX and Primary Collectors are shown in Figure 4.1. Then the graphite is heated to its maximum temperature. When the Thermal Energy Storage (TES) system reaches its maximum temperature, the propulsion process, shown in Figure 4.2 begins. Hydrogen enters the inlet chamber at low pressures and temperatures (1). It then passes through a series of heat exchangers in the TES system (2) until it reaches its maximum temperature and pressure conditions at (3). Finally, the hydrogen is expanded through the nozzle (4). The resulting specific impulses are significantly higher than conventional chemical propulsion, but thrust is higher than typical electric propulsion systems.

Figure 4.1: SOTV illustrating location of Receiver/Absorber/Exchanger and Primary solar collectors

H2, in 1

QSo lar 34H2, out

Figure 4.2: Theoretical solar thermal propulsion engine The purpose of the propulsion design is to optimize the solar thermal propulsion system for the mission of LEO-to-GEO transfer. Solar thermal propulsion systems behave differently than cold gas, chemical, or electric systems, creating new challenges in its design. The characteristics, problems, and proposed solutions to this type of behavior are discussed in this section.

4.2 Overview In the design of a solar thermal engine for the mission of LEO-to-GEO transfer, the ideal engine is one that will: • Minimize fuel consumption • Minimize transfer time • Minimize heat loss • Minimize heating time By minimizing fuel consumption, the SOTV becomes more feasible and profitable. It is reasonable to assume that savings in fuel consumption are obtainable because of the significantly higher specific impulses produced by solar thermal propulsion. By minimizing transfer time, the SOTV will demonstrate further advantages over electric propulsion. The RFP specifies a transfer time goal of 60 days, tolerates a maximum time of 90 days, and requests a study of a fast-track mission of 30 days. Heat loss affects performance by dissipating heat energy that would otherwise be transferred to the propellant. Heat loss prevention is a significant challenge in the RAX because of the extreme tem-

26 perature (2400 K) as compared to the cooler temperatures of the neighboring components. Hydrogen will be stored at cryogenic temperatures and therefore must be thoroughly insulated or isolated from the RAX before its use. Minimizing heating time will allow the RAX to operate more often during a transfer; however, the rate of temperature increase is limited by the size of the primary concentrator. On a system level, an SOTV with an RAX that reheats quickly may also be inefficient because of the added mass of the concentrator. Compromises between these design requirements lead to a SOTV-RAX design combination that minimizes fuel consumption while meeting the transfer time requirements. The major component in the propulsion system is the RAX engine. Shown in Figure 4.2, the RAX consists of an inlet, a low temperature stagnation chamber, a TES system consisting of heat exchangers running through a graphite core, a high temperature stagnation chamber, and an outlet. The RAX will consist primarily of graphite, with an approximately 1 mm thick rhenium coating to prevent the hydrogen from corroding the graphite. Rhenium is preferred for use in interfaces, nozzles, and feed lines because of its stability at both high and low temperatures, and because it will not react with hydrogen0.

4.3 Modeling

4.3.1 Heat Transfer Heat transfer is modeled using a simple energy balance formula as an initial estimate. The ratio of temperature, T, over maximum TES temperature, TMax (in this case, the maximum temperature of the TES) follows the relation:   &  T C p,H mH = −  2  2  (4.1) 1   t TMax  CP,Graphite  mTES  where C is specific heat, m& is mass flow rate of hydrogen propellant, m is the mass of the thermal P H 2 TES energy storage core, and t is time. This useful relationship gives us an estimate of the temperature variations with respect to time, and allows us to estimate the variation of specific impulse, ISP, with respect to time using the relationship:

= T (4.2) I SP I SP,Max TMax

where ISP,Max is the maximum obtainable specific impulse, determined through testing or estimated through the empirical relation:

27 = T (4.3) I SP K MW

where K is a constant based on the ratio of specific heats and Mw is the molecular weight of the propellant. (Hydrogen is therefore ideal, because its molecular weight is the lowest of any material on earth.) Using the function of temperature with respect to time and specific impulse with respect to temperature, we can now determine thrust, F, as a function of time if we assume steady flow: F = m& I g P SP (4.4a)

1 / 4   C  m&   (4.4b) = &  −  p,H 2  H 2   F m P I SP,Max g 1    t   C P,Graphite  mTES  

2 & where g is the acceleration of gravity on Earth’s surface, 9.81 m/s and mP is the mass flow rate of propellant. The above relations gave an estimate of the thrust variations of the RAX. However, when compared with more detailed numerical analyses as performed by Tong0 and experimental data obtained from BWX Technologies, 0 error of more than 20% resulted. Figure 4.3 shows a comparison of theoretical energy balance model results-vs.-experimental results. The phenomenon is attributed to temperature variations throughout the core. Because of the variations, the maximum temperature of the core remained 2400 K even though the average temperature determined through the energy balance method is well below the maximum. Additional heat losses to the surroundings caused further deviation from the predicted energy balance model. However, because the performance of the engine is dependent on maximum temperature and not average temperature of the core, a region of maximum performance known as the “performance plateau” occurred. It is difficult to predict the plateau region without a complex finite element analysis. However, based on the experimental data below, we used an estimate of 9 minutes, in which the engine would perform with maximum thrust and specific impulse.

28 Temperature Profile, Receiver/Absorber/Exchanger 2400

2200

2000

1800

1600

1400 Temperature, K Temperature, 1200

1000 TES Avg. Temp. Exhaust Temp. 800 Energy Balance Model

600 0 5 10 15 20 Time, min.

Figure 4.3: Comparison of energy balance model with experimental results

4.3.2 Orbital Motion Orbital motion is modeled using numerical methods in Matlab and Mathematica. The following equations of motion are used: µ & 2 &r& = rθ − (4.5) r 2 & && − 2 r&θ F θ = + (4.6) r m r & && where r and are polar coordinates with respective derivatives r& , θ , and r&& , θ . F is thrust, and is the gravitational constant for Earth, 3.986012 × 105 km3/s2. These equations are integrated using Matlab’s “ode45” function, a fourth- and fifth-order predictor-corrector method. While burning, the mass variation is assumed equal to the rate of loss of propellant: ∂m = −m& (4.7) ∂t P

29 The first analysis models a constant-thrust transfer, in which thrust is constant at 70 N. Later analyses include temperature variation models in the TES system and coasting arcs during which the TES is reheating. Experimental results obtained from Miller determine burn time.0 Reheating time is determined by: C ()T − T m = P,Graphite Max Min TES (4.8) treheat  Q&    S  A  Sun

 Q&  where S is the effective area of the solar collector and   is the rate of heat transfer per unit area A   Sun

2 from the sun, 1,350 W/m , TMax is the average maximum average temperature of the TES, and TMin is the minimum average temperature of the TES. During the transfer, the TES system heats to 2400 K, then propels the spacecraft until the exhaust temperature drops below the performance plateau range. The process repeats until the spacecraft reaches GEO. The simulation plots the orbital motion, and measures and integrates local velocities and accelerations at each point. Each evaluation measures total transfer time and propellant consumption. Optimizations are also performed, varying TES mass and solar collector area. A more detailed analysis of the optimization can be found in section 2.2.

4.3.3 Alternative Transfers Alternative fuel-saving may also be possible. A simple computer algorithm evaluates a multiple- Hohmann transfer, in which the vehicle thrusts only at periapsis and apoapsis of the coasting arcs. The following equation calculates ∆v:  − &  ∆ = −  m0 m tburn  (4.9) v I Sp g0 ln   m0 

The mass of the vehicle before the burn is m0. Burn time is five minutes, which is short enough to stay within the performance plateau region for maximum performance. The Hohmann transfer arc is calculated using an initial ∆v, and a secondary ∆v is applied at apoapsis until the vehicle reaches its geostationary orbit.

4.3.4 Propellant Feed System The storage tank stores hydrogen in its liquid form at a relatively constant temperature and pressure. Either a pressure relief valve or a regenerative cooling system maintains the pressure in the tank. The system also eliminates the need for a regulated propellant feed such as a blowdown system, although the primary purpose of the pressure control system is to prevent ruptures and leaks.

30 4.4 Performance In the continuous-thrust transfer model, the SOTV completes the transfer in 8-9 days and achieves a mass ratio (Total mass/Burnout mass) of approximately 1.9. Intermittent thrusting has little or no effect on propellant usage, but transfer time ranges from 8 to 45 days depending on solar collector size. Varying TES mass has little or no effect on transfer time and propellant usage. Increasing mass of the TES yields little benefit because the additional burn time gained also results in increased reheat times. Increasing mass results in no gain in time; therefore the mass is minimized in order to minimize requisite structural mass. Increasing the surface area of the solar collector enhanced performance by increasing the heat transfer from the sun, and these mass tradeoffs are beneficial. Multiple Hohmann transfers result in a propellant mass savings of approximately 3%. Transfer time is adversely affected, however, ranging from 65-190 days.

4.5 Summary The propulsion analysis consists of thermal modeling of the engine, orbital motion modeling, and propellant feed system analysis. The following conclusions are drawn: • The specific impulse of thermal propulsion systems are limited by burn time. • Energy balance model assumptions result in error on the order of 20%. • Continuous thrust transfers are not practical using solar thermal propulsion. • Intermittent thrusting has very little effect on the overall fuel consumption of the SOTV as compared to fuel consumption resulting from continuous thrust transfers. • Multiple impulse Hohmann transfers may reduce propellant slightly, but increase transfer time significantly. The preferred engine for the design is the RAX concept developed by BWX technologies, with a mass of approximately 150 kg and a mass flow rate of 0.01 kg/s, allowing for a thrust of 70 N.

4.6 Next Steps There are a number of different thrusting schemes that the solar thermal propulsion system can be used for. Future study of this propulsion system should include: • Continuing to optimize transfers for minimum fuel consumption • Evaluating feasibility of alternative missions, such as interplanetary missions, space station support, and rescuing of client spacecraft • Developing a plan for refueling procedures • Further evaluating the thermal behavior of the engine, by finite element analysis and prototype testing.

31 In the process, we will continue to work on minimizing the mass of the vehicle in order to maximize the performance. As shown in Figure 4.3, the energy balance model acts as a crude approximation for the actual temperature profile of the RAX. A more detailed thermal model, preferably an experimental one, will be necessary for further research.

32 Chapter 5: Thermal

5.1 Thermal system design The thermal system design focuses on three key areas: • The fuel tank must maintain a sufficient amount of fuel for the mission lifetime. • The RAX must minimize parasitic heat loss. • The receiver must be designed to accept and absorb the incoming solar energy. This is accomplished by making the cavity as close to a black-body as possible. The receiver must also be designed to conserve the energy it absorbs and not re-radiate it to the cold space environment. The Thermal Energy Storage (TES) is insulated using graphite felt and Multi- Layer Insulation (MLI). MLI and vapor cooled shields are also used to insulate the fuel tank. Finally, consideration of the interaction between each of the above systems is accounted for. The following sections describe the design considerations within each of these areas.

5.2 Fuel tank The fuel tank is designed to store liquid hydrogen for the SOTV’s mission lifetime, defined as 6 to 9 months by the mission profile. The tank is a vented design employing MLI and Vapor Cooled Shields (VCS) to minimize the boil off of fuel. The resulting boil off passes through a thermodynamic venting system into space. The insulation design takes into account the energy from the Sun, the Earth’s albido and infrared emissions. Heat loss is minimized by three methods: • The exterior shell is an aluminum alloy painted with white enamel, which serves as a barrier serves to protect the MLI from damage and reduces the maximum exterior temperature. • The two large primary concentrators shield the tank, thus yielding a maximum exterior surface temperature of 150 K. • MLI is used to between the inner and outer walls of the tank in conjunction with a vapor-cooled shield (Figure 5.1). The combination of both the MLI and Vapor Cooling Shield (VCS) reduces the fuel boil off to 0.5 kg per day. For a mission lifetime of 9 months this corresponds to 125 kg of fuel lost to boil off. The equations for modeling the performance of MLI are:

T − T Q = k 2 1 (5.1) ;

 3  2   1  ⋅ e ⋅ T   T   T  =  +  h  +  c   + c  Keff hs 1   1    −   N/   2 e   Th   Th  (5.2)

Outer shell

∆ • X2 MLI Q2 VCS • ∆ MLI • X1 Q1g Q1

Inner shell

Figure 5.1: Vapor cooled shield

Reducing the effective conductivity is beneficial to a certain point. Beyond this point, the MLI application becomes too bulky. Vapor cooled shielding allows for a reduction of the heat transfer to a fraction of a watt with a negligible increase in mass or volume. The process removes heat by circulating liquid hydrogen through a tube cooling the shield to which it is brazed. The shield is made of 2mm thick aluminum to provide cooling through conduction. As the mass flow rate of the coolant increases, more heat is transferred away from the inner shell. The equations to determine the performance of the VCS are:

• • • = + Q2 Qg1 Q1 (5.3) • • = Q1 mg hfg (5.4) • • = − Q1g mg Cp (T2 T1 ) (5.5)

The latent heat (hfg) of liquid normal H2 = 445.6 kJ/kg and Q1,2 are calculated from equation 5.1. Cp values for gaseous parahydrogen (boil off of H2) are found in Table 5.1.

Table 5.1: Specific heat for parahydrogen

T,K Cp, kJ/kg K 20.3 12.14 30 10.83 40 10.58 50 10.54 60 10.71 80 11.77 100 13.43

Keff vs. MLI thickness

0.12

0.1

0.08

0.06 Keff

Thickness (cm) 0.04

0.02

0 0 0.0005 0.001 0.0015 0.002 0.0025 0.003 0.0035 0.004 Keff (W/m^2*K)

Figure 5.2: Effective conductivity vs. MLI thickness

Figure 5.3 shows the heat transfer as a function of mass flow rate for a 6 cm application. Provid- ing 6.5 mg/sec of flow rate reduces the heat transfer to the tank to less than 1 Watt.

35 Heat flux (W)

140

120

100

6cm

Q (W) 60

0 2 2.5 3 3.5 4 4.5 5 5.5 6 6.5 7 Mass flow (mg/s)

Figure 5.3: Heat flux vs. mass flow rate

Interaction between the RAX and the fuel tank is a concern; the energy leaving the RAX is approximately 3 kWe. Incorporating a radiation shield between the tank and RAX minimizes the heat flux from the engine to the tank. The shield also serves to re-radiate some of the lost heat back to the RAX.

5.3 Receiver/Cavity Design The RAX receiver is designed to collect and absorb the incident solar energy provided by the solar collectors. The receiver must also be well insulated because of the high temperatures it is required to maintain. There are two primary sources for energy loss in this system: re-radiation losses from the aperture, which is the largest concern, and heat flux to space. These losses are minimized to improve the efficiency of the RAX. Refractory MLI reduces heat loss from the exterior surfaces of the graphite receiver. Minimizing the aperture size of the secondary collector reduces aperture losses. The practical limit for the concentration ratio is determined by cost and technology challenges. Overall concentration ratios in the range of 6000:1 - 7000:1 are baseline with approximately 2000:1 for the primary concentrator and 4:1 on the secondary concentrator. Additionally, the surface area ratio of the cavity is sufficiently large to approximate a black-body cavity. The cavity-to-aperture ratio is 100:1, effectively making it black-body radiator0.

36 The energy absorbed by the cavity walls must be conducted efficiently throughout the mass of the graphite. Fortunately, graphite has a relatively good conduction of 40 W/m-K at 2300 K. During thrusting cycles, the propellant flows into the receiver through a tube and fills the inlet plenum. The propellant then flows uniformly through the axial holes drilled in the receiver. The axial section heats the propellant and discharges it into the exit plenum where it collects and flows out the outlet tube to the nozzle.

5.4 Insulation The material properties of graphite limit the maximum temperature of the receiver to 2,400 K. At 2,400 K the engine begins its “burn” until the propellant at the outlet cools to 2000 K. Multiple layers of MLI using refractory metals and graphite felt reduce thermal losses. Piping, thermocouples, and support pins require penetrations through the insulation. These penetrations reduce the effectiveness of the insulation. Reducing the effective conductivities and emmisivities by 20% account for the penetrations. Re-radiation losses are approximately 3.6 kW at 2400 K. Parasitic losses are on the order of 3kWt. Minimizing heat loss reduces the requirements for solar input power, thus reducing the size of the primary concentrator. Insulation also serves to protect the thermionics from overheating. Directly exposed to 2500 K the thermionics would fail in a short period of time.

