Cygnus Satellite LLC Design Report

Aman Sharma | John Gehrke | Brandon Keeber

Eduardo Asuaje | Jacob Korinko | Vaibhav Menon

Ira A. Fulton Schools of Engineering

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Contents

List of Figures 4

List of Tables 6

1 Nomenclature 7

2 Introduction 92.1 Executive Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

3 Preliminary Design 103.1 Potential Approaches . . . . . . . . . . . . . . . . . . . . . . . . . . . 103.2 Stakeholders . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.2.1 Mission Objective Stakeholders . . . . . . . . . . . . . . . . . 113.3 Top-Level Requirements . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.3.1 Mission Overview . . . . . . . . . . . . . . . . . . . . . . . . . 113.3.2 Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123.3.3 Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123.3.4 Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123.3.5 Thermal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123.3.6 Attitude Determination and Control (ADCS) . . . . . . . . . 123.3.7 Telemetry, Tracking, and Control (TT&C) . . . . . . . . . . . 123.3.8 Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123.3.9 Command and Data Handling (C&DH) . . . . . . . . . . . . . 123.3.10 Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133.3.11 Risk . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4 Mission Analysis 134.1 Concept of Operations . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4.1.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134.1.2 Mission Phases . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4.2 Orbit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164.2.1 Orbit Design . . . . . . . . . . . . . . . . . . . . . . . . . . . 164.2.2 Orbit Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

5 Payload Design 195.1 Gain . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195.2 Beam Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195.3 Antenna Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . 205.4 Frequency Band . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 225.5 Link Budget Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 235.6 Hardware and Bandwidth . . . . . . . . . . . . . . . . . . . . . . . . 25

6 Subsystems 276.1 Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 276.2 Thermal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 326.3 Attitude Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

6.3.1 Star Tracker . . . . . . . . . . . . . . . . . . . . . . . . . . . . 396.3.2 Reaction Wheels . . . . . . . . . . . . . . . . . . . . . . . . . 40

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6.3.3 Inertial Measurement Unit (IMU) . . . . . . . . . . . . . . . . 416.3.4 Disturbance Torques . . . . . . . . . . . . . . . . . . . . . . . 42

6.4 Telemetry, Tracking, and Command . . . . . . . . . . . . . . . . . . . 446.4.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . 446.4.2 Assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 446.4.3 System Interfacing . . . . . . . . . . . . . . . . . . . . . . . . 446.4.4 Cygnus TT&C . . . . . . . . . . . . . . . . . . . . . . . . . . 46

6.5 Propulsion System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 496.5.1 Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 496.5.2 Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 496.5.3 Apogee Kick Motor . . . . . . . . . . . . . . . . . . . . . . . . 506.5.4 Reaction Control System (RCS) Thrusters . . . . . . . . . . . 516.5.5 Propellant Manifold . . . . . . . . . . . . . . . . . . . . . . . 52

6.6 Power System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

7 Risk and Cost Analysis 627.1 Risk and Reliability Analysis . . . . . . . . . . . . . . . . . . . . . . . 627.2 Cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64

8 Gallery 66

References 79

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List of Figures

1 Cost Summary of Satellite . . . . . . . . . . . . . . . . . . . . . . . . 92 Power Summary of Satellite . . . . . . . . . . . . . . . . . . . . . . . 103 Concept of Operations . . . . . . . . . . . . . . . . . . . . . . . . . . 154 Orbit Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175 LEO . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 186 GTO . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 187 Launch Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . 198 Spot Beam vs. Broad Beam[12] . . . . . . . . . . . . . . . . . . . . . 209 Parabolic Reflector Dish . . . . . . . . . . . . . . . . . . . . . . . . . 2110 Shaped Reflector Antenna . . . . . . . . . . . . . . . . . . . . . . . . 2111 Range of Frequency Bands . . . . . . . . . . . . . . . . . . . . . . . . 2212 Cost vs. Availability . . . . . . . . . . . . . . . . . . . . . . . . . . . 2313 Link Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2414 Amplifier Trade Study[23] . . . . . . . . . . . . . . . . . . . . . . . . 2515 TWTA Amplifier [10] . . . . . . . . . . . . . . . . . . . . . . . . . . . 2616 Polarization Architecture . . . . . . . . . . . . . . . . . . . . . . . . . 2617 Static Stress Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 2818 Deformation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2919 Side View of Satellite Structure . . . . . . . . . . . . . . . . . . . . . 3020 Structure Mass Summary . . . . . . . . . . . . . . . . . . . . . . . . . 3121 Temperature Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3222 Temperature Variation for Satellite’s Antennas . . . . . . . . . . . . . 3323 Temperature Variation for Satellite’s Solar Arrays . . . . . . . . . . . 3424 Temperature Variation for Satellite’s Main Structure . . . . . . . . . 3525 Temperature Variation for Satellite’s Batteries . . . . . . . . . . . . . 3626 Temperature Variation for Satellites Transponders . . . . . . . . . . . 3727 Temperature Variation for Satellite’s Propellant Tanks . . . . . . . . 3828 ADCS Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3929 Trade Study between Attitude Determination Systems . . . . . . . . 3930 Jena-Optronik Astro APS . . . . . . . . . . . . . . . . . . . . . . . . 4031 Honeywell HR12s Reaction Wheel . . . . . . . . . . . . . . . . . . . . 4132 ASTRIX 1090 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4133 Angle Perturbation due to Solar Torque . . . . . . . . . . . . . . . . 4234 Wheel Properties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4335 System Interfaces [23] . . . . . . . . . . . . . . . . . . . . . . . . . . . 4536 TT&C Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . 4637 TT&C Block Diagram [23] . . . . . . . . . . . . . . . . . . . . . . . . 4738 TT&C Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4739 TT&C Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4840 Propulsion System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5041 Apogee Kick Motor (AKM) . . . . . . . . . . . . . . . . . . . . . . . 5042 Apogee Kick Motor Specs[2] . . . . . . . . . . . . . . . . . . . . . . . 5143 RCS Thrusters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5144 RCS Thruster Specs[3] . . . . . . . . . . . . . . . . . . . . . . . . . . 5245 Propellant Manifold Block Diagram . . . . . . . . . . . . . . . . . . . 5346 Propulsion Subsystem Weight and Power . . . . . . . . . . . . . . . . 5447 Sum of all Subsystem Power Requirements . . . . . . . . . . . . . . . 55

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48 Candidate Solar Cell Properties[23] . . . . . . . . . . . . . . . . . . . 5649 Triple Junction GaAs Cell [26] . . . . . . . . . . . . . . . . . . . . . . 5750 Solar Array Exploded View . . . . . . . . . . . . . . . . . . . . . . . 5751 Calculation of Solar Array . . . . . . . . . . . . . . . . . . . . . . . . 5752 Single DOF Solar Gimbal Motor [27] . . . . . . . . . . . . . . . . . . 5853 PCDU [28] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5954 Depth of Discharge vs. Cycle [23] . . . . . . . . . . . . . . . . . . . . 6055 Trade Study of Battery Characteristics . . . . . . . . . . . . . . . . . 6056 Lithium Cobalt Oxide Battery Unit [18] . . . . . . . . . . . . . . . . 6057 Cost of Solar Cell Technologies [23] . . . . . . . . . . . . . . . . . . . 6158 Cost Calculation of Carbon Fiber Cloth [15] . . . . . . . . . . . . . . 6159 Cost of Various Types of Batteries . . . . . . . . . . . . . . . . . . . 6160 Battery Capacity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6161 Failure Graphs [8] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6262 Probability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6263 Impact . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6364 Impact . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6365 Impact . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6366 USCMB Non-Recurring . . . . . . . . . . . . . . . . . . . . . . . . . 6467 USCMB Recurring . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6568 Application of USCMB . . . . . . . . . . . . . . . . . . . . . . . . . . 6569 Side Exploded View . . . . . . . . . . . . . . . . . . . . . . . . . . . 6670 Isometric Packed Payload . . . . . . . . . . . . . . . . . . . . . . . . 6671 Side Packed Payload . . . . . . . . . . . . . . . . . . . . . . . . . . . 6772 Space Side Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . 6773 Communications Compartment . . . . . . . . . . . . . . . . . . . . . 6874 Power Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . 6875 Packed Upper Compartment . . . . . . . . . . . . . . . . . . . . . . . 6976 Emergency and Launch Antenna . . . . . . . . . . . . . . . . . . . . 6977 Battery Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7078 Antenna Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7079 Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7180 Propulsion Compartment . . . . . . . . . . . . . . . . . . . . . . . . . 7181 Transmitter & Receiver Antennae . . . . . . . . . . . . . . . . . . . . 7282 Thrust and Interlock ESPA Manifold . . . . . . . . . . . . . . . . . . 7283 Gyroscope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7384 Input/Output Multiplexer . . . . . . . . . . . . . . . . . . . . . . . . 7385 Power Condition and Distribution Unit . . . . . . . . . . . . . . . . . 7386 RCS Thruster . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7487 Reaction Wheels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7488 Star Tracker . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7589 TT&C Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7590 Emergency Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . 7691 Battery Array . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7692 Apogee Kick Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7793 Solar Array Gimbal . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7794 Transponder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7895 Travelling Wave Tube Amplifier Array . . . . . . . . . . . . . . . . . 78

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List of Tables

1 Frequency Band Trade Study . . . . . . . . . . . . . . . . . . . . . . 22

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1 Nomenclature

Aeff = Effective Areaλ = Wavelengthη = EfficiencyD = Diameter of AntennaPsa = Solar Array Power Output RequirementPe = Spacecraft Power Requirement during EclipsePd = Spacecraft Power Requirement during DaylightTe = Eclipse PeriodTd = Daylight PeriodXe = Power System Path Efficiency during EclipseXd = Power System Path Efficiency during DaylightPo = Solar Array Power Output at Beginning of LifePBOL = Solar Array Power Output at Beginning of LifePEOL = Solar Array Power Output at End of LifeId = Inherent DegradationI = CurrentR = ResistanceVbus = Selected Bus VoltageDOD = Depth of Dischargeηb = Efficiency Between Battery and Loadηcell = Efficiency of Gallium Arsenide Solar CellNbat = Number of Batteries RequiredCbat = Battery Energy Capacity RequiredCactual = Actual Capacity of Selected BatteryUe = Exhaust SpeedIsp = Specific Impulsego = Gravitational acceleration of Earth at Sea LevelMp = Propellant massOF = Oxidizer to Fuel RatioMox = Oxidizer MassMfu = Fuel MassVp = Propellant Tank Volumeρp = density of propellant at 323 Kρox = density of oxidizer at 323 Kρfu = density of fuel at 323 KMT = Tank MassMgas = Mass of Pressurant GasP = End of Life Tank PressureRgas = Specific gas constant of Pressurant GasT = End of Life Tank Temperatureρ = Density of Pressurant GasVpres = Volume of pressurant gasTF = Thrust∆V = Change in velocitymo = Beginning of Life Mass of Satellitemf = End of Life Mass of Satellitefn = Fundamental Frequency

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do = Outer Diameterdi = Inner DiameterE = Modulus of ElasticityLs = Structure LengthMs = Structure MassIA = Area Moment of Inertia

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2 Introduction

The objective of this mission was to design, analyze, and build a direct broadcastsatellite (DBS) for a primary customer. The satellite had to be capable of providingbroadcast television to specific end users situated anywhere within the contiguousUnited States. It was necessary to design robust electronic and mechanical compo-nents capable of surviving the harsh space environment for a designated number ofyears. Additionally and most importantly, the satellite needed to provide dependableand consistent service to the end user in order to remain a competitive consumer op-tion. This report was written from the perspective of Cygnus Satellite LLC, taskedwith the responsibility of performing all necessary trade studies and analyses regard-ing orbit and payloads, in addition to cost, risk, and schedule analysis. By utilizingcomputer software such as, SolidWorks, Thermica, MATLAB etc., the company wasable to design a sophisticated satellite and assemble a comprehensive report detailingsubsystem configurations and other essential pertinent information.

2.1 Executive Overview

Figure 1: Cost Summary of Satellite

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Figure 2: Power Summary of Satellite

3 Preliminary Design

3.1 Potential Approaches

The first step of the initial analysis was to decide the functionality of the satellite.This depended on the stakeholders of the satellite and the consumer market. The firstoption was to create a satellite to broadcast internet. These satellites are typicallyput into a geosynchronous orbit in a constellation configuration. This orbit allowsfor greater coverage but it also increases the signals latency to roughly 20 times thatof a terrestrial internet network. Operating at this orbit also significantly increasesthe delay to receive the data. This data speed can be on the order of two magnitudesslower than current high speed internet services. To overcome this problem a largelow earth orbit constellation would be needed to provide coverage to the entire UnitedStates. Creating a constellation like this would be expensive and not practical forthis mission.

The next option was a telephone satellite. Similar to satellites that provideinternet access, a large constellation at LEO or GEO would be needed to providecomplete coverage across the country. This type of satellite also required a significantamount of ground based stations and other infrastructure to work properly. Inaddition, there was a significant amount of competition already in this space withlarge constellations that provide coverage to the entire planet.

The last option was to broadcast television. There are two types of televisionsatellites, fixed satellite service (FSS) and direct broadcast service (DBS). FSS pro-vides service either directly to a home satellite dish or to a ground station. Thistechnology has been in use since the 1970s and has since been replaced by DBS. FSSsatellites require much larger satellite dishes than their DBS counterpart and areless widely used by satellite television providers. In some cases, television providersreceive data using a fixed satellite service but then re-broadcast the data using adirect broadcast satellite. Based on this, a direct broadcast satellite was the bestoption.

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Currently, most satellite television providers use direct broadcast satellites. DBSrequires a significantly smaller satellite dish than FSS which makes it easier for theconsumer to have at their home. This cuts down on infrastructure costs of havingan Earth based station and increases access to consumers. Direct broadcast satel-lites currently operate on the Ku band but there have been experimental satellitesprovided by NASA and DirecTV that operate on the Ka band. This research andadvancement could open up a wide range of bandwidths in which to broadcast datain the future. Recently, there have been advancements in providing mobile receptionfor airlines, recreational vehicles and multiple military applications. These marketsprovide potential markets for future company growth. Overall, the design of a directbroadcast satellite was the best option due to reduced infrastructure needed, a wideunsaturated consumer market, and the potential for future company expansion.

3.2 Stakeholders

The stakeholders for a communication satellite include the primary customer, sec-ondary customer, operator, and end user. In most cases, the primary customer is thestakeholder that finances and owns the communication satellite, and facilitates thetransmission of data to it. The secondary customer is a stakeholder that may havefinanced or launched the communication satellite, and benefits from it. The operatoris responsible for overseeing most or all functions of the communication satellite, andis held responsible if the satellite fails. The operator is therefore also considered astakeholder. The end user is the final stakeholder because they receive the data fromthe primary customer.

3.2.1 Mission Objective Stakeholders

In this paper, a scenario is considered in which DirecTV has requested Cygnus Satel-lite LLC to build a state of the art communication satellite that is capable of receivingand transmitting television broadcast to its customers in the 48 contiguous UnitedStates. While DirecTV will need Cygnus to operate the satellite in geosynchronousorbit (GEO), a third party will be employed to provide launch services.

For this scenario, the stakeholders will include DirecTV which is the Primarycustomer, with Cygnus Satellite LLC acting as the Operator. For the secondaryuser, SpaceX has been selected to act as the launch provider, and television-watchingcustomers will play the role of end users.

3.3 Top-Level Requirements

3.3.1 Mission Overview

The satellite shall provide 552 stations of uninterrupted High-Definition Television(HDTV) via direct broadcast downlink to customers in the Contiguous United States(CONUS). The mission duration shall not exceed 15 years from final orbit acquisition.All electronic systems aboard the satellite shall be radiation hardened and have thecapability to function fully in the space environment for the entirety of the missionduration.

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3.3.2 Sizing

In terms of physical restraints, the satellite shall not exceed the allowable massdetermined by the launch vehicle capabilities. The dimensions of the satellite shallalso not exceed the dimensional constraints determined by the launch vehicle fairinggeometry.

3.3.3 Orbit

The mission orbit for the satellite shall be kept at a geosynchronous equatorial orbit.The inclination of the satellite shall be 0 degrees with an allowable error of plus/minus0.05 degrees, measured from the center of Earth. Longitudinal drift shall not exceedplus/minus 0.05 degrees. Disposal of the satellite at EOL shall consist of a graveyardre-orbiting maneuver of at least 300 km.

3.3.4 Structure

The structure of the satellite shall be capable of withstanding all loadings duringlaunch, deployment, and nominal operations.

3.3.5 Thermal

The thermal subsystem of the satellite shall be able to regulate and collect/transmitdata for the thermal environment aboard the satellite. All subsystems shall be keptwithin their respective operating limits.

3.3.6 Attitude Determination and Control (ADCS)

The attitude of the satellite shall be determined and controlled by a three-axis stabi-lization system consisting of reaction wheels and RCS thrusters. A pointing accuracyof a TBD amount shall be set to meet communications payload and TT&C require-ments.

3.3.7 Telemetry, Tracking, and Control (TT&C)

Telemetry, tracking, and control of the satellite shall be handled via a single groundstation during nominal operations, accessible 24/7. TT&C during launch operationsshall be handled by a capable third-party satellite tracking network.

3.3.8 Propulsion

The satellite propulsion system shall be capable of delivering the spacecraft to missionorbit and be able to maintain this station for the entirety of the mission duration.

3.3.9 Command and Data Handling (C&DH)

The spacecraft computer shall have the capability to collect, interpret, and transmitall necessary data aboard the satellite.

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3.3.10 Power

The power subsystem shall be able to provide constant power needed to keep thesatellite active during both nominal operations and eclipse periods. Onboard bat-teries shall be capable of sustaining satellite systems through the duration of themission without exceeding an acceptable depth of discharge. Solar panels shall becapable of recharging the batteries and supplying power to the satellite during non-eclipse periods. The choice for solar panels shall be made with the consideration ofdegradation in such a way that the EOL power does not fall below acceptable levelsat any point during the mission.

3.3.11 Risk

Risk shall be mitigated effectively in all aspects of the satellite system. Factors ofsafety, budget limits, and redundancy shall be investigated in order to minimize thepossibility of system or mission failure.

4 Mission Analysis

4.1 Concept of Operations

4.1.1 Overview

The Cygnus Satellite Concept of Operations consists of four mission phases. Pre-launch operations includes all functions up to the actual launch of the satellite system.Launch and deployment of the satellite system includes all functions and operationsup to the deployment of the system to geosynchronous equatorial orbit (GEO). Theon-orbit portion is by far the most important part of the mission, as it includes allnominal operations of the system, and is crucial to achieving the mission objective.Finally, the end-of-life (EOL) phase of the mission is a small but necessary part ofthe operations, and includes the steps that will be taken to ensure compliance withall mandates regarding spacecraft disposal.

4.1.2 Mission Phases

4.1.2.1 Pre-Launch Operations The Cygnus Satellite system shall be assem-bled and shipped in such a way that it will require minimal handling and mainte-nance prior to launch. The batteries will be charged and a check of all subsystemfunctionality must be performed. Solar panels should be folded into their launch con-figurations, and the explosive bolt system for the deployment of the solar panels willbe primed, with the torsional springs already installed and tensioned. The satellitewill be attached to the launch vehicle fairing, along with six other small cube-sat sys-tems from groups that will share the launch cost. The launch vehicle payload fairingwill be attached to the main body and prepared for launch. All Cygnus subsystemswill be inert during launch.