5.4.1 Engine performance The RAX performance is based on the mass of the TES system. The more massive the TES, the longer the engine is able to thrust. Burn times, shown in Figure 5.4, are based upon a minimum exit gas temperature of 2000 K. This temperature represents the lower Isp limit at which the SOTV.

37 Burn time

3000.0

2500.0

2000.0

80 N 1500.0 160 N

Temperature (K) Temperature 1000.0

500.0

0.0 0 2 4 6 8 10 12 14 16 18 20 22 Time (minutes)

Figure 5.4: Burn time to 2000 K

Once in orbit, the SOTV must charge the TES. This is a critical time for the spacecraft. Minimizing initial heat time will save fuel, since station keeping operations are performed at sub-optimal Isp levels. Figure 5.5 shows heating times as a function of TES mass.

38 Heat time vs. block mass

18 150 m^2 collector 100 m^2 collector 16 120 m^2 collector Time (hours)

10 100 140 180 220 260 300 340 380 Block mass (kg)

Figure 5.5: Heat time vs. block mass

Re-heat times for the TES directly affect trip time. Re-heat times are calculated based upon the energy input from the Sun minus the parasitic heat loss of the engine (Figure 5.6)

39 Re-heat time vs. collector size

250

200

150

150 Kg TES

100 Time (minutes)Time

0 50 100 150 200 250 300 Collector size (m^2)

Figure 5.6 Reheating times as a function of collector size

5.5 Conclusions Parasitic heat loss is managed by conventional MLI applications and vapor cooled shields for the fuel tank. The design results in less than 1 watt of heat transfer into the tank, and a boil-off of 0.5 kg per day. Based upon a 9 month mission lifetime the resulting fuel loss is approximately 1.25 % over the mission lifetime. The thermal system mass for the LH2 storage is 200 kg including the VCS and MLI system. Conventional MLI and graphite felt applications minimize heat loss in the TES system. Heat loss due to radiation from the aperture is approximately 3.6 kW at 2,400 K. Parasitic heat loss is slightly less at 3 kW. The engine is capable of 10.5 minute burn times at 80 N of thrust and maintains nearly constant specific impulse. Thermal system mass for the RAX is 18 kg.

40 Chapter 6: Fuel Tank/Primary Structure

6.1 Overview The main objective in developing the structure of the SOTV is maximizing the propellant storage capacity while maintaining an overall SOTV mass that is compatible with the launch vehicle capabilities. The maximum amount of propellant that can be stored is 11,000 kg, based on mass estimates of all the components of the SOTV. The cylindrical fuel tank is designed to fill the diameter of the static payload envelope and most of the height of the cylindrical part of static payload envelope leaving room in the hollow launch vehicle adaptor for the spacecraft bus and room in the sloped part of the static envelope for the RAX and the inflatable solar collectors. See Figure 13.9 in Chapter 13 for a diagram of the stowed configuration of the SOTV. The liquid hydrogen fuel tank constitutes the majority of the structural mass of the SOTV and serves as its primary structural member. As the primary structure, the fuel tank bears the loads due to the acceleration of the entire spacecraft during launch.

6.2 Hardware The tank is constructed of aluminum-lithium alloy 1460. With a lower density, higher modulus of elasticity, and higher strength, the aluminum-lithium alloy out performs aluminum 2xxx, 6xxx, and 7xxx alloys. Aluminum-lithium alloys are still relatively new and do not have an extensive flight heritage. The most significant uses of Al-Li in the space industry have been in the Centaur upper-stage and payload adaptors for the Titan IV launch vehicle. A disadvantage of Al-Li is that it is difficult to weld. Also, it has anisotropic properties, meaning the modulus of elasticity is not the same in all directions0.

6.3 Modeling The complex structure of the SOTV includes many components that attach to the fuel tank such as the RAX support system, solar collector support structure, control moment gyros, grappling arms, and the spacecraft bus. The only accurate way to design the fuel tank to be compatible with the structural requirements due to the payload envelope size is to use a complicated finite analysis method. However, in this design report the structure is idealized as a cantilevered cylinder with uniform mass distribution so that simple equations can be used to design the structure. A cantilevered cylinder has one free end and one clamped end. See Figure 6.1 for a diagram of the SOTV fuel tank modeled as a cantilevered cylinder. Approximating the tank as cantilevered cylinder produces a good estimate for the necessary thickness of the fuel tank for the following reasons: • The fuel tank is the majority of mass of the SOTV • The mass of the tank is close to evenly distributed

41 • The tank is clamped to the launch vehicle adaptor, similarly to the clamped end of the cantilevered cylinder. The process used in this report to design the fuel tank is similar to the process described in Chapter 11 of Reference 0.

Figure 6.1. Laterally and Axially Loaded Cantilevered Beam0

4.47 m

9.91 m

Figure 6.2: SOTV fuel tank modeled as a cantilevered beam

To account for room necessary for stringers and insulation on the outside tank, the diameter of the tank is set at 4.47 m. The following equation gives the corresponding length of a cylindrical fuel tank that contains 11,000 kg of liquid hydrogen propellant:

42 mH L = 2 (6.1) πR2 ρ H2 r where mH2 is the mass of propellant, R is the tank radius, and H2 is the density of liquid hydrogen, which is 70.7 kg/m3. The height of the tank is calculated as 9.91 m.

Table 6.1: Fuel Tank Characteristics and Requirements

Geometry: Cylinder Length = 9.91 m Cylinder Diameter = 4.47 m Distributed Mass = 14,200 kg Requirements:

Envelope: The SOTV fits within the Atlas V 522 medium payload fairing envelope (stretched an additional 3 m in length from the standard length). Assume satisfying rigidity requirements will keep the spacecraft’s deflection from violating the fairing’s dynamic envelope.

Mass: Assume that the mass of the SOTV is 14,200 kg

Load Factors: Axial = 5.5 (steady state) + 0.5 (dynamic) = 6.0, Lateral = 2.0 (load factors can be found in Table 13.1 of Chapter 13)

Rigidity: The first axial frequency of the spacecraft must be above 15 Hz. The first lateral (bending frequency) must be above 2.5 Hz.2

Internal Pressure: The tank must support a minimum internal pressure of 5 atm without plastically deforming. A gas pressure release valve will expel propellant when the propellant vapor pressure reaches a level too high for the tank to support.

Factors of Safety: 2.0 (ultimate) and 1.60 (yield) according to recommended factors of safety in Space- craft Mission Analysis and Design for a structure that has no flight history or past structural tests. 0

Material Properties: Aluminum-lithium alloy 1460 is chosen.

43 Young’s Modulus E 80109 N/m2 Poisson’s Ratio 0.30 Density 2,590 kg/m3 6 2 Ultimate Tensile Strength Ftu 760 10 N/m Yield Tensile Strength 6 2 Fty 560 10 N/m

Two design options are chosen for analysis. The first is a monocoque cylinder. A monocoque cylinder has skin only, no stringers or ring frames. Next, a cylinder design with stringers and circumferential rings is studied. The cylinder that withstands the necessary loads and has the smallest mass is chosen.

6.3.1 Monocoque Design Sizing for Rigidity to Meet Natural Frequency Requirement The tank has a uniform thickness and has no rings or longitudinal stiffeners. The natural frequency of the tank is described by the following equation: 1 k f = S (6.2) nat 2π m where m is mass and kS = stiffness = load/deflection (spring constant). The spring constant, ks, is found from mg k = (6.3) S δ where is the deflection. The following equations apply because the fuel tank is approximated as a uniform beam (see Figure 6.2) with a clamped and a free end: Lateral Beam:

 m L3  (6.4) δ = 0.125 B ng  EI 

= EI f nat 0.560 3 (6.5) mB L

Axial Beam:

 m L  (6.6) δ = 0.5 B ng  AE 

= AE fnat 0.250 (6.7) mB L

Parameters: n load factor g gravitational acceleration mB mass of beam (uniformly distributed) I area momentum of inertia of the beam’s cross-section E modulus of elasticity A cross-sectional area of the beam L beam length First, the thickness to satisfy the axial rigidity requirement must be determined. Solving Equation × -3 2 6.7 for A, using an fnat of 15, a value of 6.34 10 m results. The required cylindrical thickness, t, to obtain this A is 0.451 mm. Next, the thickness to satisfy the lateral rigidity requirement is calculated. Equation 6.6 is used × -3 4 with a fnat of 2.5 to solve for the necessary I. The necessary area moment of inertia, I, is 3.45 10 m . The following equation determines the relationship between area momentum of inertia of a cylindrical cross section and its thickness: I t = (6.8) πR 3 where R is the cylinder radius. The required thickness is found to be 0.098 mm. These calculations show that axial rigidity is a more important design factor than lateral rigidity.

6.3.2 Applied and Equivalent Axial Loads The limit loads on the SOTV are calculated using the load factors given for the Atlas V 522 in the Atlas V Mission Planner’s Guide, which are shown in Table 6.2:

45 Table 6.2: Limit loads

Type of Load Weight (N) Distance (m) Load Factor Limit Load Axial 1.39×105 - 6.0 8.36×105 (N) Lateral 1.39×105 - 2.0 2.79×105 (N) Bending Mo- 1.39×105 4.96 2.0 1.38×106 (N-m) ment The moment arm is taken to be half of the cylinder length, for a length of 4.96 m. The equivalent axial load is found using the equation: 2M P = P + (6.9) eq azial R where Paxial is the axial limit load and M is the bending moment limit load. The equivalent axial load is calculated as 2.07×106 N. Next the ultimate load is found using the equation:

Limit Load Ultimate Factor of Safety = Ultimate Load (6.10)

The Ultimate Load is calculated as 4.14×106 N.

6.3.3 Sizing for Tensile Strength The cylinder is sized for axial stress using the equation: P P σ = = (6.11) A 2πRt where is axial stress.

The limit load, Peq, and aluminum lithium alloy’s allowable stress, Ftu, are used to find the necessary thickness. The required thickness is calculated as 0.388 mm.

6.3.4 Sizing for Stability (Compressive Strength) The tank is now sized to prevent axial buckling. The equations used to calculate the cylinder buckling strength are:

 Et  σ = 0.6γ   (for = 0.3) (6.12) cr  R 

where cr is elastic bucking stress, and is the reduction factor used to relate theory to test results given by: γ = − − −ϕ 1.0 0.901(1.0 e ) (6.13)

46 where is found from: 1 R ϕ = (for R/t < 1,500 and L/R < 5) (6.14) 16 t

The critical buckling load, Pcr, is calculated Equation 6.11. Once Pcr is calculated, it is compared with the design load, which is the axial limit load, by calculating the margin of safety (MS) as: MS = P /Ultimate Load – 1.0 cr (6.15)

The margin of safety must be greater than or equal to zero for a design that can withstand the ultimate load. The process to find the required thickness to prevent buckling is iterative. An initial t is chosen and then the corresponding MS is calculated. Different thicknesses are chosen until the value of MS is zero. The t that makes MS zero is the smallest t that will prevent buckling and provides the tank design with least mass. The minimum thickness to prevent buckling is calculated as 6.09 mm. This thickness necessary to prevent buckling is greater than the minimum thickness to meet natural frequency and tensile strength requirements, indicating that compressive strength requirement is the critical design factor for the tank.

6.3.5 Internal Pressure Hoop stress determines the maximum internal pressure that the tank can withstand. Boiling of the liquid hydrogen propellant will create gas pressure on the walls of the tank. The equation for calculating hoop stress is: pR σ = (6.16) h t Where p is internal pressure. The limit pressure and ultimate pressure are calculated by using the values of Fty and Ftu, respectively. The limit pressure is the maximum pressure that elastically deforms the tank and the ultimate pressure is the internal pressure that causes the tank to fail. Values calculated for the limit pressure and ultimate pressure are 15.0 atm and 20.4 atm, respectively.

6.3.6 Mass and Volume of Monocoque Tank The mass of the empty monocoque tank is calculated as:

= ρ π + ρ π 2 me al 2 RtL al 2 R t (6.17)

where al is the density of Al-Li alloy. The mass of the empty tank is calculated to be 2,690 kg. The total mass of the full tank is calculated by adding the mass of the empty tank and the propellant.

47 m = m + m tot e H2 (6.18)

The mass of the full propellant tank is 13,690 kg. As described in Chapter 13, the maximum mass allowable for the SOTV is 14,400 kg, leaving 710 kg of available mass for the other components of the SOTV. These components are estimated to have a total mass of 1,400 kg. Therefore, either the skin-stringer tank design is considered.

6.3.7 Skin-Stringer Design In the skin-stringer design, longitudinal stiffeners bear the compressive loads and prevent buckling. The circumferential rings allow for stiffeners to have a smaller cross sectional area and still resist buckling by shortening the buckling length of the stiffeners. Because the stiffeners prevent axial buckling, the skin no longer needs to bear these loads and does not need to be as thick. The factors that determine the minimum skin thickness are now the acoustic environment of the launch vehicle and the hoop stress due to internal pressure. Modeling acoustic loads are difficult and generally require a computerized finite element analysis.0 Because the SOTV fuel tank is full of liquid, a significant damping of the tank skin is provided, which helps prevent skin failure. The axial acceleration of the liquid in the fuel tank during launch will also help the fluid support the tank skin. Therefore, the critical factor in determining the minimum skin thickness for the SOTV tank is hoop stress. Modeling the hoop stress for a skin-stringer tank differs from modeling a monocoque tank and would require a finite element analysis. Using the monocoque cylinder hoop stress equation, Equation 6.16, will provide a conservative approximation for the internal pressure capabilities for a skin-stringer cylinder. The fact that the tank must be able to withstand an internal pressure of 5 atm before yielding is the primary criterion for determining the minimum skin thickness. Equation 1.16, solved for thickness, uses the yield stress and a pressure of 5 atm. No factor of safety is used in this calculation because this equation already gives a conservative estimate. A minimum thickness of 2.02 mm is calculated. From sizing the monocoque cylinder for axial rigidity, the minimum skin thickness needed is 0.451 mm and from sizing it for lateral rigidity, the minimum necessary skin thickness is 0.098 mm. Both of these thicknesses are less than the minimum thickness required for skin internal pressure capabilities. Therefore there is no minimum cross-sectional area necessary for the stiffeners because the skin alone meets the rigidity requirements. In the SOTV fuel tank design the skin panels are not allowed to buckle. The critical stress for compressive buckling of skin panels is given by:

kπ 2 E  t 2 σ =   (6.19) cr 12()1 −υ 2  b 

48 where b is the spacing between the stiffeners given by: 2πR b = (6.20) n S where nS is the number of stringers used. The number of stiffeners is varied along with stiffener cross-sectional area to find the critical buckling stress. The margin of safety equation, Equation 6.15, ensures that the design is adequate. If the design is adequate, the mass of the tank is calculated and compared to other combinations to determine which has the least mass. The general trend observed is that maximizing the amount of stringers and minimizing the cross-sectional area of each stringer leads to a less massive design solution. The minimum possible cross-sectional area for a stringer due to manufacturing and fastening constraints is 2.0 cm2. Determining the number of ring frames necessary to stabilize the stiffeners is a complex problem that would involve a finite analysis model. Approximately 1 m is an appropriate length for each bay of the tank. Therefore, an 11 ring configuration divides the tank into 10 bays. The estimated mass for the frames and the fasteners that attach the stiffeners and rings to the skin is 25% of the skin and stiffener mass.0 Taking the ring and fastener estimation into account, the mass of the tank is: = ( π ρ + π 2 ρ + ρ )× me 2 RtL al 2 R t al As L al ns 1.25 (6.21)

where As is the cross-sectional area of one stiffener. The optimal configuration for the skin-stringer tank is shown in Table 6.3:

Table 6.3: Optimal skin-stringer tank configuration

Parameter Symbol Value

Number of stiffeners nS 112 Stiffener spacing b 0.125 m Buckling coefficient k 4.3

2 Stiffener cross-sectional area As 2.0 cm

Mass of empty tank me 1,840 kg

Mass of full tank mtot 12,840 kg

The mass of the empty skin-stringer tank is 850 kg less than that of the monocoque tank, which translates into a 32% mass reduction. Therefore, the skin-stringer concept is the most appropriate design for the SOTV propellant tank. The estimated mass of the all the components of the SOTV (excluding the

49 tank) is 1,400 kg, which gives a total mass of the SOTV of 14,240 kg. The mass ceiling of the SOTV design is 14,400 kg, which leaves a 160 kg margin of error for the mass estimate. A Z-shape stringer cross section allows for easy attachment to the skin because the part of the stiffener where the rivets are placed is unobstructed by other parts of the stiffener. Figure 6.3 shows a drawing of the cross-section of the stiffener:

Figure 6.3: Z-Shaped Stiffener (2 Square cm Cross-Sectional Area)

The stiffeners are placed on the inside of the tank so that the outside surface is smooth for insulation requirements. Figure 6.4 depicts the tank cross-section, showing the configuration of the stiffeners.