4.1.2.2 Launch and Deployment The spacecraft will be deployed into geosyn-chronous transfer orbit (GTO) by the launch vehicle via an ejection system integratedinto the launch vehicle coupler. During this time, a third party spacecraft trackingnetwork will be tracking the system. At the time of Cygnus deployment, the six

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cube-sats will also deploy and carry out their various missions separately; the pay-load fairing will separate, and the solar panels will deploy. When the system detectsseparation from the launch vehicle, all TT&C components will come online, and theADCS system will autonomously attempt to orient the antennas to open groundstation communication. Once a stable orientation is found and the satellite is ableto establish TT&C communication, the system will be ready to initiate the apogeekick maneuver to circularize and obtain the required mission orbit. Following theapogee kick maneuver, the system will be ready to proceed with on-orbit operations.

4.1.2.3 On-Orbit Operations The Cygnus Satellite payload will begin nom-inal operations once antenna pointing is verified. The ground station will uplinkdata to the satellite where it is processed and then transmitted to the user (down-link). During this time, the ADCS and TT&C systems will be working to ensureproper attitude, altitude, and health of the spacecraft. The power subsystem will en-sure power distribution throughout the life of the satellite, including solar array useduring nominal operations and battery use during eclipse periods, and the thermalsubsystem will regulate the thermal environment during these periods.

4.1.2.4 End of Life Operations After 15 years of nominal operations, the satel-lite system will be re-orbited to a graveyard orbit. Once the satellite is re-orbited,Cygnus will end operations with this specific spacecraft.

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Figure 3: Concept of Operations

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4.2 Orbit

4.2.1 Orbit Design

In choosing which type of orbit to place a communications satellite (comsat) in, anumber of factors needed to be considered in light of the desired mission objectives.Factors such as inclination, altitude, eccentricity, perigee, and ascending node allinterdependently affect the structure of the comsat program and must be consideredwhen making all major design decisions.

Deciding the basic mission profile is a direct antecedent to the orbit design. Analternative mission profile to the one assigned was considered in which a comsat wouldbe placed in orbit around the Earth-Moon L2 Lagrange point, and act as a targetedrepeater station for high-bandwidth signals between the Earth and a hypotheticallunar colony. The trajectory necessary to place such a satellite in the required orbitwould be highly advanced, consisting of a unique parabolic Earth escape-trajectoryand a complex series course alterations. Based on the highly time-limited nature ofthis class, this mission will not be attempted at this time.

In the case of a communication satellite providing HD TV coverage to the con-tiguous United States, orbit type can be either Low Earth Orbit (LEO), MEO, GEO,or Elliptical. Each orbit has distinct advantages and disadvantages. In the case of aLEO or MEO orbit, the decreased distance between the ground stations and satelliteis ideal for applications with a low tolerance for signal delay. Decreased signal lagwould of course also increase the efficiency of an HD TV broadcast, but due to themodern practice of introducing a censor delay into TV broadcasts, it would be ratherunnecessary. In addition, by placing a comsat in LEO, the decreased Earth surfacecovered per satellite necessitates that a constellation (fleet) of satellites be employedin order to ensure continuous coverage, a vastly more expensive and high-risk option.

For these reasons, the most cost-effective and efficient orbit to place Cygnus intowould be a geostationary one. One positive implication this has on the missionprofile is a reduced orbital maintenance cost, reducing the need for large amounts ofmaneuvering propellant. At the altitude required for GEO ( 35,000 km), atmosphericdensity nears that of a perfect vacuum, imparting a negligible drag force, even overlong periods of time. The 0◦ latitude of the GEOs equatorial orbit minimizes theJ2 anomalys perturbing effects on the satellite, leaving only solar radiation pressureand 3-body perturbations to account for.

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Figure 4: Orbit Trade Study

4.2.2 Orbit Analysis

Presented in fig. 5. are the ∆V estimates for the satellite transfer from a parkinglow Earth orbit of 555.6km to a geosynchronous orbit of 35,789km. In order tomake the decision for a final launch site we investigated six different launch locationsaround the world. The initial mass of the satellite is assumed to be 2669.0 kg with theapogee kick motor operating at an Isp of 315 sec. For the trade study presented in thefollowing figures an analysis was conducted using several different launch locationsat various inclinations. The selection of the launch site depends upon the launchprovider and the amount of propellant and ∆V that is needed.

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Figure 5: LEO

The second option is to use the launch provider to place the satellite into geosyn-chronous transfer orbit (GTO) where we would then only be responsible for con-ducting the final burn once the satellite reaches the desired orbit and to make theinclination plane change. This option is more expensive in terms of cost for thelaunch but it provides many other benefits. Using a GTO provided by the launchvehicle we can use a significantly less amount of fuel which will reduce cost andweight of the satellite. In addition, this method is very reliable. Outlined in Figure 6is the propellant mass estimate and the ∆V required for the transfer.

Figure 6: GTO

In both of the figures above, the ∆V estimates are based on a combined secondburn and inclination change in one maneuver. One option to change the inclinationof the satellites orbit is to conduct the first burn to place the satellite in the desiredorbit and then conduct another burn to place it into the right inclination of 0 degrees.The other option is combine these two maneuvers into a single burn. This techniqueproves to be more efficient by using less ∆V which in turn reduces the amount ofpropellant that is needed.

Based upon these results several conclusions can be made. For the most efficienttransfer, based upon the amount of ∆V required, using a launch location such asFrench Guiana and India would be the most beneficial. However this provides severalchallenges that must be overcome to launch from these locations. Due to the locationof these sites it would be very expensive and would contain a large amount of riskshipping the completed satellite to these locations. Launching the satellite fromthe Cosmodrome in Kazakhstan would not be a good option because of its largeinherent inclination and its distance from the United States. The final two optionsare to launch from Vandenberg California or Cape Canaveral Florida. These twolocations provide very similar characteristics and must be decided based upon thefacility the launch provider uses.

Based upon the mass requirements and finding a launch provider that can placethe satellite into a geosynchronous transfer orbit, SpaceX’s Falcon 9 launch vehicle

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will be used. This launch meets every criteria necessary to place the satellite intothe desired orbit. The launch will take place from Cape Canaveral Florida with thefinal characteristics outlined in fig. 7.

Figure 7: Launch Characteristics

5 Payload Design

5.1 Gain

Gain is the measure of directivity of an antenna. Gain is proportional to effective areagiven by eq. (1) and eq. (2). For large antennas, the effective area is approximatelyequal to the real area of the antenna.

G =4 ∗ π ∗ Aeff

λ2(1)

G = η ∗ π ∗Dλ

2

(2)

Standard antenna efficiency (η) is usually between 55% to 70% and the standardground satellite antenna diameter is 0.5334 m but can range in size depending on theneed [5]. Typical LNB noise of satellite antenna is 1 dB, and as the EIRP increases,the area of the antenna decreases [11]. It is important to know that EIRP helpsshape the coverage area when designing the antenna. At a bandwidth of 6 GHz,the gain of a 10 m antenna will approximately equal 53.3 dB [20]. Also, for thecommunication satellite to be most efficient, it is best to have an EIRP no less than40 dBW.

5.2 Beam Trade Study

According to Tech-FAQ [9], a spot beam is a signal that is directed towards a specificarea on the surface. The advantage of using spot beams is that is allows a satelliteto target a specific area, which averts data interception and minimizes the powerutilized. Another advantage of spot beams is the capability to reuse a frequencyfor different locations without interference at the receiver. This allows for morechannels to be carried on the same frequency which is then operated in several areas.However, using spot beams to cover too many areas such as the entire continentalUnited States is not recommended because it take up too much power and may causedata interference because the beams are grouped closely.

On the other hand, wide beams cover large geographical areas and a wide beamover the continental United States is also known as CONUS [12]. An advantage ofwide beam is that it is more omni-directional than spot beams. This means that

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Figure 8: Spot Beam vs. Broad Beam[12]

an antenna does not need to be pointed accurately in order to create a connection.CONUS is also much simpler and more reliable than spot beams.

Cygnus Satellite LLC. will use CONUS for its communication satellite, whichmeans that it is necessary to choose between installing parabolic reflector dish andshaped reflector antenna.

5.3 Antenna Trade Study

In order to use a parabolic reflector dish for CONUS it would be essential to adjustthe pointing and operating point of the reflector so that the gain at the edges of thecoverage area is within the requirements [23]. However, this means that gain overmost of the coverage will be greater than the minimum requirement, which requires alot of power. Another disadvantage of utilizing a parabolic reflector dish for a broadbeam is that covers areas that are not in the continental United States. Non-essentialcoverage areas include oceans, Canada, or Mexico.

The above mentioned disadvantages of utilizing a parabolic reflector dish forbroad beam can be fixed creating a shaped reflector antenna specific to cover CONUS.Shaped reflector antenna produces a broad beam that conforms more closely tothe coverage area by limiting transmitted power, which is why a a shaped reflectorantenna is better. Examples of the parabolic reflector dish and shaped reflectorantenna are observed in fig. 9 and fig. 10, respectively.

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Figure 9: Parabolic Reflector Dish

Figure 10: Shaped Reflector Antenna

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5.4 Frequency Band

Although Ku-band systems are more abundantly used today, Ka-band systems arean up and coming competitors. With both systems offering certain advantages, it isbeneficial to compare the two bands in order to ensure an optimal design. The figurebelow [7] shows the ranges of both the Ku and the Ka-band along with the effectsof rain on signal dissipation. The table and figures below show a more in-depth

Figure 11: Range of Frequency Bands

comparison of Ku-band systems and Ka-band systems. After being introduced inthe 1980s, the coverage of Ku-band systems has significantly increased over the past30 years [16]. However, this has resulted in less carrier frequencies being available fornew systems. The new Ka-band systems are able to offer higher downlink data ratesbut fall short in regions with high rain weather [1]. As can be seen in the figuresbelow, for harsh regions, cost for Ka-band systems drive-up exponentially which cancause problems when providing service to areas in CONUS where this type of weatheris common. Even in temperate regions the cost drives up exponentially as higheravailability is demanded. In conclusion, Ku-band systems are superior due to beinga well-established system and due to its higher overall reliability.

Criteria \Band Ku KaCost per BPS (bitsper second)

Offer competitive cost per BPScompared to same spot beam sizeKa-band systems

Provide same cost per BPS forsmaller spot beam systems. Linkperformance deteriorates as spotbeam coverage increases

Coverage Provide same coverage as Ku-band large spot beams. EIRP isthe same for both bands, but Ku-band systems have high signalgain. Frequency reuse increasescoverage

Ka-band small spot beams pro-vide significantly less coveragethan Ku-band systems. Lowersignal gain for similar size anten-nas. Lack of frequency reuse lim-its coverage

In Case of SystemFailure

The existence of a large-numberof Ku-band satellites allows forreallocation of service to othersatellites

The scarcity of Ka-band satel-lites denies Ka-band systems thesame benefits as the Ku-bandsystems

Weather Less energy dissipation due torain. Less overall cost vs avail-ability in most regions (see fig-ures below)

High signal loss due to rain. Costare escalated as higher availabil-ity is demanded (see figures be-low)

Table 1: Frequency Band Trade Study

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(a) Temperate

(b) Tropical

Figure 12: Cost vs. Availability

5.5 Link Budget Analysis

In order to conduct the link budget analysis several parameters and assumptionsneeded to be made. In order to accommodate the high bandwidth nature of receivinga HD television broadcast, a home satellite dish size of 0.5 meters was used. Thissmall home dish was deemed a reasonable sized product for consumers to be expected

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to buy. In addition, the receiving and transmitting dishes on the satellite were sizedto be 1.5 meters in diameter each, with a gateway dish size of 10 meters in diameter.All dishes were sized to allow for a greater link margin, which helped account for anyattenuations due to signal losses or inclement weather conditions. Beyond bandwidthconsiderations, the home dish size was influenced by the volatile nature of weather-induced signal attenuations, which if made too small would prevent clear broadcastat the specified data rate. The link budget analysis can be seen in fig. 13

Figure 13: Link Budget

In order to conduct the analysis, assumptions also had to be made about the lossesinvolved in the system, which were assumed to be worst case scenario values [23].Incorporated losses included those due to the antenna, circular depolarization losses,atmospheric losses, pointing losses, and freespace losses from a geosynchronous orbit.In addition, each antenna was assumed to operate at an efficiency of 70% for thesatellite antennas and the worst case efficiency for the uplink ground antenna and thehome receiver of 55%. For further investigation into the equations and assumptionsmade, refer to the source code in the appendix section.

These results from the analysis are very reasonable when compared to othercommunication satellites. From this analysis we can see that we have a link marginof 14.1038 dB which provides an additional tolerance for any attenuations betweenthe transmitter and the receiver. This high link margin allows for indirect signalsto be able to bounce off of any other surfaces and still be received by the grounddish. To some companies this link margin may be considered to be too large but wewant to ensure that the link is made between the satellite and the user under anycircumstance.

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5.6 Hardware and Bandwidth

In satellites, back-end amplifiers are used to increase the strength of processed signals.Signals received through the antenna are filtered from the array, and passed throughthe transponder with the appropriate bandwidth allocation. The transponder multi-plexes the signal frequency by shifting (translating) its center frequency down to thefrequency necessary for downlink, then passes the signal through a power amplifier,which increases the signal strength. Finally, the signal is passed to the downlinkantenna, where it is transmitted to the end user. Traditionally used in satellites,vacuum tubes amplify signals into the Ultra-High-Frequency (UHF) or even the Mi-crowave frequency range via a generally analogous principle. A series of resonatorsare resonated with the signal of interest, and the energy from an electron beam pass-ing through the resonators is transferred to the signal, increasing its strength. Insolid state amplifiers, this same effect is accomplished through careful manipulationof electron flow via control of the semiconductor pathways within the material itself.There are a number of drawbacks to the use of vacuum tubes in satellites, the mostproblematic of which being an inherently low durability due to the fragile nature oftheir construction. Additionally, amplified signals suffer from non-linear distortionsoutside of a very narrow bandwidth, severely limiting their transmission capabilityand therefore increasing the number of amplifiers/transponder pairs needed. Solid-state devices in general are smaller, more reliable, and cheaper than their tubularancestors, seemingly making them the ideal candidate for future Cygnus satellites.

It is worth noting that a number of experts on the subject still to this dayprefer vacuum tube amplifiers due to a measurably higher quality thermionic energyconversion than can be achieved in solid-state amplifiers. This slight increase inefficiency however is offset by the non-linear wide-band distortions introduced intosignals amplified by vacuum tubes. In order to make the decision between Solid StatePower Amplifiers (SSPA) and Travelling Wave Tube Amplifiers (TWTA), a numberof traits had to be considered (Figure 14). Despite the increased mass, TWTAsexhibit a significant increase in efficiency over SSPAs in applications which require ahigh data rate and satellite altitude. TWTAs are however less durable, bulkier, andintroduce more signal distortion at the ends of their bandwidth range than SSPAs.Despite all of these drawbacks, the extremely limited power available ultimately playsthe largest role, resulting in the decision to use TWTAs in our satellite. Increasedpower consumption, as will later be shown, has a cascading effect on mass, cost, andrisk.

Figure 14: Amplifier Trade Study[23]

In the Ku band, bandwidth allocations are typically 500 MHz wide. In theCygnus satellite being designed, a circular polarization scheme was chosen in order toincrease the amount of stations being broadcast to the end user by utilizing the boththe vertical and horizontal electromagnetic wave components. Research into powersystem aboard the Horizons-1 satellite yielded power and bandwidth information onits TWTA transponders. We have therefore assumed a 36 MHz bandwidth, and108 Watt power consumption per transponder [21]. Transponder mass values were

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attained from the product page of the L-3 transponders which were chosen for thissystem (fig. 15).

Figure 15: TWTA Amplifier [10]

In accordance with SMAD values, an inter-channel gap bandwidth and station-keeping bandwidth of 4 MHz each was assumed. fig. 16 displays the architectureof signal polarization. A Matlab Script incorporating all these values yielded thefollowing results:

• Total number of channels = 23• Total number of stations per channel = 24• Total number of digital TV stations transmittable = 552• Total number of transponders including spares = 30• Total mass of transponders and amplifiers including spares and power convert-

ers = 153.9 [kg]• Total power of transponders and amplifiers including = 2484.0 [W]

Figure 16: Polarization Architecture

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6 Subsystems

6.1 Structure

The space environment is very demanding on materials. The combination of hugeswings in temperature from thermal cycling, exposure to large amounts of radiation,oxidation, outgassing, and large forces during launch make for a difficult materialsproblem. In addition, the structure needs to act as a skeleton for the assembly ofthe subsystems so it must be easily integrated with all other systems. For the costof the structure to remain low, the structure needed to be made out of componentsthat are compatible with current manufacturing processes.

In order to combat the wide range in thermal cycling it was important to use amaterial with a very small thermal expansion. The temperature the spacecraft wouldface can range from -160 C to +180 C which demanded the material to be able towithstand extremely hot and cold temperatures. In addition, electromagnetic andparticle radiation from the radiation belts, solar emissions and other cosmic radiationhad to be taken into effect. However, these effects were considered negligible dueto the small impact they had on the structure. In addition, the material neededto be strong to overcome the accelerations, acoustic effects and thermal effects fromlaunch. The loads at launch proved to be the largest strain on the structure during itsmission lifetime. While the material needed to be strong, almost more importantly,it also needed to be very lightweight to reduce the costs associated with launch.

There were several possible materials that met these requirements. Currently,the majority of spacecraft structures are made out of aluminum alloys. Some alloysmeet most of the requirements as outlined above. They have a high stiffness, mod-erate thermal expansion, moderate cost and are readily available in many forms. Inaddition, their stiffness to density ratio is very high which results in high strengthfor relatively low mass. Titanium also presented another good option because itsurpasses the properties of aluminum in most areas. Titanium is extremely strong,has an even higher stiffness to density ratio than aluminum and also has a moderatethermal expansion coefficient. However, titanium is very expensive, hard to machineand not as readily available as aluminum [6].

In addition to aluminum and titanium there were several other exotic metalsthat could be used, the first option being beryllium. With a stiffness to densityratio much higher than titanium it could be a very light and useful material for thestructure. However, beryllium is very brittle and does not maintain its propertieswell at low temperatures. In addition, the material is very expensive to purchase andextremely expensive to machine due to the toxicity of the dust created during themanufacturing process. Another option was magnesium. This metal has a similarstiffness to density ratio as aluminum but operates poorly at low temperatures.

The final and best option were composite materials. With a stiffness to densityratio far exceeding any metal, composite materials are significantly stronger andlighter than any other options. In addition, composite materials have a negativethermal expansion coefficient making it suitable to operate in the thermal extremesof space. However, rapid temperature changes can cause strains in the materialmeaning there would have to be insulated with another material [13].

There are many different types of composite materials but the best for this ap-plication was carbon fiber. Carbon fiber has high strength and stiffness, fatigueinsensitive, very light and relatively low cost [17]. Currently, there are many carbonfiber manufactures that can provide the structural members that would be necessary

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for the main structure of the satellite. It was therefore decided that the structurewould be made out of carbon fiber reinforced polymer square tubes. The structureconsisted of 1 inch square tubing with a wall thickness of 0.022 inches.

The carbon fiber structure was encapsulated in 5 mm thick aluminum 6061 panelsthat served as additional structural support and a surface for attachments. At thebottom of the satellite a thrust cone was used to house the main apogee kick thruster.This cone consisted of the same carbon fiber as the main structure with a heat-resistant lining to help isolate internal components from the massive amounts ofheat given off by the main apogee burn. The entire structure, including the antennabus structure, the back of the antennas and the back of the solar panels, were thencovered in multi-layer insulation (MLI) for thermal insulation. MLI reduces heatlosses due to thermal radiation by increasing thermal resistance and reducing therate of heat transfer.