Figure 6.4: Cross-Section View of Stiffener Configuration

50 Chapter 7: Attitude Determination and Control

7.1 Overview

7.1.1 Background Attitude determination and control are vital operations for a stable spacecraft that is required to operate with precise pointing accuracies while performing delicate maneuvers. The attitude determination portion of the Attitude Determination and Control System (ADCS) provides information about the orientation of the spacecraft with respect to another object. For example, the attitude of the spacecraft may be determined with respect to the Earth, Sun, moon, or stars. Depending on the accuracy requirements and operational altitudes of the mission, one designs an Attitude Determination System (ADS) that will accomplish the mission of the spacecraft in a safe and cost effective manner. In a docking mission, it is imperative that the ADS provide reliable accurate attitude information. Redundancy in the design of the system increases the reliability of the system. In the event that one sensor fails, there are other means of providing attitude information that meet the requirements the system. The attitude control portion of the ADCS provides the spacecraft with the ability to rotate in a manner to fulfill the mission requirements. For example, a solar collector of a spacecraft may be required to point directly at the sun during certain operations; such a requirement may require a rotation of the spacecraft. Another example of an operation that requires rotational maneuvering involves docking with another spacecraft. Rotation about any combination of the three axes of the spacecraft may be required to obtain the proper orientation for rendezvous. Attitude control is also needed for spacecraft stabilization. Disturbance torques act continuously on a spacecraft while in orbit. Gravity gradient, magnetic, solar radiation, and aerodynamic disturbance torques all act to destabilize the spacecraft. These disturbance torques must be counteracted with either torques internal to the spacecraft or by torques applied externally. Destabilization of a spacecraft may render it inoperable, thus incurring a loss to the owner.

Figure 7.1: Spacecraft Body Reference Frame

Before describing each of the disturbance torques, it is advantageous to define the reference frame utilized in this section. In Figure 2.1, a spacecraft in the shape of a rectangular parallelepiped is used to describe the body reference frame. The origin of this frame is defined as the mass center of the object. The mass center signifies the location around which the spacecraft will rotate. Assuming the object is of constant density throughout, the mass center is the dimensional center of the object. With the mass center of the spacecraft defined, each axis can be further explored. The x-axis is the roll axis of the spacecraft and is positive moving from the mass center to the front of the spacecraft. The y-axis is the pitch axis and is defined as positive moving away from the mass center to the right-hand side of the spacecraft. Finally, the z-axis is the yaw axis and in accordance with the right-hand rule, is defined positive down from the mass center. Further definition of roll, pitch, and yaw can be found in Etkin and Reid14. Gravity gradient disturbance torques generally affect spacecraft that are not symmetric about all three axes. In the case of the generic object shown in Figure 7.1, the distance from the mass center to the furthest point in the x-direction is much larger than the furthest point in either the y or z directions. Be- cause of this physical characteristic and the fact that this object is assumed to have a constant density throughout, the mass moments of inertia about both the y-axis (Iyy) and z-axis (Izz) are much larger than that about the x-axis (Ixx). A further explanation of the mass moments of inertia is found in Meriam and Kraige33. Due to this significant difference in mass moment of inertia, the gravity of the Earth has a stronger effect about the y- and z-axes of the spacecraft than the about the x-axis. Thus, the gravity tends

52 to rotate the spacecraft about the y- and z-axes. Equation 7.1 models such a gravity gradient disturbance torque applied on a spacecraft. From this equation one can derive the worst possible gravity gradient disturbance torque as Equation 7.2. 3µ T = u × ()I ⋅u (7.1) g R e e 3µ T = I − I sin()2θ (7.2) g,worst 2R3 z x dev µ 14 3 2 Where is the Earth’s gravity constant, 3.986 × 10 m /s , R is the orbit radius (m), ue is the unit vector in θ the nadir direction, dev is the greatest deviation in radians of the z-axis from the local vertical, and I is the moment of inertia tensor of the spacecraft measured in kg·m2. Worst-case gravity gradient disturbance torques are calculated using the largest and smallest moments of inertia, Izz and Ixx, respectively. Magnetic disturbance torques affect the motion of spacecraft primarily in LEO. This torque depends on the magnetic field strength of the Earth, which decreases with altitude. This magnetic field has a stronger effect on a spacecraft while in polar orbit due the orientation of the magnetic field. The magnetic disturbance torque also depends upon the residual magnetic dipole of the spacecraft. This dipole is created as current passes through the wiring in the spacecraft. As the dipole encounters the magnetic field of the Earth, a torque is applied on the spacecraft in the direction that satisfies the right- hand rule. This worst-case magnetic field disturbance torque can be determined from Equation 7.3. T = DB m,worst (7.3)

2M B = (7.3a) R 3

Where D is the residual dipole of the spacecraft, B is the Earth’s magnetic field, and M is the magnetic moment of the Earth, 7.96 × 1015 T·m3. The solar radiation torque applied to a spacecraft is heavily dependent on the reflectivity of the materials that are exposed to the sun. Equation 7.4 can be used to determine this torque. This equation reduces to Equation 7.5 for worst-case scenario.

  2  (7.4) T = K ()()u ⋅u Au α + r + u 2r + r  × s sr S s n  s d n  s 3 d  c T = F(c − cg) sr,worst ps (7.5)

53 F F = s A ()1+ q cosi (7.5a) c s

-6 2 Where Ks is the solar pressure constant 4.64 × 10 N/m , sc is the vector from the mass center of the α spacecraft to an area A, un is an unit vector normal to A, us is a unit vector in the direction of the Sun, is the absorptivity coefficient of the spacecraft’s surface, rs is the surface specular reflectance coefficient, rd is the surface diffuse reflectance coefficient, cps is the center of solar pressure, cg is the center of gravity, 2 8 Fs is the solar constant 1,367 W/m , c is the speed of light 3.0 × 10 m/sec, As is the surface area of the spacecraft, q is the reflectance factor, and i is the angle of incidence of the Sun. Disturbance torques due to aerodynamic drag also primarily affect spacecraft in LEO. The density of the atmosphere remains significant at these altitudes, thus leading to drag forces on the spacecraft. This torque is determined using Equation 7.6, which reduces to Equation 7.7 for worst-case scenario. 1 T = ρV 2C A()u × s (7.6) a 2 d v cp T = F(c − cg) a,worst pa (7.7)

1 F = ρC AV 2 (7.7a) 2 d

ρ Where is the atmospheric density, Cd is the drag coefficient, V is the velocity, uv is the unit vector in the velocity direction, A is the area perpendicular to uv (surface area), scp is the vector distance from the center of mass of the spacecraft to the center of pressure, F is the drag force acting on the spacecraft, and cpa is the center of aerodynamic pressure.

7.1.2 Requirements The ADCS for the SOTV must meet certain performance criteria specified by the RFP in Appendix D. One such criterion is that the SOTV must have a pointing accuracy of ± 1 degree while having a fine-pointing accuracy of ± 0.1 degrees. This requirement affects the attitude determination system design due to the need of sensors to provide attitude accuracies of better than ± 0.1 degrees. The pointing requirement also requires the attitude control actuators to accurately maneuver the SOTV. The safety of the client satellite also restricts the operation of the ADCS. According to the RFP, the SOTV can only use cold gas propulsion for attitude maneuvers in the vicinity of the client satellite. The thermal engine cannot therefore fire expel hot hydrogen gas in the direction of the client. Gas at such an extreme temperature could damage the delicate sensors on the client. However, the thermal engine can

54 be used for small translational movements when docking, as long as the gas is propelled away from the client. Caution must be used when firing cold-gas propulsion systems in the vicinity of the client as well. However, the probability of damaging the sensors using cold-gas propulsion is much lower than that when using hot-gas propulsion. The loading induced on the client satellite is another safety issue addressed in the RFP that must be addressed in the design of the ADCS. The SOTV must not generate a loading greater than 0.1 g on the client during any phase of the mission. This constraint affects the time required to reach the rotational speed at which the SOTV can maneuver. This rotational speed is referred to as the slew rate. The faster the SOTV-client system changes rotational speed, the more inertial loading is induced on the client. An analysis of the slew acceleration is necessary to guarantee the structural stability of the client while minimizing the time to rotate the SOTV-client system. This analysis is performed using a worst-case scenario for the mass and dimensions of the client satellite.

7.2 Hardware Examples

7.2.1 Attitude Control An attitude control system must be designed to meet two separate needs, spacecraft maneuverability and stability. These needs can be satisfied using a variety of equipment both on an individual basis and in combination with each other. Examples of attitude control systems are listed in Table 7.1.

Table 7.1: Attitude Control Systems49

Weight Power Actuator Typical Performance Range (kg) (W) Cold-gas Thrusters <5N variable N/A 0.4 to 400 N-m-s for momentum wheels at 1,200 to 5,000rpm; Reaction and Momentum Wheels max torques from 0.01 to 1N-m 2 to 20 10 to 110 Control Moment Gyros (CMG) 25 to 4200 N-m of torque >10 kg 90 to 150 2 Magnetic Torquers 1 to 4,000 A-m 0.4 to 50 0.6 to 16

I. Internal Torques: One means of providing attitude control is the use of an internal torque. This means that a torque is created within the spacecraft that acts to rotate the spacecraft in the opposite direction of the external torque. This internal torque can be created using a number of actuators such as reaction wheels, momentum wheels, and control moment gyroscopes. Reaction wheels are used to induce rotation of the spacecraft. An individual actuator can only provide a torque about one axis, the axis around which it rotates. The axis of rotation of the reaction wheel cannot be altered. The internal torque obtained by the rotation of the reaction wheel produces a rotation of the spacecraft about the same axis, but in the oppo-

55 site direction. Combinations of three or more reaction wheels positioned in strategic locations provide the spacecraft with maneuverability about any axis of rotation. The momentum wheel is a similar actuator, also used to provide an internal torque. This actuator is used to balance any disturbance torques that the spacecraft may encounter. This balance is accomplished by spinning the momentum wheel at a rate that produces an internal torque that provides the necessary rotational stiffness about each axis to balance the external torque. The resultant rotational stiffness is about the axis around which the momentum wheel rotates49. One downfall of both reaction and momentum wheels is the small torque range capability. These actuators are suited for operations that require up to 2 N·m of torque. This limits the slew rate capabilities of the spacecraft. For torque requirements greater than 2 N·m one must use multiple wheels per axis of rotation or must select other means of attitude control49. Another negative aspect to the use of reaction and momentum wheels is the need for momentum dumping, required when the actuator reaches its spin rate operational limit. Operating the wheel at speeds greater than this limit could cause damage to itself due to heat or structural failure. Thus, an external force must be applied to the spacecraft to create a moment opposite of the existing torque. This is accomplished using cold-gas propulsion or magnetic torquers. By applying this external force the actuator rotates in the direction to balance the torque that is generated by the external force. Thus, the wheel speed decreases or de-spins. Control Moment Gyroscopes (CMG) provide internal torques within a spacecraft as well. In con- trast to reaction and momentum wheels the CMG can provide torques up to 500 N-m. This actuator operates at a constant spin rate and is gimbaled and therefore, provides the capability of varying the axis around which the resulting torque is applied. Therefore, instead of requiring a separate actuator for each axis of rotation, the CMG can be gimbaled to provide the necessary torque about multiple axes. Disad- vantages of the CMG include large mass, high cost, and large power requirements. In addition, some CMG’s have a maximum gimbal angle at which it can operate without damage. Like a reaction/momentum wheel momentum dumping is required once this operational limit has been reached. Momentum dumping for CMG’s also requires the use of a cold-gas propulsion system or magnetic torquers.

II. External Torques: Attitude control can be accomplished using external torques. This torque can be created using cold-gas thrusters placed at strategic locations. Cold-gas thrusters are simply a means of releasing pres- surized gas to produce a resulting force on the spacecraft. Examples of cold-gas propulsion are included in Table 7.2. The gas is never exposed to a combustion process. Such thrusters can be used to provide rotation of the spacecraft about any axis depending on their location. However, thrusters must be coupled

56 during rotational maneuvers to prevent translational motion. The use of only one thruster for rotation in space not only provides rotation about the desired axis, but also translates the spacecraft in the direction opposite of the propelled gas.

Table 7.2: Examples of Cold-Gas Propulsion49

Mass of Initial Volume Weight of Tank the gas Total Pressure [liter] (empty) [kg] [kg] Mass [kg] [PSI] Technology Titanium & 13 6.7 5 11.7 6,000 Graphic opoxy 11 12 4 16 6,000 Titanium 9.5 17 4.5 21.5 10,000 Titanium Cold-gas thrusters can also provide attitude stability. This is accomplished by providing routine thrusting operations that balance the disturbance torques encountered by the spacecraft. In addition to stability control, cold-gas propulsion is used for momentum dumping when using actuators to create internal torques for attitude control. The need for this capability is discussed in the previous section. Here we focus on the method for which momentum dumping is accomplished. A momentum wheel reaches its maximum spin rate by continually increasing its spin rate to counteract disturbance torques. Therefore, cold-gas thrusters are used to create an external torque in the opposite direction than that already acting on the spacecraft. The momentum wheel tries to balance this torque by decreasing its spin rate. This momentum dumping operation is accomplished using a sequence of impulses from the thrusters until the spin rate of the momentum wheel is well below the operational limit. Cold-gas propulsion is much more reliable than the attitude control actuators previously mentioned. However, the propellant mass required severely restricts the lifetime of the spacecraft. Once the propellant is consumed the attitude control operation can no longer be performed unless a backup system is in place or a refueling option exists.

7.2.2 Attitude Determination: An attitude determination system must be designed to meet the pointing accuracy requirements of the spacecraft. Examples of attitude determination systems are listed in Table 3.1. One type of sensor used for attitude determination is an earth horizon sensor. This type of sensor is only used in Low Earth Orbits (LEO). Even at these low altitudes the sensor is only good for coarse attitude measurements. Therefore, this type of sensor is used for missions where pointing accuracy is not a major design parameter. A sun sensor is an accurate means of determining a spacecrafts attitude. However, this instru- ment cannot provide attitude determination in three axes. It must be coupled with another type of sensor.

57 In addition a sun sensor cannot be coupled with another sun sensor to provide attitude information in all three axes.49 A star tracker is the most accurate of the attitude determination sensors. This type of attitude sensor searches for star patterns that resemble those in its internal star catalog. Once it identifies a pattern the tracker uses the orientation of the spacecraft with respect to the pattern to determine the attitude of the spacecraft attitude with respect to another object. The disadvantage of a star tracker includes high cost and large power requirements. In addition, star trackers have difficulty operating when pointing in the direction of the Sun.

Table 7.3: Attitude Determination Systems49

Mass Range Power Sensor Typical Perforamance Range (kg) (W) Sun Sensors Accuracy = 0.005 deg to 3 deg 0.1 to 2 0 to 3 Attitude Accuracy = 1 arc sec to 1 arc Star Sensors minute (0.0003 deg to .01 deg) 2 to 5 5 to 20 Horizon Sensors <0.1 deg to 1 deg 0.5 to 4 0.3 to 5 Magnetometers Attitude Accuracy = 0.5 deg to 3 deg 0.3 to 1.2 <1

7.3 Modeling and Performance

7.3.1 Modeling I. SOTV Characteristics: The ADCS design is strongly dependent on the mass, dimensions, and location of every component of the SOTV. This is made more difficult by the total mass decrease due to propellant consumption. This variance leads to a change in mass moment of inertia and mass center of the spacecraft. II. Attitude Control: The modeling process for the attitude control is analyzed in two stages. First, each scenario is modeled for a simplified version of the SOTV to determine whether a further analysis is necessary. This analysis focuses on the performance requirements placed on the ACS due to slewing maneuvers while docking with the client spacecraft. Figure 7.2 demonstrates that this simplified configuration consists of only the fuel tank of the SOTV and the fuel within the tank. Rotational effects of the fuel are not considered in this analysis. This simplification places the mass center of the system at its dimensional center point. Based on this configuration the mass moments of inertia about the mass center of the object are 32,500 kg·m2, 121,000 kg·m2, and 121,000 kg·m2 about the x-, y-, and z-axes, respectively.

Figure 7.2: Spacecraft Configuration used for preliminary analysis

If the results obtained from the first stage of the modeling process gives reason to explore the option in more detail, the analysis is extended to include all components of the SOTV. For the complete system case the moments of inertia about the mass center of the object are 61,000 kg·m2, 172,000 kg·m2, and 172,000 kg·m2 about the x-, y-, and z-axes, respectively. III. Scenario AC-1: Slewing maneuvers during the docking process can be accomplished by using a cold-gas propulsion system. This analysis produces estimates for thrust requirements, propellant requirements, and slew maneuver capabilities. To estimate the amount of thrust required from each thruster we first must define the placement of each thruster. This is shown in Figure 7.2 as well. In this analysis it is assumed that each cold-gas thrusters is located equidistant from the mass center as it is thruster counterpart. Propulsion systems 1 and 2 provide rotation about the x-axis and are both located 2.235 m from the objects mass center. Pro- pulsion systems 3 and 4 provide y-axis rotation and are located 4.955 m from the mass center of the object. Like systems 3 and 4, the distance from the mass center to the propulsion systems 5 and 6 is 4.955 m, yet 5 and 6 provide a rotation about the z-axis. Figure 7.3 presents the amount of thrust required from each thruster as a function of slew rate. This is performed assuming that the thrusters use nitrogen as propellant and provide thrust for one second with a specific impulse of 70 seconds. At the end of one second the object has reached its desired slew rate.

59 Thrust Required vs Slew Rate

250

200

150 Roll Pitch & Yaw 100 Thrust Required (N)

0 0 0.10.20.30.40.50.60.70.80.91 Slew Rate (deg/s)

Figure 7.3: Thrust required vs. Slew rate

Figure 7.3 shows that for a very small slew rate such as 0.01°/s a thrust of approximately 2 N is required from each thruster. However, this slew rate is not reasonable for docking procedures that require relatively fast slew rates. Also, shown in the graph is that a slew rate of 1°/sec requires approximately over 200 N of thrust from the cold-gas propulsion system for pitch and yaw maneuvers. This is an unrea- sonable amount of thrust to be produced by a cold-gas propulsion system. Another result obtained from this analysis is that a cold-gas propulsion system cannot meet the ten-year lifetime expectancy of the SOTV without the ability to re-fuel the propulsion tanks. From Table 7.2 the largest tank considered holds 5 kg of propellant. Figure 7.4 shows the relationship between the mass of the propellant required for each maneuver and the slew rate of the SOTV. In addition, Figure 7.5 demonstrates how the number of slewing maneuvers the spacecraft is capable of completing about the roll, pitch and yaw axis based on the propellant mass estimate of 5 kg varies with slew rate. For a slew rate of 1°/s the spacecraft is only able to perform one complete roll maneuver before refueling is required. This includes the thrust required to initiate the rotation and that required to cease the rotation. For pitch and yaw maneuvers there is enough propellant to initiate one rotation about each axis, but not enough to cease the maneuvers once the appropriate orientation is obtained. This is another negative aspect of this system.

60 Mass of Required Cold-Gas Propellant vs Slew Rate

0.7

0.6

0.5

0.4 X-axis Y-axis 0.3 Z-axis

0.2 Mass of required propellant (kg)

0.1

0 0 0.2 0.4 0.6 0.8 1 1.2 Slew Rate (degrees/sec)

Figure 7.4: Mass of propellant vs. slew rate

Number of possible maneuvers vs Slew Rate

800

700

600

500 X-axis 400 Y-axis Z-axis 300

200 Number of possilbe maneuvers

100

0 0 0.2 0.4 0.6 0.8 1 1.2 Slew Rate (degrees/sec)

Figure 7.5: Possible slewing maneuvers vs. slew rate

61 A cold-gas propulsion system has many positive attributes regarding reliability and cost- effectiveness. However, disadvantages such as small slew rates and constant refueling outweigh the advantages. Therefore, this system is ruled out without any further analysis. IV. Scenario AC2: Another means for accomplishing slewing maneuvers uses momentum/reaction wheels. A Con- stellation HR14 reaction wheel (Figure 7.6) made by Honeywell is used to perform this analysis. This reaction wheel provides one of the largest torque capabilities of any of its kind on the market today. Even so, this actuator only produces 0.2 N·m of torque during normal operation.

Figure 7.6: Constellation HR14 reaction wheel28

In the previous section the thrust required from a cold-gas propulsion system was estimated to determine the feasibility of the option. For momentum/reaction wheels we must determine the amount of torque required to rotate the spacecraft. To calculate this torque one must review some basic principles of dynamic motion. To estimate the torque required to rotate a system at a given slew rate one must understand that the torque applied to a spacecraft (g) is the time rate of change of the angular momentum of the spacecraft in an inertial reference frame ( h ) (Equation 7.8). The time rate of change of the spacecraft is defined as

⋅ the sum of the time rate of change of the spacecraft in a fixed reference frame ( h ) and the product of the angular velocity of the spacecraft (ω ) and the angular momentum of the spacecraft in the fixed body frame ( h ) (Equation 7.9). Next, one must define the angular momentum of the spacecraft in a fixed reference frame. As shown in Equation 7.10, h is defined as the sum of the internal angular momentum of the spacecraft due to actuators such as reaction wheels and the product of the mass moments of inertia of the body and actuator I and the angular velocity of the spacecraft. For this estimate it is assumed that

62 the internal angular momentum of the spacecraft is zero. Therefore, h in Equation 7.10 becomes only the product of the mass moments of inertia of the body and actuator and the angular velocity of the spacecraft (Equation 7.11). For this estimate we will assume that the angular velocity is equal to a rotation θ angle ( r) per unit of time (t) (Equation 7.12). Using this definition, Equation 7.11 is rewritten as Equa- tion 7.13. This equation is then differentiated with respect to time to get Equation 7.14. Now, reinserting

⋅ ω into this equation produces Equation 7.15 for h . This equation is then inserted back into Equation 7.9

⋅ to obtain Equation 7.16 for h . Now, it is assumed that currently the spacecraft is not rotating, thus the second term of Equation 7.16 goes to zero. Equation 7.17 is the estimate that is used in this section to estimate the amount of torque that is necessary for a rotation of the spacecraft at a specific slew rate and slew acceleration.

⋅

g = h (7.8)

⋅ ⋅ h = h + ω × h (7.9)

= + ω h h w I (7.10)

h = Iω (7.11)

θ r ω = (7.12) t θ r h = I (7.13) t

⋅ θ r h = I (7.14) t 2 ⋅ ω h = I (7.15) t

⋅ ω h = I +ω × h (7.16) t

⋅ ω h = I (7.17) t

Only the diagonal elements of the mass moment of inertia matrix will be used for estimation purposes. This estimation includes Ixx, Iyy, and Izz and is a reasonable approximation because these moments of inertia are much larger than the off-diagonal elements in the matrix, thus having a much larger influence on the required torques. As previously described, the moments of inertia of the spacecraft in Figure 7.2 are 32,500 kg·m2, 121,000 kg·m2, and 121,000 kg·m2 about the x-, y-, and z-axis, respectively. From Equations 7.12, 7.13, and 7.14 it can be seen that ω is differentiated with respect to time. The resultant is approximated as the angular acceleration (α ). In this case the time variable found in Equa- tion 7.17 is the time the spacecraft takes to accelerate from its initial rotational speed of zero to the desired rotational speed. For this estimation we want an angular velocity of 1°/sec and an acceleration time of 5 seconds. To obtain the correct units for torque (N·m2) one must convert the angular velocity into radians. Using the values described above, the torque required for rotating a spacecraft from an angular velocity of 0°/sec to 1°/sec in 5 seconds about the x-axis is approximately 110 N·m. The required torque about the y- and z-axes are both approximately 420 N·m. As discussed previously, the maximum torque capability for momentum wheels is approximately 0.2 N·m. Momentum/reaction wheel attitude control systems are excellent means of attitude control for small spacecraft. They are reliable, cost-effective, light, and do not require a lot of power. However, the torque capability of this type of system is insufficient for large spacecraft like the SOTV. To demonstrate this, one can plug 0.2 N·m of torque into Equation 7.17 to determine the maximum slew rate possible for the fuel tank while keeping all of the previous assumptions. The result is a slew rate of 0.003 °/sec about the x-axis and 0.0007°/sec about the y- and z-axes. These slew rates are extremely small for the docking procedure of the SOTV. Therefore, this system is ruled out without any further investigation. V. Scenario AC3: Slewing maneuvers can be accomplished using control moment gyroscopes (CMG). A single- gimbaled M1700 CMG (Figure 7.7) made by Honeywell is used to perform this analysis. This CMG provides up to 3170 N·m of torque while having an unlimited range about its gimbal axis. From the momentum/reaction wheel modeling section, the torque required for rotating a spacecraft from an angular velocity of 0°/sec to 1°/sec in 5secs about the x-axis is approximately 110 N·m. The required torque about the y- and z-axes are both approximately 420 N·m. These torque requirements fall well within the limits of the CMG.

Figure 7.7: M1700 CMG528

Although the CMG has disadvantages compared to cold-gas propulsion and momentum/reaction wheel systems such as more mass, higher cost, and higher power requirements, the CMG is the best fit for the performance requirements of the SOTV because it is the only means of producing the torque needed for the SOTV’s docking maneuvers. An estimation is made using the full configuration without the client. As mentioned in the beginning of this section, the moments of inertia of the entire SOTV without the client attached is 61,000 kg·m2, 172,000 kg·m2, and 172,000 kg·m2 about the x-, y-, and z-axes, respectively. Using these moments of inertia, torques of 200 N·m, 600 N·m, and 600 N·m must be applied to the SOTV about the x-, y-, and z-axes, respectively to increase the angular velocity from 0 to 1°/s in five seconds. This remains well within the capabilities of the CMG. Note that the addition of the client is not important for this calculation. The most stringent slew requirements apply during the docking procedure with the client. Once this procedure is accomplished the rate at which the SOTV/client system slews is not as important. Therefore, it is not imperative that this system rotate at 1°/s, and thus the increase of moments of inertia due to adding the client to the calculation is balanced by the decrease in angular velocity requirements for rotation. In addition to the torques due to slew rate requirements the CMG must provide the ability to balance disturbance torques. For this analysis a client spacecraft with a mass of 5000 kg is included in the

65 calculations. The mass center of the SOTV/client system is located 2.51 m, 0.09 m, and 0.06 m from the center of the propellant tank in the x-, y-, and z-directions, respectively. The mass moments of inertia of this SOTV/client system about its mass center are 82,000 kg·m2, 499,000 kg·m2, and 499,000 kg·m2 about the x-, y-, and z-axes, respectively. Figure 7.8 demonstrates the effect of worst-case disturbance torques on the SOTV as a function of altitude. These torques are calculated using Equations 7.1 through 7.7a. As seen from the graph the largest disturbance torques occur at lower altitudes, where the gravity gradient torque dominates. At an altitude of 300 km the gravity gradient disturbance torque applied to the SOTV is 0.84 N·m. This torque must be added to the torque requirement for slewing maneuvers to produce the maximum torque needed

Disturbance Torques vs Altitude

1.00

0.90

0.80

0.70

0.60

0.50 Gravity 0.40 Gradient 0.30 Disturbance Torques (N-m) 0.20

0.10 Aerodynamic 0.00 0 5000 10000 15000 20000 25000 30000 35000 40000 Altitude (km) for a rotation at 300 km.

Figure 7.8: Worst case disturbance torques vs. altitude

Another issue that must be addressed when determining the performance of the SOTV is the total mass decrease due to propellant consumption throughout the transfer of the client. This variance leads to a change in mass moment of inertia and mass center of the spacecraft. Assuming the client is being transferred from an altitude of 300 km to an altitude of 35,000 km,

66 Propellant Mass vs LEO-GEO Transfer Altitude

12000

10000

8000

6000 Propellant mass (kg) Propellant 4000

2000

0 0 5000 10000 15000 20000 25000 30000 35000 Altitude (km)

Figure 7.9: Mass of hydrogen gas propellant as a function of altitude in a LEO-GEO transfer

demonstrates how the mass of propellant decreases as a function of altitude. This calculation is performed using a fifth-order approximation of this relationship given by Equation 7.18. Figure 7.9 shows the change in mass center of the SOTV as a function of altitude:

− − − − (7.18) M (alt) = −9.8496×10 19 R5 +1.3367 ×10 13 R4 − 7.1196×10 9 R3 +1.8894×10 4 R2 − 2.6315R + 22014.778117

Where R is the orbital radius of the spacecraft in kilometers.

Propellant Mass vs LEO-GEO Transfer Altitude

12000

10000

8000

6000 Propellant mass (kg) Propellant 4000

2000

0 0 5000 10000 15000 20000 25000 30000 35000 Altitude (km)

Figure 7.9: Mass of hydrogen gas propellant as a function of altitude in a LEO-GEO transfer

Variance in Center of Mass as a Function of Altitude for a LEO-GEO transfer

in x-direction 3 in y-direction in z-direction

2 Distance from center of fuel tank (m)

0 0 5000 10000 15000 20000 25000 30000 35000 40000 Altitude (km)

Figure 7.10: Change in mass center of the SOTV as a function of altitude in a LEO-GEO transfer

The last analysis that is performed to ensure that the RFP is satisfied is an estimation of the loading imposed on the client satellite. A very basic dynamic loading test is performed on the SOTV/client system to determine the maximum loading on the client during maneuvers. This maximum loading occurs on the structure of the client that is located the farthest from the mass center of the system. From this analysis the maximum loading applied on any feature of the client is approximately 0.025 g, one- fourth of the maximum loading of 0.1 g allowed by the RFP. To obtain an accurate performance model for the CMG’s used on the SOTV one needs to develop control laws for the system. Such analysis is beyond the scope of this report. Ford and Hall0 discuss this subject in detail. The goal of the design of the ACS is to provide ample design margin in the component selection to meet the requirements of the SOTV. To improve this design margin one can use four CMG’s positioned in a square pattern each tilting inward at an angle (Figure 7.11). This provides each CMG the ability to contribute in rotations about multiple axes. This reduces the amount of torque that must be generated by each CMG to produce spacecraft rotation. The use of four CMG’s also increases the reliability of the system. If one CMG ceases to operate the combination of the other three can provide complete attitude control for the SOTV.