A structural analysis of the structure was conducted using the SolidWorks sim-ulation tools. Using the Falcon 9 user guide, a design load factor of 6 was appliedto the center of gravity of the structure [19]. A load factor of 6 was then applied inthe axial direction to account for the worst case scenario the structure will face uponlaunch. The results of the stress analysis and the displacement can be seen in fig. 17and fig. 18 respectively.

Figure 17: Static Stress Analysis

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Figure 18: Deformation

From these results the greatest displacement of 0.6244mm was seen to occur atthe center apex of the structure. This displacement was an acceptable result is it wasinsufficient of causing any problems for the integrity of the structure. In addition,the stress distribution would not have a significant impact on the structure as themaximum stress on the structure was found to be 8.502 ∗ 106 N

m2 . This resultedin a factor of safety margin of 24. This structural analysis demonstrated that thestructure was well within safety margins and would easily withstand any forces itencountered upon launch.

The satellite was designed to connect to the Falcon 9 fairing using the EELVsecondary payload adapter (ESPA). This adapter not only mated the Cygnus satelliteto the fairing but allowed for the addition of up to six small satellites with a maximummass of 180 kg to be attached to the adapter [14]. By including this capability, itallowed for launch costs to be shared with other consumers in order to minimize cost.The ESPA adapter has become a standard in the industry and is a notably reliablesystem. The antenna bus structure was then constructed using carbon fiber, andattached to its motorized antenna deployment mechanism (ATM) made by AirbusDefense and Space. This technology is a proven and reliable deployment system. It iscritical that the ATM does not fail due to the mission critical role that the antennasplay.

In fig. 19 the general configuration of the subsystems inside the satellite arevisible. All of power components for the solar panels as well as the housing for theapogee kick motor were fixed to the the bottom platform. Attached to the secondplatform were the propellant tanks for all of the RCS thrusters and the main apogeekick motor. The top section included all of the other components associated with

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the communication system, TT&C, attitude, and thermal subsystems. At the verybottom of the spacecraft the ESPA adapter was attached to six small satellites. TheESPA will remain attached to the Falcon 9 fairing while the satellite is safely deployedaway from the system. Once the satellite is away from the fairing the cube-sats willdeploy from the fairing, leaving behind the ESPA adapter ring forever attached tothe launch vehicle.

Figure 19: Side View of Satellite Structure

In addition to withstanding static loads, the structure required the ability toretain its stiffness in its launch environment. Stiffness is often measured by thenatural fundamental frequency of a given structure. The fundamental frequencyof the satellite was required to be greater than the launch vehicles fundamentalfrequency to prevent dynamic coupling. Dynamic coupling can result in catastrophicdeconstruction, stemming from amplification of launch vehicle input loads, and is tobe avoided at all costs. A preliminary frequency analysis was done using equationseq. (3) and eq. (4).

IA =π ∗ (d4o − d4i )

64(3)

fn =1

2 ∗ π∗

√3 ∗ E ∗ IMs ∗ L3

s

(4)

The area moment of inertia value was based on the circular ESPA launch vehi-cle interface. This assumption was made because a structure will typically have its

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largest strain energy at the launch vehicle interface site [23]. One difficulty encoun-tered in performing frequency analysis is due to the fact that area moment of inertiavaries throughout the entire structure. Therefore, it was assumed that bending stiff-ness was based upon the launch vehicle interface. One major constraint that dictatedthe fundamental frequency was the length of the payload, due to the cubed lengthterm in eq. (4). The light carbon fiber frame and the large modulus of elasticity didhowever act to mitigate the effects of this mathematical result.

The ESPA had an outer diameter and inner diameter of .9609m and .9394mrespectively, and a modulus of elasticity of 1.75 ∗ 1011. The length of the payload is3m and the mass of the structure is assumed to be 128.94kg. At a payload lengthof 3 meters and 128.94 kilogram, the area moment of inertia was estimated to be.0036216m4, resulting in a final fundamental frequency of 117.618Hz. This valueis a reasonable preliminary estimation for the fundamental frequency as it is largerthan the fundamental frequency of the launch vehicle. However, further frequencyanalysis should be done using current software in order to better simulate the launchenvironment and structural properties. The final mass estimates can be seen infig. 20. The fastener estimate was based on 15% of the dry mass of the structure[23].

Figure 20: Structure Mass Summary

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6.2 Thermal

Satellites in orbit are required to operate in an environment with constant and ex-treme temperature changes. The spacecraft absorbs heat through sunlight, albedo,and planet emitted radiation. It also produces heat internally through power dissi-pation, mainly from electrical components. The thermal subsystem is responsible forsetting and maintaining the temperature range for the satellite and its components.This thermal control is important because all of the components require a specificoperating temperature range. These operating temperatures are maintained by ac-tive systems such as radiators and heaters, and passively by coating components withspecific emissivity and absorptivity properties [4, 23].

The thermal analysis was done using Thermica in order to analyze temperaturechanges undergone by components throughout the satellite. This information wasthen used to design thermal maintenance systems This analysis included a roughmodel of the overall design of the spacecraft. The model included a main struc-ture, antennas, solar arrays, radiators, electronic components, batteries, and thepropellant tanks. Although only two electronic components and three batteries werecreated, their properties and parameters were modeled to represent the overall num-ber of each component. Each of the satellites main components were modeled to itsunique physical and thermal parameters. The physical properties included materi-als, thickness, dimensions, density, and coatings. The thermal properties includedspecific heat, conductivity, emissivity, and absorptivity. The electronic componentssuch as transponders, radiators, and batteries also included their own power dissi-pations which represented the heat they radiated inside the satellite. After runningthe simulation, preliminary results were obtained which provided an estimate of thesatellites maximum and minimum temperatures as well as the temperature gradientsfor the components. These results needed to be compared to the operating tempera-tures of the components, and the design was modified in cases where the temperaturereached temperatures below or above the operating range.

Figure 21: Temperature Range

Figure 21 represents the preliminary results for maximum and the minimum tem-peratures the components experienced during one orbit. The orbit used in the sim-ulation was a geostationary orbit during the spring equinox. This orbit was uniquesince during this time the satellite had to travel through Earth’s shadow. This eclipsehad a duration of 72 minutes and the satellite experienced its lowest overall temper-ature. All the main components fit into two distinct outcomes depending on theirtemperature variations. The antennas, solar arrays, and the main body experienceda wide range of temperatures both positive and negative. The batteries, transpon-ders, and propellant tanks had a more compact temperature difference. The main

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difference between these two outcomes was due to the fact that the components withgreater change did not have active thermal controls since their operating range wasbroader. Under this configuration, all satellite components experienced temperatureranges that fell within their operating limits.

Figure 22: Temperature Variation for Satellite’s Antennas

Figure 22 represents the temperature of both the antennas throughout a 24 hrperiod. From the plots we observed where the maximum and minimum temperaturesoccurred. For the antennas, the minimum temperature of −148.5◦C occurred whenthe satellite was directly in Earth’s shadow. The maximum temperature of 35.6◦Chappened when the backside of the antennas faced the Sun and the satellite was atits point in orbit closest to the Sun.

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(a) Complete

(b) 72 min Eclipse

Figure 23: Temperature Variation for Satellite’s Solar Arrays

Both of these plots demonstrate the temperature range undergone by the solararrays. From the top plot we observed that the six main solar panels experiencedalmost identical temperature distributions. Similarly to the antennas, the lowesttemperature was measured during the eclipse. The bottom plots display the time ittook for the solar arrays to go from their maximum temperature of 48.2◦C to theirlowest, at −179.7◦C. It took 72 minutes, the entire duration of the eclipse, for thesolar panels to cool down and another 72 minutes to heat up again. The solar arraysshowed an almost constant temperature for the rest of their orbit due to the fact thatthey tracked the sun as they orbited earth. As the Sun’s incidence angle changedthroughout the geostationary orbit, the solar panels closest to the main structure ofthe satellite had different temperature profiles than the rest of the panels as a result

34

of the satellite’s shadow.

Figure 24: Temperature Variation for Satellite’s Main Structure

Figure 24 represents the temperature of the six panels that made up the mainbody of the satellite. Just like the antennas and the solar arrays, the lowest temper-ature (−48.4◦C) for all panels was registered during the eclipse phase. In contrastwith other components, the highest temperature for each of the different panels of thesatellite depended on its orientation relative to the sun. There were four panels thatexperienced a peak temperature (∼ 54.4◦C) when facing directly towards the Sun.The other two panels, which faced parallel to the equatorial plane, only achieveda maximum temperature of about 3◦C. This occurred because the incidence anglewith sunlight is 0◦ so most of their heat was absorbed from the internal components.

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Figure 25: Temperature Variation for Satellite’s Batteries

The batteries of a communication satellite have different requisites than the othercomponents. Batteries generate internal heat through power dissipation and generatemost of this heat when they are providing the satellite’s power rather than beingrecharged. Batteries are also only used at very particular points during a mission.They are only required during eclipses, where the solar panels are not able to providethe necessary power for system functions. At most, batteries will be required tooperate for 72 minutes which represent the largest eclipse the satellite will experience.In order to maintain the batteries at an adequate temperature, the batteries werealso modeled as heaters for the propellant tanks.

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(a) Complete

(b) 72 min Eclipse

Figure 26: Temperature Variation for Satellites Transponders

These graphs represent the temperature of the transponders inside the satellite.Although there were only two elements, it was possible to represent all the necessarytransponders. From this plot, it is clear that these components experienced a muchdifferent environment than the rest of the sections. The overall temperature ofthese components remained within a 17◦C range. This was due to the fact thatelectronic systems generated internal heat by power dissipation and were kept insidethe satellite, which provided a certain amount of insulation. Since these elementsalso had a very limited operating temperature range, active thermal controls wererequired. Radiators were used to dissipate excess heat and prevent temperatures fromreaching higher values which could affect the performance of the communicationsystem. Similar to the solar arrays, the transponders experienced their greatestchange during the 72 minutes during which the satellite was in Earth’s shadow.