Figure 7.11: Positioning of control moment gyroscopes49

70 7.3.2 Attitude Determination 7.3.2.1 Star Trackers/Sun Sensors/Infrared Earth Sensors One possible scenario for attitude determination of the SOTV involves using a combination of star trackers, sun sensors, and infrared Earth sensors. Star Trackers, as previously discussed, are the most accurate sensors available for attitude determination. Most can provide accurate attitude information in all three axes without the aid of another sensor. Unfortunately, the star trackers cannot operate if they are pointing within a 30-60 degree range of the Sun. Operation under these circumstances could be fatal. A redundant system must be in place to ensure accurate attitude determination in the case of star tracker failure. A combination of sun sensors and Earth sensors can provide the needed redundancy. Sun sensors are very accurate, but only provide information regarding two of the three vectors needed for complete attitude determination. They also cannot operate when the spacecraft is in the Earth’s shadow and thus, cannot detect the sun. The sun sensor cannot be coupled with one of its kind to produce full attitude determination. The sun sensor can be coupled with an infrared Earth sensor to fill its shortcomings. This scenario provides the third vector needed for complete attitude determination. The accuracy of this system is one of its disadvantages. An infrared Earth sensor is not as accurate as the sun sensor and its accuracy decreases as the altitude of the spacecraft is increased. In addition the use of these infrared sensors does not remedy the problem of incomplete attitude information while in the shadow of the Earth. Like the Sun sensors, the infrared sensors cannot provide complete attitude information alone. 7.3.2.2 Three Star Trackers Star trackers can be used alone to provide complete attitude information for the spacecraft. As previously discussed they cannot operate while pointing in a direction within 30-60 degrees of the Sun. However, the use of multiple star trackers pointing in different directions provides the opportunity to operate one tracker that is pointing is an operable direction with respect to the Sun, while the others remain protected from the Sun. Star trackers have disadvantages such as high cost, relatively low reliability, and pointing angle restrictions. However, the accuracy of the tracker outweighs the disadvantageous. Using multiple star trackers to obtain attitude information mitigates reliability concerns. Therefore, three CT-633 Stellar Attitude Sensor4 are chosen as the ADS for the SOTV. Each sensor has a 30° shade attached to the end to protect the delicate components from direct sunlight. 4 As previously discussed star trackers are inoperable and very sensitive to damage while focused within a certain range of the Sun. The redundancy of the ADS allows for the operation of only the sensor that is not pointing in the direction of the Sun, while continuing to meet the accuracy requirements of the SOTV.

71 The loss of a sensor would not decrease the pointing accuracy of the SOTV. It would only increase the performance requirements on the ACS. In this case the ACS would rotate the spacecraft to position the operable star sensors away from the Sun. This adjustment to the operational requirements of the ACS ensures the accuracies stipulated by the RFP.

72 7.4 Summary

7.4.1 Attitude Control The attitude control system of the SOTV performs two major operational functions of the SOTV. The ACS provides the capability to rotate the spacecraft about any axis at a slew rate of 1°/sec. In addition, the ACS provides stability against disturbance torques. This operation of the system is extremely important in LEO due to the strong influence of disturbance torques in this region. Four M1700 Control Moment Gyroscopes provide the required torque to accomplish attitude control operations. The specifications for this CMG are found in Appendix D. Each actuator is placed in a pyramid orientation inside the spacecraft bus at an orientation also shown in Appendix D. A design margin is used to size each actuator due to the lack of a complete model of the CMG. The use of four actuators improves the reliability of the system because any three of the CMG’s can meet the performance requirements stated earlier. Therefore, if one CMG should fail the attitude control performance of the SOTV is unaffected.

7.4.2 Attitude Determination The attitude determination system of the SOTV provides complete attitude information with an accuracy of 40 arc seconds. This information is available at all times due to the use of three CT-633 Stellar Attitude Sensors.4 The placement of these sensors is shown in Appendix D. The sensors are attached to the spacecraft at the top of the fuel tank and are spaced 120° from each other. This placement ensures that attitude determination accuracy is possible at all times. The specifications for this sensor are also found in Appendix D.

73 Chapter 8: Secondary Structure

The secondary structures of the SOTV are the RAX support system and the spacecraft bus structure. These structures do not bear the loads generated from the acceleration of the entire SOTV mass as does the fuel tank, the primary structure. Both RAX support system and the spacecraft bus structure are constructed of aluminum-lithium alloy 1460 because of its high strength and low density. The RAX support system holds the RAX in place at a safe distance away from the fuel tank and interfaces with the solar arrays to provide proper alignment. See Figure 8.1 for a picture of the RAX support system. The system consists of hollow circular beams that are 5 cm in diameter and 2 mm thick. The two ring shaped components are 1.2 m in diameter. The distance between the two rings is 0.6 m and the distance between the tank and the top of the ring shaped components is 1 m creating a distance of 1.4 m between the tank and the top of the RAX. The RAX support also incorporates a curved heat-shielding piece between the RAX and tank. Four rhenium wires attach the RAX to the support structure. The wires attach to eyelets on both the RAX support and the outer covering of the RAX. The two ring structures provide a slip-ring joint between the structure and the solar collector canisters allowing them to rotate. The solar array drivers are mounted to the top of the heat shield to provide rotation to array. See Figure 8.2 for a diagram of the solar array interface with the RAX support structure.

Figure 8.1: RAX support system

Figure 8.2: RAX support structure and solar array system

The spacecraft buss structure supports the CMG’s and grappling arms while allowing enough room on the topside end of the fuel tank for other components such as antennas, star sensors, and the computer. See Figure 8.3 for a picture of the buss support structure alone and Figure 8.4 for a picture of it along with the other components on the top end of the SOTV. The main design criterion used in designing the support structure is that it must support the arrangement of the components so that there is no interference with the launch vehicle adaptor. See Figure 8.5 for a picture of the spacing between the launch vehicle adaptor and the components on the topside of the fuel tank.

Figure 8.3: Spacecraft bus support structure

Figure 8.4: Bus support structure with components

Figure 8.5: Launch vehicle adaptor and topside components

76 Chapter 9: Grappling Arms and Mechanical Hand

9.1 Overview The Solar Orbit Transfer Vehicle is required to be able to grapple satellites and hold them during the transfer from LEO to GEO. This will be accomplished through the use of a mechanical arm and grappling hand. Several function and interface needs are inherent to our grappling system: it must be able to interface with some portion of the client satellite in order to firmly grasp the client for purposes of transportation, it must be capable of manipulating the satellite for positioning purposes, and it must not inhibit crucial functions of the client satellite. In addition to these requirements, the arm and hand must conform to standards applicable to the entire SOTV system: it must be durable (capable of lasting 5-10 years), it must damage neither the SOTV nor the client satellite, and it must exert a force less than a 0.1 g upon the client satellite. Since nearly every satellite has a Launch Vehicle Adapter (LVA), the client’s LVA will serve as the point of interface. A common part of many launch vehicle adapters’ is a Marmon clamp, which resembles a ring raised apart from the body of the client satellite with a thickness of 2.5 cm to 12.5 cm. The Marmon clamp interfaces with the LVA inside the payload faring and will be used as the point of interface between the SOTV and the client satellite.

9.2 Hardware For maximum capabilities and to enhance the system’s reliability, two arms are mounted on the SOTV, each a model of the Magnum 7 Function Arm, manufactured by International Submarine Engi- neering, Ltd. This arm has a history of over fifteen years of use in ocean environments, where it adequately deals with temperature and pressure constraints. Each arm mounts on a rotating base with a pivot joint, and has two further pivot points and a gripper hand at its end that mounts on a wrist with a rotation joint. The Magnum 7 robotic arm is shown in Figure 9.1.

Figure 9.1: Magnum 7 robotic arm bending and rotation diagram

The gripper hand consists of a base and a thumb-like prong that has a 7.5 cm wide grip, which will be used to interface with Marmon clamp of the client satellite with a precision of ±13 mm. Each arm is made of aluminum with stainless steel fittings, is one and a half meters long, requires 100 W to operate, and has a mass of 71 kg. During launch, the arm stores flattened against the truss at the upper end of the SOTV. Owing to the frigid environment of outer space it is necessary to employ both heaters and some form of insulation to protect the three hydraulic pumps utilized by the arm. For insulation, several layers of atomic oxygen-resistant polymer wrap around the arm; this insulation was chosen for its efficiency as well as its thinness, which will avoid inhibiting the movement of the arm. The thicknesses of the MLI for the three hydraulic cylinders are 5.625 mm, 4.125 mm, and 3.125 mm, which will allow only 2 W of heat to escape into space. The heat lost will be compensated for through the use of Thermofoil™ Heaters custom made by Minco Products, Inc, which will attach directly onto the cylinders underneath the insulation. When used in combination with the insulation, the heaters will maintain the temperature of the hydraulic pumps at a temperature around 10°C, well above the temperature of -29°C needed to keep the hydraulic fluid from solidifying. Equation 9.136 solves for the heat dissipated through the MLI insulation, which results in the wattage of the heaters. π − 2 kL(T1 T2 ) Q =  D  ln outer   Dinner  (9.1)

In addition to the hardware that permits docking, cameras will attach to the wrist joint as part of a communications system to monitor maneuvers. These cameras assist in docking procedures, providing

78 information needed to avoid collisions. The positions of the cameras provide the view that is least im- peded by the movement of the arm. Several design tradeoffs arose with the decision to use the Magnum 7 arm since there are lighter and more inexpensive arms are in production and development and more complex arms are in use, such as the Canadarm and Canada Hand (used on the space shuttle and the ISS, respectively). However, the Magnum 7 arm has an impressive range of motion (see Degrees of Freedom), as well as a proven record of use and reliability. The arms will be attached to the inner structure of the truss (see SOTV model), and will operate in tandem to obtain maximum stability and accuracy. Due to their placement, the SOTV will be able to move the client so as to allow sunlight to reach the client (thus avoiding inhibiting solar-powered necessary functions), and will be able to secure the client for the duration of transfer.

Figure 9.2: Degrees of freedom

79 9.3 Results and Summary Based upon the needs for the grappling system and SOTV as a whole, the Magnum 7 arm is the best choice to meet these requirements. The heater and insulation system work together to maintain maximum productivity and safety of the arm during performance, while managing to remain light and efficient (power draw on the system is relatively small). The range and degrees of freedom of the pair of arms and hands is more than adequate for the purposes of the SOTV, and their accuracy will prevent damage to the client satellite. In addition, when maneuvering the client satellite the SOTV will not be limited to the capabilities of the arm, but rather the capabilities of the ADCS and the needs of the client satellite.

80 Chapter 10: Communications

The communications system of the SOTV includes the receivers, transmitters and antennas. Re- ceivers and transmitters are sized to accommodate the size and gain of the antennas, the amount of data being passed through it (a.k.a. data rate) and the orbit (i.e. altitude) of the SOTV at the time of transmission. On board antennas are selected and sized based on data rate, orbit, power requirements, ground station antenna size and beam width desired. Two phases of the mission are considered when designing the communications system. One phase is the normal operation of the SOTV while on orbit or during an orbit transfer, referred to as Te- lemetry, Tracking, and Command (TT&C). For TT&C a store and forward method can be used for communication to the ground station. The spacecraft can store data on-board until it passes over the ground station and can transmit the data.38 The other phase is when the SOTV is docking with the client satellite, referred to as Downlink. During docking downlink a continuous communications link is required. One option for this is to use a cross-link communications satellite network. The Tracking and Data Relay Satellite System(TDRSS) is an example of a cross-link network. The TDRSS is not an option for the SOTV mission because TDRSS is located in a medium earth orbit and can only relay data from low earth orbit satellites. Another option for a cross-link satellite network is to send information to a secondary ground station which will then be routed via the Comsat network around the globe to the primary ground station. The SOTV will utilize the second option involving the Comsat network. During the mission planning phase a secondary ground station will be selected that will be in line of sight of the docking maneuver. A powerful transmitter and parabolic antenna will transmit the video data to the secondary ground station. A simple transmitter and helix antenna will also transmit the TT&C data to the secondary ground station. Both data feeds will be transferred via the Comsat network to the primary ground station where a pilot and satellite controllers are located. Table 10.1 gives an overview of the communications hardware on the SOTV which will be discussed in more detail ahead.

Table 10.1: SOTV communications hardware overview

Communications Hardware Overview

TT&C S-Band Transmitter S-Band Receiver S-Band Omni-directional Antenna S-Band Quadrifilar Helix Antenna Downlink Ku-Band Transmitter Ku-Band Parabolic Steerable Antenna

10.1 Telemetry, Tracking and Command Communications Throughout the SOTV mission the spacecraft must communicate with the ground station. During normal operation of the SOTV a store-and-forward technique will be used, where the spacecraft will only communicate while in view of the ground station. The S-Band frequency, ranging from 2.0 to 2.69 GHz, is commonly used for TT&C communication. S-Band communications allow limited data rate transmissions but is sufficient for most TT&C systems. S-Band hardware has been selected for use on the TT&C system of the SOTV The TT&C communications system utilizes an S-band receiver to receive data from the ground station. This receiver is used to collect data relayed from the ground station to the SOTV. The TT&C receiver requires about 5 Watts of power and has a mass of 150 grams. Two transmitters are on board the SOTV. One is dedicated to handle the downlink of TT&C information for the duration of the mission. The other transmitter is used during the docking phase of the mission. The TT&C transmitter operates in the S-band at a frequency of approximately 2.25 GHz. The TT&C transmitter utilizes two antennas. The TT&C transmitter requires approximately 5 Watts of power and has a mass of 150 grams. The primary TT&C antenna is an S-Band Quadrifilar Helix Antenna developed by Condor Sys- tems. The helix antenna is 15.24 cm in diameter, has a height of 8.89 cm and has a mass of 272 grams. It can operate in a frequency range of 2.1 – 2.3 GHz. A picture of the antenna and other system characteristics can be found in the appendix. The secondary TT&C antenna is an S-Band Omni-directional antenna, also produced by Condor Systems. The secondary antenna is used for reliability purposes including failure of the primary antenna

82 and events when the primary antenna is not pointed at the ground station. The omni-directional antenna can operate in a frequency range of 1.2 – 2.5 GHz. The antenna has a diameter of 15.2 cm, a height of 8.9 cm and a mass of 340 grams. A picture of the antenna and other system characteristics can be found in Appendix D.

10.2 Docking Procedure Communications The docking procedure will consist of a pilot watching a video feed of the docking maneuver. The SOTV will have two video cameras mounted to view the docking procedure. The docking transmitter will send the video feed to the ground station and must be capable of transferring the required data rate. The docking communications system requires large data rates to be transmitted to the ground station. The camera produces a data rate of approximately 50 Mbps. The Ku-Band frequency, ranging from 11.0 to 20.0 GHz, is commonly used for high data rate communication. Ku-Band transmitters require more power because of the higher frequencies and the large amount of data being transferred. High data rate communications can require high powered, large parabolic antennas to send the signal to the ground station. The power and size of the antenna can be reduced because the signal can be directed in a narrow beam at the ground station. The gain of an antenna is a measure of the beamwidth. An antenna with a narrow beamwidth is referred to as a high gain antenna. A high gain antenna requires less power and a smaller dish compared to a low or medium gain antenna48. The docking Ku-Band transmitter has a mass of approximately 150 grams and consumes nearly 15 Watts of power. The docking antenna is a parabolic antenna with a diameter of approximately 1 meter and a mass of 6.0 kg. The parabolic antenna is mounted to the side of the bus and must be deployed from a launch position to a usable position before the first docking procedure. Specifications for the communications hardware can be found in Appendix D.

10.3 Conclusions The communications system of the SOTV is unique in that it includes two phases. One phase is the TT&C system to be used throughout the mission. The second phase occurs during docking maneuvers. The TT&C phase, operating in the S-Band, utilizes a transmitter, receiver, a primary helix antenna and a secondary omni-directional antenna. The docking phase, operating in the Ku-Band utilizes a transmitter and a high gain parabolic antenna. The hardware designed for this system meets or exceeds the requirements to communicate and control the SOTV from the ground station.

83 Chapter 11: Command & Data Handling

The SOTV Command & Data Handling (C&DH) system is the computer of the spacecraft. Table 11.1 lists the requirements of the software to keep the SOTV operational. The table shows the size and throughput required by each application in the computer. Each application will be stored in the computer memory so adding the size of each application gives an estimate for the required memory of the computer. The throughput (give in KIPS or Kilo-Instructions Per Second) of each application gives an estimate for the required processing power of the computer.