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Figure 27: Temperature Variation for Satellite’s Propellant Tanks

The propellant tanks represented a greater challenge for the thermal subsystemsince their operating temperatures were very restricted. From the plots we were ableto observe that there weren’t any major changes in temperature. But for certain partsof the orbit, the temperature fell outside the limits. This was not really a problemsince the propulsion system was also used for very small time intervals. This meantthat the satellite could wait until the temperature of the tanks reaches the optimalvalue and then perform any of the attitude control burns it is required. As mentionedin the analysis for the batteries, tanks were kept at moderate temperatures using theheat generated by the batteries.

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6.3 Attitude Control

Due to the presence of disturbances while the satellite was in operation, the orienta-tion, or attitude, of the satellite was perturbed from a desired or optimal location.This was where an attitude control system was implemented in order to ensure thatthe satellite was oriented properly. The attitude was adjusted by using utilizing acombination of star trackers, gyroscopes, momentum wheels, reaction wheels, and re-action control thrusters. In addition, the satellite was capable of being spin-stabilizedor three-axis stabilized each of which required a different combination of the afore-mentioned hardware.

Figure 28: ADCS Summary

For this mission, Cygnus LLC decided to design a satellite that was three-axisstabilized using a combination star trackers, reaction wheels, and thrusters.

6.3.1 Star Tracker

A star tracker utilizes a camera to measure the position of star(s) and is able providethree-axis stabilization using the acquired data. The figure below [22] shows a tradestudy that was performed in order to choose between Star Tracker, Earth Sensors,and Sun Sensors. For this mission, the star tracker utilized was provided by Jena-Optronik.

Figure 29: Trade Study between Attitude Determination Systems

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Figure 30: Jena-Optronik Astro APS

6.3.2 Reaction Wheels

Reaction wheels were utilized in order to change the orientation of the satellite byspinning the wheels along a specified axis. Reaction wheels are also capable ofbeing utilized as momentum wheels by spinning them at a constant angular speedwhich builds up angular momentum and greatly reduces any disturbance torquesthat operating on an axis parallel to the rotational axis of the reaction wheel. Forthis mission, four reaction wheels were utilized to provide accurate control. Theorientation of the wheels was determined by analyzing the results of a study doneby University Putra Malaysia [29]. In the study, three and four reaction wheelorientations were analyzed and the chosen orientation produced the lowest amountof torque. The reaction wheels chosen for this mission were Honeywells HR-12. Theirdesign and orientation is shown below. These wheels have a maximum momentumof 50 N-m-s. The saturation rate was conservatively approximated to be 5 days permomentum unloading. This gave a propellant mass of ∼ 47kg that will be utilizedthroughout the lifetime of the satellite in order to unload momentum. This wascalculated using the saturation rate of the reaction wheels, which is once every 5days, and the duration the thrusters will be fired, which is 1 second [24].

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Figure 31: Honeywell HR12s Reaction Wheel

6.3.3 Inertial Measurement Unit (IMU)

An inertial measurement unit (IMU) combines several different sensors such as ac-celerometer, gyroscope, and magnetometers in order to obtain orientation data andeven data on gravitational forces. The IMU used for this mission was the Airbus De-fence & Spaces Astrix 1090. This unit had a very low rate drift rate 0.01 degrees/hourover one hour and 0.10 degrees/hour till end of life. This helped in providing accu-rate data and reduced software complexity as it had to account for less error. Thisunit was also capable of providing three-axis data. The satellite contained two ofthese units, one as a main unit, and one as a backup. The figure below shows theAstrix 1090.

Figure 32: ASTRIX 1090

41

6.3.4 Disturbance Torques

The Gravity Gradient Torques occur when the center of gravity of a spacecraft isnot aligned with its center of mass. For our satellite, the center of gravity wasonly perturbed in the z-direction by 1.375m. This combined with Ixx and Iyy valuesand a worst case θ value of 45◦ produced a torque that approximately equaled to7.6 ∗ 10−6 N −m.

The Solar Pressure torque occurs due to the momentum in the sunlight that heatsa specific area of the satellite. The effect of the solar pressure torque depends on thetype of material used and the location of the solar radiation pressure. By using areflectance factor of 0.6 and a surface area of 4.5m2, the solar pressure torque wasapproximately 4.51∗10−5N−m. Using a single-axis model applied in MATLAB, thechange in angle, the required wheel torque, and required wheel angular momentumdue to this solar radiation torque were calculated over a period of two days. Theyare shown in fig. 33 and fig. 34.

Figure 33: Angle Perturbation due to Solar Torque

42

(a) Wheel Torque due to Solar Radiation Torque

(b) Wheel Angular Momentum due to Solar Radiation Torque

Figure 34: Wheel Properties

43

6.4 Telemetry, Tracking, and Command

6.4.1 Introduction

The telemetry, tracking, and command (TT&C) subsystem is a vital part of thesatellite system. Nearly all onboard subsystems interface with the TT&C subsystemin some way. Information regarding satellite health, tracking, and performance iscommunicated from the spacecraft to the ground facilities, where they are interpretedand analyzed to ensure that the mission is going as planned. Command functions aregenerated based on telemetry and ranging readings, and are uplinked to the satellitewhere these commands are executed. There are five main subsystem functions ofTT&C [23]:

• Carrier tracking (lock onto the ground station signal)• Command reception and detection (receive the uplink signal and process it)• Telemetry modulation and transmission (accept data from spacecraft systems,

process them, and transmit them)• Ranging (receive, process, and transmit ranging signals to determine the satel-

lites position)• Subsystem operations (process subsystem data, maintain its own health and

status, point the antennas, detect and recover faults)

6.4.2 Assumptions

For both the uplink and downlink analyses during normal operations, it was assumedthat the antennas are pointed perfectly towards each other (boresight pointing).During the launch phase and subsequent transfer, near-perfect pointing for the anti-Earth-facing antenna was assumed for the sake of simplicity. It was also assumedthat no more than one TT&C transponder/antenna will fail. All dimensioning, mass,and power assumptions were made from handbook references. Cygnus assumed singleground station control during nominal satellite operations, and a capable third-partytracking network during launch operations.

6.4.3 System Interfacing

The TT&C subsystem interfaces with every subsystem on the spacecraft with theexception of the propulsion subsystem, and must reliably pass information back andforth. A table displaying this interfacing is displayed below in fig. 35:

44

Figure 35: System Interfaces [23]

45

Therefore the requirements can be outlined as such:

Figure 36: TT&C Requirements

6.4.4 Cygnus TT&C

The Cygnus Satellite system used two-way-coherent transponders compatible with aKu-band ground tracking system. The ground station modulated a pseudo-randomcode onto the command uplink signal (similar to the method used by the Air ForceSatellite Control Network (SGLS)), and the TT&C subsystem receiver retransmitedthe code on the telemetry carrier signal back to the ground station. Based onthe turnaround time of the signal, the Doppler-frequency shift was measured andthe range and range-rate was determined. Based on pointing information from theground system, the satellites azimuth and elevation angles were determined, lead-ing to an accurate determination of the spacecrafts angular position. The TT&Csubsystem architecture will be very similar to the SMAD generic TT&C subsystem.

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Figure 37: TT&C Block Diagram [23]

Two-layer redundancy ensured mission continuation in the event of a singleTT&C transponder failure. Use of a diplexer allowed the use of one antenna forboth transmitting and receiving. While not shown on the block diagram, a low-gainhemispherical omni-directional antenna mounted on the anti-Earth-facing side of thesatellite was used during the launch phase of the mission, and during emergency op-erations. This provided a final third layer of redundancy for the Cygnus TT&Csubsystem. The modulation method used by the TT&C subsystem was BPSK/PMmodulation, where the carrier and data were transmitted at frequencies separated bythe subcarrier frequency. Data rates were taken from the suggested values in SMAD(Table 11-19) [23]. The parameters for the Cygnus TT&C system are as follows:

Figure 38: TT&C Parameters

Sizing for the TT&C subsystem was also performed via the SMAD handbook.Table 11-26 [23] lists typical parameters for TT&C subsystems; Cygnus used a Ku-band communications subsystem for TT&C, and therefore used these parameters:

47

Figure 39: TT&C Sizing

48

6.5 Propulsion System

6.5.1 Selection

Cygnus Satellite LLC will utilize a regulated hybrid propulsion system which willcontain a bipropellant NTO/MMH system for the apogee kick motor and a mono-propellant MMH system for the Reaction Control System (RCS) thrusters. MMH isa more volatile type of hydrazine, and is a capable of igniting without an oxidizer.However, an oxidizer such as NTO is utilized in addition to increase thrust. A cata-lyst is necessary in order to react with MMH for the monopropellant MMH system.Cygnus chose the industry standard S-405 as the catalyst, which must be heatedin order to be efficient. A pressurant tank containing Helium will also be used toregulate the propellant subsystem.

A regulated hybrid propulsion system is ideal for the proposed satellite becauseit provides the necessary thrust, while using less Mp than a bipropellant system. Afull bipropellant system would utilize oxidizer for both the apogee kick motor andthe RCS thrusters. The potential benefits of reducing Mp as much as possible is whythe regulated hybrid system was chosen.

6.5.2 Sizing

In order to size the propellant tanks, it is was first necessary to determine the ∆V nec-essary for Geosynchronous Transfer Orbit (GTO) to Geosynchronous Orbit (GEO)transfer and then add it to the several ∆V ’s necessary for station-keeping and at-titude control. An oxidizer-fuel (OF) ratio of 1.64 for the NTO/MMH system wasused, which allowed manufacturing companies to size the oxidizer and fuel tanks tosimilar sizes.