Table 11.1: SOTV itemized computer software requirements

Size Throughput

(Kwords) (KIPS) Communications Command Processing 1.0 7.0 Telemetry Processing 1.0 3.0 Attitude Sensor Processing Star Trackers 0.0 0.0 ADCS 31.8 144.0 Autonomy 15.0 20.0 Fault Detection 6.0 20.0 Other Power 1.2 5.0 Thermal 0.8 3.0 Kalman Filter 8.0 80.0 Operating System Executive 3.5 60.0 Run-time Kernel 8.0 - I/O Device Handlers 2.0 40.0 Test 0.7 0.5 Math Utitlities 1.2 75.0 TOTAL 80.2 457.5

84 11.1 Software Requirements The communications section of the software requirements is divided into two sections, command processing and telemetry processing. Command processing involves the commands sent from the ground station to control the spacecraft. The command processing section is driven by the docking phase of the mission because of the amount of commands being sent to SOTV during docking. The estimation for the size of algorithm to process the commands is 1 Kword (1024 words of code). The throughput is a high 7.0 KIPS because of the intensive docking procedure. The telemetry data is processed to send information about the health of the spacecraft to the ground. The telemetry processing will increase during docking to give more updates on the position and attitude of the SOTV. The Attitude Determination and Control system relies on the computer to process the information received from the sensors and determine the control maneuvers to be performed by the actuators. The CT-633 Stellar Attitude Sensors from Ball Aerospace use internal software, star catalogs and processing to achieve autonomous attitude solutions. The computer has an algorithm that will determine the orientation of the sun based on our position and attitude so the solar collector can be pointed. The attitude control processing may be shut off during docking to allow a pilot to control the spacecraft. As seen in Table 11.1, the Attitude Determination and Control system puts the largest amount of stress on the computer system. A break down of the SOTV ADCS computer requirements is given in Table 11.2. Kinematic Integration is the estimate of the attitude based on sensed rate changes from gyros. Error determination is an optimization algorithm designed to find errors in the attitude vector determined by the star trackers. Precession is the motion of the angular momentum vector and the computer must recognize the motion and compensate for it using attitude control systems (i.e. the control moment gyros). Magnetic control refers to the computers responsibility to compensate for magnetic torques acting on the body of the spacecraft. The computer must take the attitude estimate from the star trackers and control the CMGs to put the spacecraft in the desired attitude. Ephemeris is a list of the spacecrafts position as a function of time and the ephemeris control is used to correct unintended changes in the position of the spacecraft. Orbit propagation is the tendency of the satellite to move off course because of gravity and other external forces.

Table 11.2: SOTV itemized ADCS computer software requirements

Size (Kwords) Throughput (KIPS) Kinematic Integration 2.0 15.0 Error Determination 1.0 12.0 Precession Control 3.3 30.0 Magnetic Control 1.0 1.0 CMG Control (4) 6.0 60.0 Ephemeris Control 2.0 2.0 Complex Ephemeris 3.5 4.0 Orbit Propogation 13.0 20.0 TOTAL 31.8 144.0

Fault detection is important for reliability in the spacecraft. All command sent to the spacecraft from the ground station or other input devices must be checked for fault. During transmission from the ground it is possible for data to become corrupt. The fault detection algorithm checks each bit for errors and only initiates a command when the command is proven safe.

11.2 Docking Procedure Requirements The SOTV mission will require a video link to the ground station for docking control. The cameras used will be black and white with 256 shades of gray. Each picture/sample requires 8 bits with 3 million samples produced per second. This leads to a data rate of, approximately, 25 Mbps per camera with two cameras on-board. 50 Mbps is a tough load for the communications system to transmit to the ground station. Video compression is an option that could be researched in order to ease the load on the communications link. While producing a video feed is not computer intensive, compressing video is very computer intensive. The trade off of a more powerful computer to a more powerful communications link will most likely result in a decision of improving the computer system. Selecting a computer with more processing power allows for the integration of a video compression utility. Further study could be done on how to integrate video compression into the C&DH system to ease the load on the communications system. The computer selected is the Honeywell Radiation Hardened PowerPC (RHPPC) Single Board Computer. The computer is radiation hardened and based on PowerPC 603e technology licensed from Motorola. The processor can handle 210 MIPS at 150 MHz. The memory section contains 4MB of

86 SRAM (Synchronous Random Access Memory) for main memory, 4MB of EEPROM (Electrically Erasable Programmable Read-Only Memory) and 64KB of SUROM (Start-Up Read-Only Memory) for boot. The RHPPC uses a PCI (Peripheral Component Interconnect) interface for I/O (Input/Output) connections. The PCI connections will come from ADCS sensors and actuators and the communications system. The overall computer power consumption is 12.5 Watts. The size of the board is 234 mm x 220 mm. The RHPPC is designed to withstand the thermal, radiation and vibration environment of space. The RHPPC uses Wind River System's Tornado™ 2.0 tools and VxWorks™ 5.4 real time operating system.

11.3 Conclusions The Command & Data Handling system of the SOTV includes the computer and two cameras mounted on the grappling arm to be utilized during docking maneuvers. In some aspects the computer is over-designed for its application, specifically the computer processor. This is done to allow for other computer constraints to be satisfied and allows for the option of video compression, which is processor intensive. The video compression would help alleviate much of the load on the communications system during docking. The Command & Data Handling system is more than capable of handling the computing tasks necessary to operate the SOTV.

87 Chapter 12: Power system design

The electrical power system focuses on three key areas. The receiver must be designed to accept and absorb the incoming solar energy. This is accomplished by making the cavity as close to a black-body as possible. The receiver must also be designed to conserve the energy it absorbs and not re-radiated to the cold space environment. The TES is insulated using graphite felt and MLI. Finally, for thermionic consideration, the energy must be converted efficiently to electrical power. The following sections describe the design consideration within each of these areas.

12.1 Power requirements The power requirements for the SOTV are approximately 825 We (Watts-electrical). The modest power consumption merits the exploration of both thermionics and solar arrays for electrical power extraction. The RAX is a convenient source of power, and would eliminate the need of a secondary battery system on the spacecraft. A trade study is performed to conclude the advantage of either system. The power requirements are in Table 12.1.

Table 12.1: Power requirements

ADCS Power consumption Star trackers (3) 30 W Control moment gyros (4) 200 W Actuators for solar collector (2) 200 W Cold gas thrusters (4) 80 W C&DH Computer 15 W Transmitters (2) 15 W Receivers (2) 5 W Mechanisms Grappling arm 200 W Power losses 80 W Total power consumption 825 W

88 12.2 Power Conversion For applications where the SOTV needs to supply more than 800 We of power the specific power of a thermionic power system outperforms the solar array system. The results of the trade study are seen in Figure 12.1. The performance gain for the thermionic system at this power level is almost negligible.

12.2.1 Thermionic power system The RAX utilizes 60 thermionic converters surrounding the receiver. The converter specifications are shown in Table 12.2. The output voltage of the converter is 1.2 volts at 2200 K, requiring 24 converters connected in series to supply the required 28 V +/- 6 V bus voltage.

Solar array vs thermionic system mass

190

170

150

130

solar array thin film 110 thermionics solar array GaS

90 System mass (kg)

30 400 800 1200 1600 2000 Power (Watts)

Figure 12.1: System trade study results

Output voltage of the thermionics is very sensitive to temperature change. Maintaining the minimum bus voltage is one of the most important design considerations for the bimodal system. A DC-DC converter is employed to boost the voltage to the upper limit when operating

89 at 2200 K. This ensures that when the temperature drops to the lower limit the power conversion system is supplying at least 22 V. This system eliminates the need for a secondary battery system on the spacecraft.

Table 12.2: Thermionic converter specifications

Operating temperature Range (Emitter) 1900-2000 K Efficiency >10% Output Voltage >0.38 VDC Output Power 31.25 W Nominal Collector Heat Pipe Temperature 1000 K

Fault tolerance for the bimodal system is an important design criterion. Series-parallel connection of the thermionics provides fault protection in the case of open circuit failures. The failure of a couple of converters out of the 50 will not seriously degrade performance.

12.2.2 On orbit power generation Electrical power generation must be continuous even when the spacecraft is in eclipse. The length of the eclipse time depends on the orbit and the precession of the orbit around the Earth. The maximum eclipse duration for the SOTV mission profile is approximately one hour. If electrical power is to be maintained for this duration, then the RAX must have sufficient thermal energy stored for both thrusting and power modes. Also the extraction of this energy will cause an associated temperature drop. Not all of the energy stored in the RAX can be extracted because the power systems will not operate efficiently below 1900 K. With the temperature range of 300 K determined for the thermionics and a power requirement designated, the size of the TES system for the thermionics can be determined. Taking into account the heat loss and converter efficiency the mass estimate for the TES can be seen in Figure 12.2.

90 TES Mass Required (60 Minute Eclipse)

130

120

110

100

800 W 1000 W 80 1200 W 1400 W 70 TES Mass (kg) TES

30 0.1 0.12 0.14 0.16 0.18 0.2 Converter Efficiency

Figure 12.2: Receiver mass requirements

12.2.3 Cost and performance considerations Thermionic converters are a natural consideration for power conversion with the presence of a thermal energy source such as the RAX. When power requirements are greater than 800 We the advantages in mass are noticed. As power requirements become even larger the advantages of thermionics become greater. The power density for 1-2 kWe systems is approximately 100 W/kg with an increase to 200 W/kg for 10 kWe systems. Theoretical specific powers of 1000 W/kg are attainable for systems greater than 50 kWe. Specific cost for the power system exhibits the same trend as specific power. For a 1 kWe system the specific cost of the system is $1700/kWe. This cost drops to $1200/kWe for systems around 3 kWe. The only draw back to the thermionic system is an increased re-heat time due to the additional TES mass. This additional heating time increases the mission trip time unless the solar collector surface area is also increased.

91 12.2.4 Thin film solar array The specific power for the solar array system is nearly identical to the thermionic system. Choosing either system at this power level would be acceptable. However, additional considerations lead to the choice of the solar array system. This system will provide immediate power from the pre-charged secondary battery system and will not increase the re-heat time, thus increasing the trip time. To size the system the following equations are used: = − satellite life Ld (1 degradation/yr) (12.1)

The degradation of the thin film cells is approximately 2% per year. A satellite lifetime of 10 years yields a lifetime degradation (Ld) of 18.3%. This degradation increases the beginning of life (BOL) power requirements to 1 kWe. To size the array Equation 12.1 is used.

 P T P T   e e + d d   X X  P = e d sa (12.2) Td The solar array is sized for eclipse times of 60 minutes. Power requirements are identical for eclipse and sunlight, and the efficiencies Xe = 0.60 and Xd = 0.80 for peak power tracking. This analysis yields a 9 m2 array. The thin film cells are mounted on the primary solar collector, with the additional area added to the overall collector area (Figure 12.3). This increase in area corresponds to a 3 % increase in collector area.

Solar cells on outer circumference

Figure 12.3: Solar cells mounted on solar concentrators

92 12.2.5 Secondary battery system Severe operating conditions require special design consideration for the battery system. To keep mass down, the Depth of Discharge (DOD) for the cells is high. The number of charge- discharge cycles is also very high due to the nature of the SOTV mission. NiH2 cells function the best under the severe conditions. For a 60% DOD and specific power of 35 W/kg the system mass is 40 kg.

12.3 Conclusions Specific power for a thermionic or solar array system is nearly identical for an 825 We power requirement. The solar array is chosen for its instant power availability and zero affect on re-heat time. The EOL requirements size the array at 9 m2 adding 3 % to the primary concentrator area and 12 kg of mass. Power in eclipse is provided by a nickel-hydrogen secondary battery system. The mass of this system is approximately 40 kg based upon a 60% DOD. Total mass for the power system is 90 kg including power converters, regulators, and wiring.

93 Chapter 13: Launch Vehicle

13.1 Overview The SOTV RFP limits the launch vehicle selection to either a Boeing Delta IV Medium (5,4 or smaller) or a Lockheed-Martin Atlas V (522 or smaller). The SOTV will store 11,000 kg of liquid hydrogen propellant. Storing 11,000 kg of liquid hydrogen requires a large volume. The main criterion used in the launch vehicle selection is static payload envelope volume, which determines the usable volume of the payload fairing. The Atlas V 522 is chosen to launch the SOTV because its medium size payload fairing (unstretched) provides a static envelope with significantly more volume (approximately 20 m3) than the static payload envelope of the Delta IV Medium (5,4).

13.2 Hardware The Atlas V design incorporates the recently developed 3.8 m diameter Common Core Booster (CCB) powered by a single RD-180 engine.2 Adding up to five Solid Rocket Boosters (SRB) allows the user to customize the Atlas V 500 series to the desired mission. Each Atlas V consists of a CCB, up to five strap-on rocket boosters, a stretched Centaur upper stage (CIII) with either a single engine (SEC) or dual engines (DEC), and a payload fairing (PLF). See Figure 13.1 for a diagram of the layout of the Atlas V components. The Atlas V 500 series also provides the user the option to use either the short or medium size version 5 m diameter payload fairing to accommodate the volume of the spacecraft. The three digit naming convention of the Atlas series is as follows: the first digit represents the diameter in meters of the payload fairing, the second digit stands for the number of solid rocket motors on the vehicle, and the third digit identifies the number of Centaur engines.2 The Atlas V can launch from either Cape Canaveral Air Station (CCAS) or Vandenberg AFB. The launch vehicle is still under development and is scheduled to fly in 2002.

Figure 13.1: Atlas V 500 Launch System2

The interface between the spacecraft and the launch vehicle consists of the payload support system and the payload fairing. The payload support system supports the payload faring on top of the launch vehicle and the payload fairing encapsulates the spacecraft and protects it during ground operations and ascent. Lockheed Martin offers several options for the payload support system that attaches to the spacecraft interface plane (SIP), depending on the spacecraft configuration. The main components of the fairing are the fixed conical boattail that attaches the PLF to the launch vehicle, the base module that encapsulates the Centaur stage, and the cylindrical module that transitions into a constant radius nose section topped by a spherical nose cap that encapsulates the spacecraft.2 See Figure 13.2 for a diagram of the payload fairing.

Figure 13.2: Atlas V Payload Fairing2

As mentioned in the overview, the main reason that the Atlas V launch vehicle is chosen is because of its large payload static envelope found in the medium size PLF. The static payload envelope defines the useable volume for a spacecraft inside the Atlas V PLF. The tank of the SOTV must store 11,000 kg of liquid hydrogen, inducing a large volume requirement on the PLF. The medium size PLF provides an additional 2.74 m of envelope length, which translates into an extra 45 m3 of usable storage volume. See Figure 13.3 for a comparison between the payload static envelopes of the short and medium version payload fairings. The medium payload fairing can be stretched an additional 3 m if properly coordinated with Lockheed Martin.2 The medium payload fairing used for the SOTV is stretched an additional 3 m in length to accommodate its large fuel tank. Stretching the medium PLF 3 m corresponds to a 49 m3 of volume.

Figure 13.3: Atlas V Payload Envelopes2

The payload support system provides the attachment point for the spacecraft and adapts the spacecraft to the Atlas V SIP.2 The SIP for the Atlas V when launching the SOTV will be on the launch vehicle-provided payload support truss. The aft end of the launch vehicle adaptor is attached to the SIP. A payload support truss is necessary when launching heavy spacecraft such as the SOTV. The payload support system provides the mechanical and electrical interface between the spacecraft and the launch vehicle. Its design is tailored to accommodate the orientation of the SOTV in the PLF. See Figure 13.4 for a diagram of Atlas V SIP.

Figure 13.4: Atlas V Standard Interface Plane (SIP)2

The payload adaptor supports the spacecraft on the launch vehicle and provides separation between the payload and launch vehicle when the payload arrives at the proper orbit. There are several types of payload adaptors that Lockheed Martin can provide. However because of the SOTV’s center of gravity location, a customized payload adaptor must be designed. The center of gravity of the 14,300 kg SOTV will be about 4.7 m above the separation plane. Figure 13.5 shows the cg location of the SOTV is not compatible with standard payload adaptors. However, the cg location of the SOTV falls within Evolved Expendable Launch Vehicle (EELV) requirement, which the Atlas V program is striving to meet. The EELV is an Air Force program to design launch vehicles that will be able to support spacecraft with higher mass and cg locations than previously allowable. Lockheed Martin will design a custom payload adaptor for the SOTV that accommodates it high cg position.

Figure 13.5: Allowable Spacecraft cg Location Above Sep. Plane 2

A 2.97 m (117 in.) diameter payload truss has been conceptually developed to support heavy spacecraft and account for their high cg locations. The 914 mm (36 in.) high truss interfaces with the forward stub adapter on the Centaur upper stage.2 The truss is constructed of graphite epoxy struts, titanium alloy end fittings, aluminum alloy forward and aft brackets, and an aluminum alloy forward ring as seen in Figure 13.6.2 The payload envelope using this interface is shown in Figure 13.7. This interface will be used on the Atlas V 500 configurations for evolutionary heavyweight payloads such as the SOTV.2

Figure 13.6: 2.97 m (117 in.) Diameter Truss2

Figure 13.7: Payload Fairing Using 2.97 m (117 in.) Diameter Truss2

100 The SOTV will be stowed with the engine side of the tank at the upper end of the PLF and its “topside” end attached to the launch vehicle adaptor. This configuration optimizes the available volume of the cylindrical part of the PLF static envelope so that the large fuel tank can be accommodated. The customized launch vehicle adaptor must be 2.97 m in small diameter (so that it attaches to the top of the 2.97 m diameter truss), 4.47 m in large diameter (so that it attaches to the outer edge of the fuel tank), and 1.5 m tall (so that enough storage space is available for the topside components). The customized launch vehicle adaptor will provide internal useable volume to stow the topside components of the SOTV. See Figure 13.8 for a diagram of the customized launch vehicle adaptor. Figure 13.9 shows the SOTV stowed in the static envelope of the payload fairing and Figure 13.10 shows a close up of the launch vehicle adaptor and the SOTV components that are stowed within.

Figure 13.8: Customized Launch Vehicle Adaptor

101

Figure 13.9: Stowed SOTV Configuration

102

Figure 13.10: Stowed Bus Components in Launch Vehicle Adaptor

13.3 Performance Launching the 14,300 kg SOTV to a LEO of 300 km is a feat near the upper limit of the Atlas V 522’s capability. In order to maximize the launch vehicle performance, a direct ascent to circular orbit mission profile will be used. The launch vehicle performs to its maximum capability with planar ascent to a 28.5° inclination orbit.3 In order to achieve the necessary altitude, the SOTV will be launched at a 28.5° inclination from CCAS. Figure 13.11 shows the LEO performance capabilities of an Atlas 552 launched from CCAS. The performance of the Atlas 522, which has 3 less SRBs than the 552, is estimated using this graph. Using trends observed in the Atlas V Mission Planner’s Guide3, the Atlas V 522 performance capabilities are approximated as 5,400 kg less than those of the Atlas V 552 for circular LEO orbits. The maximum payload capability for an Atlas V 522 launching to a circular LEO of 300 km is estimated as 14,600 kg. The payload capability is defined as the mass of the SOTV, the launch vehicle adapter, and all other hardware required to support the spacecraft in the payload fairing.3 The largest standard launch vehicle adaptor that Lockheed Martin provides, the type E, has a mass of 104 kg.3 The standard package of spacecraft required hardware has a mass of 8 kg. Other hardware can be added such as PLF thermal shield, which has a mass of 4 kg.3 The custom-built launch vehicle adaptor for the SOTV and the sup- porting hardware is estimated to have a total mass of 200 kg. Based on these estimates and launch vehicle capabilities, the maximum mass the SOTV can have is 14,400 kg.

103

Figure 13.11. Atlas V 552 LEO Capabilities 2

A direct ascent mission is a one Centaur-burn mission. The Centaur main engines are ignited just after Atlas/Centaur separation and the burn is continued until the Centaur and spacecraft are placed into the targeted orbit. Centaur/spacecraft separation occurs shortly after the burn is completed.3 Using a Centaur dual engine configuration adds additional performance that is especially beneficial to large spacecraft flying to low-Earth orbits such as the SOTV.2 The DEC performance increase is primarily from the additional thrust of the second engine. Figure 13.12 shows a typical Atlas V 552 LEO ascent profile. This profile is similar to the profile of a 522, except that the 522 uses only 2 SRBs. Table 13.1 shows the maximum load factors that act on the spacecraft during ascent. From the data in Table 13.1, the maximum axial load (static and dynamic combined) is 6.0 g and occurs at booster engine cutoff (BECO). The maximum lateral load (static and dynamic combined) is 2.0 g and occurs at launch and Main Engine Cutoff (MECO). The load factors given in Table 13.1 act on the spacecraft center of mass and are used in determining if the primary structure of the spacecraft is compatible with the Atlas V. The SOTV tank, the primary spacecraft structure, is designed to withstand these load factors.

104

Figure 13.12: Typical Atlas V 552 LEO Ascent 3

105 Table 13.1: Atlas V 500-series maximum load factors

Load Condition Direction Atlas V 500 Steady State, g Atlas V 500 Dynamic, g Launch Axial 1.6 ± 2.0 Lateral – ± 2.0 Flight Winds Axial 2.4 ± 0.5 Lateral ± 0.4 ± 1.6 Strap-On SRM Seperation Axial 3.0 ± 0.5 Lateral – ± 0.5 BECO Axial 5.5 ± 0.5 Lateral – ± 1.0 MECO (Max Axial) Axial 4.8 - 0.0 * ± 0.5 Lateral – ± 0.2 (Max Lateral) Axial 0.0 ± 2.0 Lateral – ± 0.6 Sign Convention: • Longitudinal Axis: + (Positive) = Compression; - (Negative) = Tension • Pitch Axis: ± May act in either direction • Yaw Axis: ± May act in either direction • Lateral & Longitudinal loading may act simultaneously during any flight event • Loading is introduced through the center of mass of the spacecraft Note: * Decaying to zero

106 Chapter 14: Conclusions

The vehicle described in this report satisfies and exceeds the minimum requirements stated by the request for proposal. The SOTV has the ability to transfer up to 4,500 kg of payload from a 300 km altitude LEO to a 35,600 km GEO. In addition, the system is capable of performing all other missions specified by the RFP, although the system is most likely not practical for interplanetary type missions. Although the SOTV is designed to be able to accomplish the LEO to GEO transfer, such a mission is not likely to be the most profitable use of the SOTV. Completing the LEO to GEO transfer requires nearly all of the propellant that the SOTV is capable of carrying, and refueling missions will be expensive due to unavoidably high launch costs. Dispersion of debris is also not likely to be a profitable endeavor. Large pieces of debris are in- creasingly becoming a problem for functioning spacecraft in the orbital environment. However, it would be difficult to find someone who would pay the costs associated with an orbit transfer of such debris. The most profitable use of the SOTV design outlined in this report is the recovery of lost satellites. Except for the Space Shuttle, an operational SOTV would be the only vehicle capable of performing such missions. The SOTV could also be used to provide routine orbital maintenance to spacecraft whose propulsion systems have failed. Many such missions could be accomplished by the SOTV without requiring a costly refueling mission. With a proven SOTV in operation, more specialized designs could be created. Once such application would be the upper stage concept discussed earlier in the report. Applying thermionic technology to the SOTV design would allow for it to operate as a power source in addition to a propulsion system for a client satellite. As discussed in the power section of this report, the specific power capabilities of thermionics exceed that of batteries as power requirements increase. The SOTV could be a very efficient power source for spacecraft requiring large amounts of power. Many opportunities exist for the spacecraft proposed by this paper. Prospects for further design optimization include: • Reducing the structural factor of safety of the tank to increase payload capabilities and propellant efficiency. • Designing the fuel tank to be launched empty, or with a minimal amount of propellant. Such a design would further reduce the mass of the tank and improve the payload ratio and structural coefficient (thus improving the propellant efficiency) since the tank structure would not have to withstand the launch loads fully fueled. The SOTV would then be refueled before the first mission.

107 • Study of refractive secondary collector technology (as opposed to reflective technology). A refractive collector system would distribute the solar energy more evenly and efficiently. Sapphire has been suggested as a secondary collector material that would have the desired refractive and thermal properties. • Eliminating thermal energy storage altogether and changing to a direct heating scheme. The Solar Orbit Transfer Vehicle design proposed by this paper meets and exceeds the requirements imposed by the RFP. If given further study and development, an operational SOTV would represent a revolutionary advance in spacecraft technology.

108 References

1. Abramson, H. Norman (Editor). The Dynamic Behavior of Liquids in Moving Containers with Application to Space Vehicle Technology. Southwest Research Institute, Scientific and Techni- cal Information Division, Washington: 1966.

2. Atlas Launch System Mission Planner’s Guide Atlas V Appendum (AVMPG). Revision 8. Lock- heed Martin. Available: http://www.ilslaunch.com/missionplanner/. Accessed 30 April 2001.

3. Atlas Launch System Mission Planner’s Guide. Revision 7. Lockheed Martin. Available: http://www.ilslaunch.com/missionplanner/. Accessed 30 April 2001.

4. Ball Aerospace Technologies Operations. “CT-633 Stellar Attitude Sensor.” Available: www.ball.com/aerospace /ct-633.html. Accessed 10/29/2000.

5. Bates, R. R., and Mueller, D. D., and White, J. E. Fundamentals of Astrodynamics. Dover Books, New York: 1971.

6. Birur, G. C. and Tsuyuki, G. T. Cryogenics. Space Cryogenics Workshop, Vol. 32. Cleveland, OH: 1991. p. 185-190.

7. Cassini Spacecraft Website. NASA/JPL. Available: http://www.jpl.nasa.gov/cassini/english/spacecraft/. Accessed 27 Apr 2001.

8. Chmielewski, A. B. and Jenkins, C. H. "Gossamer Structures." Structures Technology for Future Aerospace Systems. Ed. Ahmed K. Noor. Vol. 188 Reston, VA: AIAA, 2000. p. 204-268.

9. Condor Systems, Capabilities and Products, Aerospace Antennas. Available: http://www.condorsys.com/cap_prod/sigint_esm/antenna/aerospasan/index.htm. Accessed 30 Apr 2001.

10. Deep Space Network Home Page. NASA/JPL. http://deepspace.jpl.nasa.gov/dsn/index.html. Ac- cessed 26 April 2001.

11. Driesbach, Frederick J. SRS Technologies, Systems Technology Group. Letter to the author. 31 Jan 2001.

12. Engberg, R. C. and Lassiter, J. O. "Dynamic Testing of Inflatable Spacecraft Structures." Sound and Vibration. Vol. 33, No. 6. Jun 1999. p. 16-20.

13. Engberg, R. C. et al. "Modal Survey Test of the SOTV 2 × 3 Meter Off-Axis Inflatable Concentra- tor." 41st AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference. 3-6 Apr 2000, Atlanta, GA. AIAA Paper 2000-1639.

14. Etkin, Bernard and Reid, Lloyd D. Dynamics of Flight, Stability and Control. 3rd Ed. John Wiley & Sons, Inc., New York: 1996.

109 15. Ford, Kevin A. and Hall, Christopher D. “Singular Direction Avoidance Steering for Control- Moment Gyros.” Journal of Guidance, Control, and Dynamics. Vol.23, No.4, Jul-Aug 2000.

16. Freeland, R. E. and Bilyeu, G. D. "In-Step Inflatable Antenna Experiment." http://www.lgarde.com/people/papers/in-step.pdf. Accessed 3 Apr 01.

17. Freeland, R. E. et al. "Inflatable Deployable Space Structures Technology Summary." http://www.lgarde.com/people/papers/spacestructs.html. Accessed 17 Jan 01.

18. Freeland, R. E. et al. "Large Inflatable Deployable Antenna Flight Experiment Results." IAF (Inter- national Astronautical Federeration) Paper 97-1301. http://www.lgarde.com/people/papers/iaf- 97-1301.pdf. Accessed 3 Apr 01.

19. Gartrell, Charles F. "Future Solar Orbital Transfer Vehicle Concept." IEEE Transactions on Aero- space & Electronic Systems. Vol. 19, No. 5. Sep 1983. p. 704-710.

20. Gary, H. P., Mullick, S. C., and Bhargava, A. K. Solar Thermal Energy Storage. D. Reidel Publish- ing Co., Boston: 1985.

21. Hall, C. D., and Thorne, J. D., “Minimum Time Continuous Thrust Orbit Transfers.” Journal of the Astronautical Sciences. Vol. 45, No. 4. Oct-Dec 1997. p. 411-432.

22. Hall, Christopher D., Associate Professor of Aerospace Engineering. Lecture materials distributed in AOE 4984, Spacecraft Attitude Dynamics and Control, Virginia Polytechnic Institute and State University, Blacksburg, Virginia, 2001.

23. Hatsopoulis, G. N. and Gyftopoulis, E. P., Thermionic Conversion. Vol. 1. MIT Press, Cambridge, MA: 1973. p. 41-43.

24. Hearn, H. C., “Thruster Requirements and Concerns for Bipropellant Blowdown Systems.”, Journal of Propulsion and Power. Vol. 4, No. 1. Jan-Feb 1988. p. 47-52.

25. Hecht, Eugene. Optics. 3rd Ed. Addison Wesley: 1998. p. 180-183.

26. High Pressure Steel SCUBA Tanks. Available: http://www.scuba.com/shop.php. Accessed 1 May 2001.

27. Hill, P. G. and Peterson, C. R. Mechanics and Thermodynamics of Propulsion. 2nd ed., Addison Wesley, New York: 1992.

28. Honeywell Space Systems Electronic Products. Available: http://content.honeywell.com:80/Space/products/Electronics.htm. Accessed 30 Apr 2001.

29. Humble, R. W., Henry, G. N., and Larson, W. J. Space Propulsion Analysis and Design . McGraw- Hill, New York: 1995.

30. Integrated Radar and Communications Subsystem (Ku Band). Boeing (Hughes) Satellite Systems. Available: http://www.hughespace.com/factsheets/scientific/ircs/ircs.html. Accessed 30 Apr 2001.

110 31. Kluever, C. A., and Oleson, S. R. “Direct Approach for Computing Near-Optimal Low-Thrust Earth- Orbit Transfers.” Journal of Spacecraft and Rockets. Vol. 35, No. 4. Jul-Aug 1998. p. 509-515.

32. Main, J. A. et al. "Post-Flight Testing and Analysis of Zero-G Foam Rigidized Struts." 40th AIAA/ASME/ASCE/HS/ASC Structures, Structural Dynamics, and Materials Conference. 12-15 Apr, St. Louis, MO. AIAA Paper 99-1524.

33. Meriam, J. L. and Kraige, L. G. Engineering Mechanics Dynamics. 4th Ed. John Wiley & Sons, Inc: New York, 1997.

34. Miles, B. J. and Rochow, R. F. Power Generation Considerations in a Solar Bimodal Receiver. Aerospace Systems and Technologies, Vol. 1. Washington, DC: 1996. p. 345-350.

35. Miller, B. G. BWX Technologies. Correspondence and site visits. Jan-Apr 2001.

36. Miller, B. G., and Rochow, R. F. “Design and Fabrication of a High Temperature Solar Receiver Cavity.” Aerospace Power Proceedings of the Intersociety Energy Conversion Conference. Vol. 1. 1995, 95CH35829. p. 755-761.

37. Moog, Schaeffer Magnetics Division. “Solar Array Drives.” Available: http://www.moog.com/schaeffer/Sections/sec4.pdf. Accessed 20 Mar 2001.

38. Pisacane, Vincent L. and Moore, Robert C. (Editor). Fundamentals of Space Systems. Oxford Uni- versity Press, New York: 1994.

39. Power Management and Distribution System Developed for Thermionic Power Converters. Avail- able: http://www.lerc.nasa.gov/Other_Groups/RT1997/5000/5450baez.htm. Accessed 30 Apr 2001.

40. Sarafin, Thomas P. (Editor) Spacecraft Structures and Mechanisms. Microcosm Press, Torrance, CA: 1995.

41. Schaub, Hans Peter and Junkins, John L. “Constellation Series Reaction Wheels.” Aerospace Elec- tronic Systems. Phoenix: 2001.

42. Schaub, Hans Peter and Junkins, John L. “Singularity Avoidance Using Null Motion and Variable- Speed Control Moment Gyros.” Journal of Guidance, Control, and Dynamics. Vol.23, No.1, Jan-Feb 2000

43. Sidi, Marcel J. Spacecraft Dynamics and Control. Cambridge University Press, New York: 1997.

44. Space Systems Group Satellite Systems Operations. “Control Moment Gyroscopes.” Honeywell Inc., Phoenix: 1993.

45. Sverdlin, Alexey. "Aluminum-Lithium Alloys for Aerospace." Advanced Materials & Processes. Vol. 153, No. 6 Jun 1998. p. 49-51.