Since fuel is needed for both the apogee kick motor and the RCS thruster, thefuel tank needed to be larger than the oxidizer tank. In order to size the oxidizer,eq. (7) was utilized to find the mass of the propellant, where 1.2 represents a 20%safety margin, and a ∆V = 1.8144km/s is required for a transfer between GTOand GEO. Next, the mass of the oxidizer and fuel were calculated using eq. (8) andeq. (9), respectively. Equation (10) was then used to calculate oxidizer (NTO) tankvolume where the ρp of NTO = 1.38 g

cm3 at 323 K. Sizing the fuel tank required a∆V = 0.704534km

sand ρp of MMH = 0.847 g

cm3 at 323 K and was substituted intoeq. (7). The resulting mass represented the fuel necessary for the RCS thrusters overa 15 year span and must be added to the calculated mass of the fuel necessary for theapogee kick motor to determine the total fuel tank volume. Equation (11) was used tofind the volume of the pressurant tank Vpres by and dividing mass of helium (Mgas)by the density of helium (ρ). Figure 40 illustrates the data utilized to constructa preliminary design of propellant tank assembly. Tank mass was dependent onmaterial used and thickness. Tanks will be constructed with 316 annealed stainlesssteel because it is less expensive than titanium and easier to wield.

Ue = Isp ∗ go (5)

Mp = 1.2 ∗mo ∗ (1− e−∆VUe ) (6)

Mox =(OF ∗Mp)

(1 +OF )(7)

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Mfu =Mp

(1 +OF )(8)

Vp =Mp

ρp=Mox

ρox=Mfu

ρfu(9)

Mgas =(P ∗ Vp)

(Rgas ∗ T − Pρp)

(10)

Vpres =Mgas

ρ(11)

(a) Propulsion System Data (b) Propellant Tank Design

Figure 40: Propulsion System

6.5.3 Apogee Kick Motor

For the transfer orbit between GTO and GEO, Cygnus Satellite LLC chose Aero-jet/Rocketdynes R-4D 490 N (110-lbf) Bipropellant Rocket Engine. This motor metall the requirements of the satellite seen in fig. 41.

(a) Propulsion System Data (b) AKM System Design

Figure 41: Apogee Kick Motor (AKM)

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Figure 42: Apogee Kick Motor Specs[2]

6.5.4 Reaction Control System (RCS) Thrusters

In all satellites, some form of a reaction control system is necessary for attitudecontrol and station-keeping. In the case of the Cygnus satellite, this role will befulfilled by a system of RCS thrusters and reaction wheels. Figure 43 shows therequirements Cygnus Satellite considered before choosing its RCS thrusters. CygnusSatellite LLC chose the MR-111C 4N (1.0-lbf) Rocket Engine Assembly which isa monopropellant system that utilizes hydrazine, however, since MMH is a morevolatile type of hydrazine, it was decided to use MMH instead. This will reduce thenumber of propellant tanks, which greatly reduced the mass of the satellite.

(a) RCS Info (b) RCS Thruster Assembly

Figure 43: RCS Thrusters

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Figure 44: RCS Thruster Specs[3]

6.5.5 Propellant Manifold

The last part of the propulsion subsystem was the propellant manifold, which encom-passed all the hardware required to regulate propellant flow between the propellanttanks and the thrusters. The propellant manifold consisted of thruster valves, linesand fittings, isolation valves, pyro valves, filters, fill and drain valves, pressure trans-ducers, and flow control orifices.

The materials most commonly used to construct the lines and fittings in satellitesare titanium and stainless steel. Titanium is lighter and more compatible withoxidizers, however stainless steel is less expensive and easier to handle [23]. It wasdecided to utilize titanium for its lines and fittings in an effort to keep the satelliteas efficient as possible.

The two essential valve types utilized in the propellant manifold included isolationvalves and pyro valves. Isolation valves have the capability to permanently openor close without a continuous power supply. This property allows isolation valvesto serve multiple functions, one of which includes isolating a group of thrusters inthe event of system failure. Isolation valves may also control spacecraft mass bycontaining a specific tank in multi-tank system. On the other hand, pyro valves areone-time use valves, which means that these valves are either normally opened orclosed. Pyro valves may serve the same function as isolation valves, however theycan only be operated once. The advantage of utilizing pyro valves include lower leakrates, decreased pressure drops, and smaller mass. Pyro valves may be used to isolatecomponents in order to satisfy safety and reliability issues, and isolate componentsafter use. A system of both types of valves must be utilized in order to create themost efficient propellant manifold.

52

It is standard practice to install filters downstream of tanks and fill/drain valves,because this is where the most particulates can be captured. The size of the fil-ter depends on the amount propellant required to pass through the filter, size ofparticulate filtration, and allowable steady state pressure drop [23].

Fill and drain valves were next installed on a manifold, and had to remain ac-cessible at all times to allow for emergency offloading. The addition of pressuretransducers allowed for the pressure monitoring required for propellant loading andpressurization. Pressures transducers were also utilized to evaluate the performanceof the system. Finally, flow control orifices were installed to equalize pressure dropsbetween the feed lines between the oxidizer and fuel lines, which ensured a stableoxidizer to fuel ratio. Flow control orifices were also used to minimize transient flow,which can cause dangerous pressure spikes. Figure 45 illustrates the basic block di-agram of the 4.5 kg propellant manifold that Cygnus Satellite LLC plans to installin its new satellite.

Figure 45: Propellant Manifold Block Diagram

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The below fig. 46 shows the weight and power the individual components of thepropulsion subsystem. There will be 12 RCS thrusters for this satellite.

Figure 46: Propulsion Subsystem Weight and Power

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6.6 Power System

Figure 47: Sum of all Subsystem Power Requirements

Once all of a spacecrafts power requirements are accounted for, its power sourcemust be designed to meet those demands, which in this case will be a solar ar-ray. While satellites that provide internet and telephone services have a fluctuatingdemand for communication system power, this apply to direct broadcast televisionsatellites. While the content of the television programming is subject to change basedon the time of day, the power requirement of the Cygnus satellites communicationwill remain constant 24 hours a day until the day it is de-orbited. Therefore, thesolar arrays must be sized to provide enough power in the daylight hours to bothsupply the normal daylight power requirements, as well as charge the batteries whichprovide the spacecraft with power during eclipses. The necessary power generatedby the solar array during daylight is given by eq. (12)

Psa =PeTeXe

+ PdTdXd

Td(12)

Pd = Pe +Cbat[W − hr]

(Td[hr])(13)

In this case, the power requirement during eclipse is simply the sum of all subsys-tem power requirements (minus battery charging). The daylight power also includesthese standard operating values, as well as the power required to charge the batteryarray (eq. (13)). Based on the wobble of Earth’s polar axis with respect to the planeof the Ecliptic, the amount of time spent by Cygnus eclipsed in Earth’s shadow eachrevolution changes throughout the year. It is at the height of these eclipse seasonsthat the solar array needed to be sized, in order to ensure adequate power production

55

at this point of highest energy storage demand. Thus, an eclipse period of 72 minuteswas used.

In order to determine which type of solar cell to use, a trade study between theproperties of various materials was analyzed (fig. 48) As can be seen from fig. 48,

Figure 48: Candidate Solar Cell Properties[23]

there are a number of major differences between the cell types. Cost is always adriving factor, though not of higher priority than product quality in the case of theCygnus satellite. This company policy meshes well with the recent advances in solarcell manufacturing which have made Triple Junction GaAs cells an affordable, high-quality option. Their high efficiency and impressive EOL properties also make thema desirable candidate for the solar arrays on our satellite. Once the solar cell typewas chosen, its efficiency was used to calculate the maximum power output per areawith the Sun normal to the array (eq. (15)). This value was then multiplied by anominal inherent degradation value of 0.72, and the normal component of sunlight atthe worst-case Sun incidence angle of 23.7◦, present during the summer and wintersolstice to find the Beginning Of Life (BOL) power output per area of the solar array(eq. (15)).

Po = 0.30 ∗ 1369W

m2(14)

PBOL = PoId cos(θ) (15)

Next, the GaAs Triple Junction cell degradation rate of 0.5%/yr was used alongwith a mission duration of 15 years in order to find lifetime cell degradation (eq. (16)).End of life power per area was then calculated (eq. (17)), and used to calculate thenecessary solar array area to provide the Cygnus satellite with adequate power forthe entirety of its missions duration (eq. (18)).

Ld = (1−D)L (16)

PBOL = PBOLLd (17)

Asa =PsaPBOL

(18)

Using a preliminary total power value of 4.846 kW, a necessary solar array areaof 31.03m2 was calculated. A Triple Junction GaAs cell manufactured and marketedby Azur Space was used as a model for this project (fig. 49). The surface featuresvisible are integrated power leads and bypass diodes, necessary to connect and elec-trically isolate the cell in case of damage or structural shadowing. Bypass diodesare necessary due to an increased resistance solar cells inherit when partially or fullyshadowed. Construction of the solar array began with an aluminum honeycomb sup-port structure. Next, a layer of carbon fiber cloth was added to provide thermal andimpact insulation to the solar cells. A mounting structure was attached to the top of

56

this layer, in which all cells were installed. The anti-sun facing side was then coveredin a layer of MLI in order to help regulated thermal dumping. All layers are fixedtogether via an aerospace adhesive. The exploded view of this design is visible infig. 50. The mass of a given solar array segment was tracked by adding the mass ofall of its components (fig. 51).