46. Tong, Wei. “Transient Three-Dimensional Thermal Analysis of a Receiver-Absorber-Exchanger System in the Integrated Solar Upper Stage Unit.” Numerical Heat Transfer. Part A, Vol. 36, 2000. p. 807-823.

111

47. Veal, G. and Freeland, R. E. "In-Step Inflatable Antenna Description." AIAA Paper 95-1193. http://www.lgarde.com/people/papers/aiaa-95-1193b.pdf. Accessed 3 Apr 01.

48. Wertz, James R., and Larson, Wiley J. Space Mission Analysis and Design. Microcosm Press, El Segundo, CA: 1999.

49. Wertz, James. R. Spacecraft Attitude Determination and Control. Reidel Publishing Company, Dordrecht: 1978.

112 Appendix A: Interplanetary Mission Considerations

An interplanetary mission that uses the SOTV will rely on the existing design, with minimal changes to the system that has been designed for LEO-GEO orbit transfers. The mission probe mass will be significantly smaller than that of client satellites for orbit transfers. The types of missions that can be performed are also limited by the SOTV’s distance from the sun. Any type of planetary mission beyond Mars will require a non-solar thrusting system, due to the decrease in solar energy intensity at orbit radii beyond that of the inner planets. The SOTV is able to boost an interplanetary probe of appropriate mass into an escape velocity from the Earth. This initial Earth-escape boost may be the best option for an interplanetary mission that uses the SOTV as it is currently designed. A constrant-thrust transfer from Earth to any planet other than Venus will require a very small probe of relatively low mass. These figures are based on the following ∆ equation that relates spacecraft mass properties to V and Isp: ∆ − V M I g (A.1) P = 1− e SP 0 + M SOTV M L

MP is the mass of the mass of the propellant, MSOTV is the mass of the SOTV, ML is the mass of the payload ∆ (space probe), V is the total velocity change available from one fully fueled SOTV, the Isp is 750 sec, and

2 g0 is 9.81 m/s . Rearranging this equation yields a probe/payload mass ML as a function of interplanetary ∆V. Note that this total ∆V includes the velocity change required to escape Earth from GEO and the velocity change required to perform an interplanetary constant-thrust transfer. M = P − (A.2) M L ∆ M SOTV − V ⋅ 1− e ISP g0

With MSOTV of 14,000 kg and MP of 11,000 kg, results of this equation are listed in Table A.1 below:

Table A.1: Available probe mass for interplanetary transfer missions

∆ Planet Total V from GEO (km/sec) Available ML (kg) Mercury 22.44 -2,453 Venus 9.61 1,086 Earth 4.36 10,603 Mars 10.01 795 Jupiter 21.08 -2,335 The results show that missions to Mercury and Jupiter are not feasible using the SOTV in its current configuration, with a full tank of hydrogen at GEO. Although the SOTV is capable of transferring a

113 small probe to Venus or Mars, this is also not practical. A probe that has a mass of 1,000 kg or less may be of little scientific value in comparison to the enormous costs associated with using a SOTV to transfer it to another planet. For example, the Cassini mission to Saturn was launched with a mass of 5,600 kg, which indicates a trend toward building and deploying larger, heavier, more complex, and more capable probes into interplanetary trajectories. SOTV can only be used to transfer probes at great cost. In the GEO-based launch scenario, the client probe and an extra SOTV fuel tank would have to be launched to GEO first. The SOTV would then consume a tank of fuel getting to GEO, at which point it would have to be refueled by the waiting tank. It may be appropriate to use the SOTV as an intermediate stage from LEO to GEO in an interplanetary mission, and use a traditional means of propulsion for the transfer to another planet. Even if the probe is light enough and cost is not a concern, communication with the SOTV will also be a problem once it is on a trajectory to another planet. The current communications system is not sufficient to transmit over space and through atmospheric attenuation to the ground station. The communications system must be slightly altered in order for it to use the Deep Space Network (DSN) space communication network. These alterations include frequency crystal changes, communications pointing accuracy, and other necessary modifications so that the SOTV can interface with the DSN. As long as the SOTV is transmitting in the right direction towards Earth, the DSN network will be able to receive telemetry data, send commands, and perform other necessary communications tasks. The bottom line is that an interplanetary mission using the SOTV is not worth the time, effort, or cost in return for the value that could be gained from the mission by using traditional thrust means with a more capable space probe.

114 Appendix B: Communications and Data Handling Hardware

B.1 Communications Hardware Specifications (Condor0)

STANDARD PERFORMANCE SPECIFICATIONS Frequency Range 1.75 to 2.3 GHz Gain Peak, 4 dBiC typ, 2 dBiC 3 dB Beamwidth 145°, Typical 110°, Minimum Power Handling 10W CW,Maximum MODIFIED PERFORMANCE SPECIFICATIONS Frequency Range 1.2 GHz to 2.5 GHz Gain -3 dBiC 3 dB Beamwidth 90°, Typical ENVIRONMENTAL SPECIFICATIONS Temperature -100° C to +100° C Thermal 1.33 x 10 -3 Newton/M 2 (10 -5 torr) at ±100° C PHYSICAL SPECIFICATIONS Size 15.2 cm in Diameter 8.9 cm in Height Mass 0.340 kg Connector SMA Female

Figure B.1: S-Band Quadrifilar Helix Antenna

Figure B.2: S-Band Omni-Directional Antenna

115 B.2 Computer Specifications (Honeywell0)

Processor RHPPC RISC (PowerPC 603e™ liscensed) 210 MIPS (Drhystone) @ 150MHz, 1.4 IPC 16Kbtye each Icache & Dcache L2 cache 512KB, look aside, write through Memory 4MByte SRAM, EDAC 4Mbyte EEPROM, super EDAC 64Kbyte SUROM (PROM) I/O MIL-STD-1553B 2 Synchronous Serial full duplex ports, 12.5Mbps (RS422) 2 UART full duplex ports, 9.6K to 1M BAUD (RS422) 16 pins programmable as interrupt or discretes JTAG (1149.1), COP, RHPPC debug Timers/counters 5, 32-bit general purpose, 4 with 8-bit prescale Form Factor cPCI 6U x 220 (9.187” x 8.661”) with 2 PMC-like slots (74 x 149 mm) Mass 1.0 kg Power 12.5W (nom), 3.3Vdc ± 5% Radiation hardness Natural space

PEC w PCI bus interface RHPPC Control w L2 cache cntl & tag RS-422/485 SIO (2) w 603e Compatible w Drivers Data Memory control w 5 execution units w SRAM w w PROM/EEPROM 16KB I & D caches Addr RS-422 UART (2) w 210 MIPS @ 150MHz w DRAM/SDRAM w Drivers Host I/F w ~2.6 Watt DRAM refresh w EDAC w SECDED 1553 1553 w Super EDAC Transceiver L2 Cache (Optional) w Auto memory scrub Transformer w 512KB, 128K X 72 w 60x bus arbitration w Parity w 1553/1773 controller DIO/IRQs (16) w SSEC 256KX16 SRAM w Syn serial ports (2) w DUART w DMA To PMCs w IEEE1355 Discretes/Interrupts IEEE1394 cable w Timers IEEE1394 bus 60X Bus, 64-bit, 50 MHz w Clocks/Reset Front Panel Data Port COP/JTAG other

PCI-PCI cPCI Bus Bridge 32-bit, 33 MHz

Main Memory EEPROM (OS, BIT) SUROM (boot) w 4MB, 512K x 72 w 4MB, 512K x 84 w 64KB, 64K x 8

w SECDED EDAC w super EDAC w UTMC 32Kx8 Bus PCI Mezzanine Local PCI-PCI w SSEC 512Kx8SRAM w Hitachi 128Kx8x4 CMOS PROM Bridge Redundant (Optional) cPCI bus

116 Appendix C: Request For Proposal

I. OPPORTUNITY DESCRIPTION Several space technology companies and the U.S. Air Force have studied the Solar Orbit Trans- fer Vehicle concept over the past few years (Refs 1–4). The SOTV concept uses a solar-powered upper stage to transfer vehicles from one orbit to an-other. Solar energy is concentrated by a collector and used to heat a massive block of graphite. Hydrogen is passed through the heated graphite block and exhausted through a rocket nozzle to produce thrust. An operational SOTV could revolutionize space missions by drastically reducing the cost of deploying vehicles and by increasing the mass that can be placed into higher orbits or into interplanetary trajectories. While there are many possibilities for the use of SOTVs, this project focuses on the following five mission objectives: A. LEO-to-GEO Transfer: Transfer of communications and other satellites from low-Earth parking orbit to geostationary orbit. B. Spacecraft Recovery: Capture of spacecraft that have been erroneously placed into the wrong orbit, followed by transfer to correct orbit. C. GEO Cleanup: Capture of non-functional spacecraft in GEO and removal to supersynchronous orbit. D. ISS Support: Delivery of large payloads to ISS and transfer of payloads into reentry orbits. E. Interplanetary Mission Deployment: Transfer of interplanetary probes to a suitable escape trajectory. II. PROJECT OBJECTIVE The objective of this project is to produce a complete system design that could serve as the basis for a proposal to deploy a solar orbit transfer vehicle technology demonstration experiment in space. The design must provide, at a minimum, the spacecraft bus, orbit and attitude control systems, a power system, a metrology system, a command and data handling system, and the primary propulsion system. The design must enable the experimenters to demonstrate conclu- sively that SOTVs can be used for the purposes outlined in Section I, and that the system satisfies the requirements given below.

117 III. REQUIREMENTS AND CONSTRAINTS The fundamental requirement is to demonstrate that a single SOTV can be used in a variety of mission applications. The goal is to demonstrate that SOTVs can be used to reduce significantly the propellant needed to transfer a spacecraft from one orbit to another while also increasing the mass that can be transferred. The SOTV must be able to perform Mission A, as well as at least 1 of Missions B, C, or D, described above. The final report must describe the trade studies leading to the chosen missions. Furthermore, the proposal must include a trade study showing how the SOTV design could be used in Mission E. Further requirements are as follows: a) Payload Capability: The system shall be de-signed to transfer a client satellite having a minimum mass of 3000 kg and a maximum mass of 5000 kg from LEO to GEO. b) Mission Trip Time: The system and mission profile should be designed to minimize overall trip time. A maximum trip time for a LEO to GEO transfer of a client satellite shall not exceed 90 days with a goal of 60 days being preferred. The proposal should also identify the requirements and/or the limitations of a fast-track mission of 30 days or less. Note that the overall trip time shall include that portion of the mission dedicated to rendezvous and capture of the client in the case of satellite rescue operation. c) Rendezvous and capture operations shall be performed based upon a pilot-in-the-loop architecture. Satellite release following orbital transfer shall also be conducted semi- autonomously. For a satellite rescue and re-positioning mission, the maximum client satellite structural loading shall not exceed 0.1g. d) Close proximity operations with a client satellite shall be performed using a cold gas propellant system to minimize potential contamination of sensitive instruments, sensors, and data communication structures. e) The overall vehicle shall have an on-sun pointing accuracy of ± 1.0 degree for gross pointing and tracking. The primary concentrator system shall provide fine pointing and tracking to within ± 0.1 degree. f) The solar thermal engine shall provide a minimum of 15 lbs thrust. The maximum bulk average temperature of the solar receiver cavity shall not exceed 2400 K (+ 25 K). g) The overall system shall be designed to launch off either a Boeing Delta-IV Medium (5/4 or smaller) or a Lockheed-Martin Atlas V (522 or smaller). h) The system shall have an overall reliability of greater than 95 %.

118 i) The system lifetime must be a minimum of five years with a design goal of 10 years. This requirement does not apply to storage capacity of the cryogenic hydrogen storage tank. Refuel- ing of the tank or long-term storage of liquid hydrogen via an on-board cryocooler are acceptable alternatives. IV. DATA REQUIREMENTS The final proposal should include the following: a) Identification of the major features of all segments of the mission architecture b) Mission planning and operations c) Launch vehicle selection d) Orbit analysis including transfer requirements, propulsion system, orbit determination e) Attitude control system including attitude sensors and actuators, and performance predictions f) Structural analysis and design, including mass properties, stress analysis, launch vehicle interface, deployment mechanisms, payload interface g) Power system requirements and design, including load, solar arrays, batteries, other required power equipment h) Heat flow analysis and thermal management system design i) Communications link requirements, frequencies, antennas, receivers, transmitters j) Command and data handling system, including processor selection, command and telemetry requirements, data storage k) Cost estimate of production, deployment, and operations l) A detailed schedule of activities for development and deployment of the system m) End-of-life disposal procedures V. REFERENCES 1. “Air Force Signs Boeing to Build $48 Million Solar Rocket,” http://www.xs4all.nl/~carlkop/sotv.html 2. “Solar Orbit Transfer Vehicle,” http://www.boeing.com/defense-space/space/propul/SOTV.html 3. “Space Power and Propulsion,” http://www.bwxt.com/Products/sotv.html 4. First Conference on Solar Orbit Transfer Vehicles, http://www-chne.unm.edu/isnps/archive/14call/b.html

119 Appendix D: Solar Orbit Transfer Vehicle Design Summary

System Mass: 14,400 kg (fully fueled), 3,400 kg dry Hardware: Propulsion: Solar Thermal Power: Solar arrays/batteries Attitude Control: Control Mo- ment Gyroscopes Attitude Determination: Star trackers Communications: Parabolic high gain, S-band qudrifiler helix antenna, S-band omni-directional antenna Performance:

Payload capacity: 4500 kg maximum LEO-GEO (300 km alt. to 35,600 km); 5,000 kg design maximum Propellant capacity: 11,000 kg Maximum Slew Rate: 1° per second

120 Attitude Determination Hardware: 3 Ball Aerospace CT-633 Stellar Attitude Sensors Power Required: 10 Watts (each) Mass: 3.5 kg (each) Performance: 40 arcsecond accuracy

Attitude Control Hardware: 4 Honeywell M1300 Control Moment Gyroscopes Power Required: 16 Watts standby (each) 50 Watts quiescent (each) Mass: 122.4 kg (each) Performance: Torque Output: 3170 N-m Single gimbaled; 360° rotational freedom about gimbal axis

121 Propulsion Hardware: Receiver Absorber Exchanger Graphite thermal energy storage 2 Silver coated Rhenium secondary collectors Rhenium casting and support structure Power Required: - Mass: 225 kg (150 kg of thermal energy storage) Performance: Specific Impulse: 720-750 seconds Propellant Mass Flow Rate: 0.01 kg/second Maximum Thrust: 80 N

Solar Collector System Hardware: 2 Moog Type 2-2 Solar Array Drive Assemblies 2 Inflatable lenticular resin- impregnated mylar collectors Aluminized mylar coating for mirror surface Power Required: 20 Watts for each drive assembly Mass: 120 kg for both collectors 3 kg for each drive assembly Performance: Drive Assembly Accuracy: +/- 0.08°

122 Command and Data Handling Hardware: Honeywell RHPPC Single Board Computer Power Required: 12.5 Watts Mass: 1 kg Performance: Processor: 20 MIPS Memory: 32 Mbytes DRAM Communications Hardware: Parabolic KU band antenna, S-band qudrifiler helix antenna, S-band omni-directional antenna, 2 Cameras Mass: 1 kg Power Required: 25 Watts (35 Watts with cameras on)

123 Grappling System Hardware: 2 International Submarine Engineering, Ltd. Magnum 7 Robotic Arms Power Required: 100 Watts (each) Mass: 70 kg (each) Performance: Range of Interface: 7.5 cm Length: 1.5 m Accuracy: +/- 13mm

124 Propellant Storage Hardware: Aluminum-Lithium Alloy 1460 Tank Mass: 1,840 kg (empty) Performance: Propellant Storage: 11,000 kg Maximum Internal Pressure: 5 atm Propellant Storage Pressure: 1 atm

Thermal Energy Management Hardware: MLI and vapor cooled shield for tank Graphite felt/MLI for RAX Mass: 200 kg for tank 18 kg for RAX Performance: Tank: ½ kg per day boil off Block: 3.6 KW aperture loss, 3 KW parosidic loss

125