Figure 49: Triple Junction GaAs Cell [26]

Figure 50: Solar Array Exploded View

Figure 51: Calculation of Solar Array

57

In order to maximize power generation, a 1 Degree of Freedom (DoF) configu-ration was chosen. In a 1-DOF system, the sun incidence angle decreases by 23.5◦

at the Winter and Summer solstices, an acceptable loss in avoiding the added struc-tural complexity associated with a 2-DOF system. A low-power, high-torque gimbalmotor manufactured and distributed by MOOG [27] was therefore chosen to providesolar array rotation (Figure 52). During the launch phase of the satellites life, thesolar array will be folded up and secured against the side of the satellite by explosivebolts. Once the apogee burn is complete, the explosive bolts will detonate, and thetorsional springs fixed to the inter-segment hinges will provide torque which extendsthe solar array fully. The next step in the power system sizing was the determinationof the Bus Voltage, an important step in the selection of a Power Conditioning andDistribution Unit (PCDU). Power losses increase as resistance and current increases,and as result of Ohms Law, power losses increases proportional to the square ofcurrent (eq. (19)).

P = I2R (19)

Figure 52: Single DOF Solar Gimbal Motor [27]

Therefore, minimizing power loss is a matter of increasing operating voltage (De-sign of Geosynchronous Satellites). Power distributors which operate at higher volt-ages are ideal, thus the Thales Group Power Conditioning and Distribution UnitMedium Power Unit [28] was selected (Figure 53). This unit has the capability tooperate in both unregulated and regulated voltage modes. In an unregulated system,individual loaded components require their own voltage regulation circuitry. Reg-ulated systems eliminate this need for redundant circuitry at the price of slightlyreduced power efficiency (Agrawal). In order to decrease overall system complex-ity, a regulated bus voltage of 50 Volts was decided upon despite the nominal dropin efficiency. The Thales Group PCDU is ideal due to its high bus voltage, whichcorresponds well with the maximum voltage produced by solar panels at the pointwhich the satellite reemerges from the eclipse. In order for these values to match, theindividual Gallium Arsenide half-cells (Figure 48) which have an EOL open-circuitvoltage 2.522 Volts each will be wired in 19-cell series to produce a voltage of 47.918Volts. A total of 537 series will then be wired in parallel and connected to the

58

Figure 53: PCDU [28]

PCDU in order to encompass all cells. In order to ensure that power losses stay ata minimum, components further away from the conditioning and distributing unitwill have power transferred to them via smaller gauge wires in order to ensure thatlocalized bus voltage never drops below too far below 50 V. Next came the designof the battery array, which provides power to the spacecraft during eclipses. Thisprocess began by finding the necessary battery capacity, which was a function ofworst case eclipse energy required, and EOL battery Depth of Discharge (eq. (20)).

CBat =CbatTe

DOD ∗ ηb= 2854.4 kW − hr = 571.49 A− hr (20)

A batterys Depth of Discharge (DOD) is defined as the total battery capacity avail-able for discharge, and varies from battery to battery based on chemistry (Figure 54)After multiple charge/discharge cycles, a batterys DOD drops, decreasing the effec-tive energy available to the subsystems.

Other pros and cons associated with the various battery chemistries also hadto be considered, important factors included energy density, energy efficiency andtemperature range, a general trade study of which was analyzed (Figure 55). Sincerequired energy capacity is independent of battery chemistry, energy density of agiven battery will affect both the mass and proportions of the power subsystem. En-ergy efficiency will affect the available battery capacity required, also affecting massand size of the battery array. Temperature range will affect the complexity of ther-mal regulation systems, a smaller range corresponding to a more precise regulationof battery temperature.

For a 15 year mission, the battery system will undergo 1350 charge/discharge cy-cles due to the two annual 45 day eclipse seasons. Nickel-Cadmium batteries undergoa significant degradation in depth of discharge over this many cycles, disqualifyingthem from consideration in this mission. At the other end of the spectrum, NickelHydrogen batteries undergo no DOD losses in a mission of this duration, makingthem ideal. Lithium-ion batteries do suffer from DOD loss, but only nominally.

In considering all these factors, the decision was made to utilize a Lithium-ionbattery, with the only drawbacks being a slight DOD loss at EOL, and an increase incomplexity in the battery thermal regulation system. Searching for a viable candi-date product yielded few results however, as it is apparently not common practice foraerospace battery companies to publish their product specifications online. Eventu-ally, a viable Lithium Cobalt Oxide battery was found, designed and manufactured

59

by the American company Eagle-Picher (Figure 56) [18]. In order to determinethe number of necessary batteries, the actual battery capacity was divided by thepublished available battery capacity per unit (200 A-hr), and rounded up (eq. (21)).

Nbat =

CBat[Whr])Vbus

Cactual= 3 (21)

Figure 54: Depth of Discharge vs. Cycle [23]

Figure 55: Trade Study of Battery Characteristics

Figure 56: Lithium Cobalt Oxide Battery Unit [18]

Final results yielded an array consisting of 3 Lithium Cobalt Oxide batteries, eachwith an available capacity of 200 Amp hours and a mass of 63.5 Kilograms. Dueto the singular nature of the product selection, the total available battery capacityended up being 15.48% beyond requirements. This generous excess in energy willprovide a comfortable margin of safety in the event that one or two cells fail duringthe duration of the mission.

In order to estimate the cost of the power system, the solar cells were first exam-ined. An approximate value of various solar cell technologies provided was examined

60

(fig. 57) [23]. The Gallium Arsenide Multijunction price per watt of $617Watt

was multi-plied by required total EOL normal power of 5.89 KW. The resulting price was $3.63million.

Figure 57: Cost of Solar Cell Technologies [23]

Next, the aluminum used in the honeycomb structure was analyzed. SolidWorksmass properties was used to determine the mass of an individual honeycomb struc-ture. This was then multiplied by the number of structures, and the most recenthigh-estimate cost-per-kilogram of Aluminum 6061 [25]. Based on the small mass ofaluminum subsisting the honeycomb structures combined with the relatively cheapprice of wholesale Aluminum 6061, its price was deemed negligible.

In order to calculate cost of the carbon fiber cloth, its total square footage wasmultiplied by number of segments and the current cost per unit area of carbon fibercost fig. 58 [15].

Figure 58: Cost Calculation of Carbon Fiber Cloth [15]

In order to calculate battery cost, cost per Kilowatt-hour for a Lithium-ion bat-tery (fig. 59) was multiplied by battery capacity (fig. 60).

Figure 59: Cost of Various Types of Batteries

Figure 60: Battery Capacity

61

7 Risk and Cost Analysis

7.1 Risk and Reliability Analysis

The risk analysis was performed using a qualitative/quantitative fever chart assess-ment scheme. First, a set of probabilities were defined for each bin (1 through 5);these cut-offs were assigned with guidance from the exhaustive study of 1584 Earth-orbiting satellites:

Figure 61: Failure Graphs [8]

Over a 15-year mission duration, the study found that the contributions of eachsubsystem to total satellite failure did not exceed 25% (with a small exception toBOL TT&C systems). Therefore, the upper bound of 25% was set for the highestProbability bin. The other subsets are described as follows:

Figure 62: Probability

The ”Impact” metrics were more difficult to set. A complete failure of the speci-fied subsystem is highly unlikely, as redundancy is built into every subsystem on theCygnus spacecraft. For the sake of analysis, however, this situation was evaluated.The quantitative mission impact metrics are described in the figure below, with thehighest bin of 5 being a total mission failure.

62

Figure 63: Impact

With the ”Probability” and Impact bins defined, bin values can be assignedto each subsystem, and the results can be displayed on the fever chart. The binassignments are displayed in the following figure:

Figure 64: Impact

Figure 65: Impact

63

It is apparent that the TT&C subsystem poses the largest risk to the mission.Aside from the ADCS system, the TT&C subsystem is the largest contributor fortotal satellite failure; therefore efforts must be made to ensure that a high-qualityand robust TT&C subsystem is acquired, and that redundancy is maximized. Forthis reason, the Cygnus TT&C subsystem applies two layers of redundancy; twoseparate transponders ensure that, even if one were to fail or perform poorly, the othercould function properly. Multiple antennas also mitigate the effects of pointing errordue to ADCS system malfunction, and the system includes an emergency/back-uphemispherical aft antenna that could function during an uncontrolled spin or reversein direction. The details of the TT&C subsystem are outlined in section 6.4.

7.2 Cost

Cygnus Satellite LLC. will utilize the Unmanned Space Vehicle Cost Model, version8, (USCM8) in order to approximate the cost to build a state-of-the-art satellite.USCM8 estimates cost of satellite based on non-recurring cost and recurring cost.Non-recurring Cost Estimate Relationship (CER) predicts the cost of design anddevelopment, manufacturing, testing, and support equipment. Recurring CER pre-dicts the cost of fabrication, manufacturing, integration, assembly, and test of spacevehicle flight hardware.

The equations used to calculate the non-recurring and recurring CER are observedin fig. 66 and fig. 67 , respectively.

Figure 66: USCMB Non-Recurring

64

Figure 67: USCMB Recurring

Figure 68 Figure 68 shows the application of both USCM8 methods. The totalestimated or approximated cost of Cygnus-1 is $798,970,688.42.

Figure 68: Application of USCMB

65

8 Gallery

Figure 69: Side Exploded View

Figure 70: Isometric Packed Payload

66

Figure 71: Side Packed Payload

Figure 72: Space Side Assembly

67

Figure 73: Communications Compartment

Figure 74: Power Compartment

68

Figure 75: Packed Upper Compartment

Figure 76: Emergency and Launch Antenna

69

Figure 77: Battery Assembly

Figure 78: Antenna Assembly

70

Figure 79: Computer

Figure 80: Propulsion Compartment

71

Figure 81: Transmitter & Receiver Antennae

Figure 82: Thrust and Interlock ESPA Manifold

72

Figure 83: Gyroscope

Figure 84: Input/Output Multiplexer

Figure 85: Power Condition and Distribution Unit

73

Figure 86: RCS Thruster

Figure 87: Reaction Wheels

74

Figure 88: Star Tracker

Figure 89: TT&C Antenna

75

Figure 90: Emergency Antenna

Figure 91: Battery Array

76

Figure 92: Apogee Kick Motor

Figure 93: Solar Array Gimbal

77

Figure 94: Transponder

Figure 95: Travelling Wave Tube Amplifier Array

78

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