Space Based Solar Power Mission

7/30/2019 Space Based Solar Power Mission

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The Pennsylvania State University

Department of Aerospace Engineering

HELIOS

Space Based Solar Power Mission Design and Architecture

Final Report

AERSP 401A – Preliminary Spacecraft Design

7 December 2012

Team Members

Andre ColemanChen ZhangEmily Wolf

Joseph WieserJohn Baum

Raymond Rolston



Abstract

Mission Helios will design and build a space-based solar power generation system that

will transmit the collected solar energy back to Earth to provide sustainable power. The aim is to

provide the most electrical power possible given the mission budget, but at the lowest possiblecost to the consumer. The system will be made up of a large modular satellite in a modified

Molniya orbit with an orbital period of 7.063 hours, an apogee distance of 30,000km (less than

that of GEO), and an inclination of 63.4°. This will allow for maximum solar power collection,

as the satellite will be exposed to the sun constantly, while minimizing radial dispersion energy

loss during microwave transmission over long distances. The phases of the mission include

research, development, design, launch, mission life, and maintenance. Initial plans aspire to

launch critical satellite operational subsystems and the first solar power generation module by

2040, and immediately begin transmitting power to the ground station after power system

deployment. Additional power modules, which will be Brayton cycle solar dynamic turbo

alternators, are to be launched and autonomously added to the satellite as launch and budget

capabilities allow. This clean source of energy would be under privatized control and could be

distributed to anywhere in the United States by the existing ground-based electrical grid. Once

established as a viable form of sustainable energy production, the potential for additional

collection satellites could be explored.

The system is comprised of three major components: the solar collection power

generation module, the power conversion and microwave transmission module, and a ground-

based receiving station. These components break down to nine subsystems to perform the

various tasks essential for mission success.

The main structural component of the satellite will be a modular open aluminum truss

system, that will support all critical satellite equipment as well as allow for the overall power

generation capacity to be easily expanded. The truss will be a FAST system (fully articulated

square truss) that will pack efficiently in the launch vehicle and then rigidly expand in orbit to

allow the attachment of the solar dynamic power generation modules (SDMs) across a long

distance. The Falcon Heavy, the NASA SLS, or a combination of both launch vehicles will be

used for this mission due to their low cost per launch, the payload mass capability, and the

payload fairing dimensions. Key requirements of the propulsion subsystem are to insert the



satellite into the Molniya orbit, keep the transmission antenna constantly pointed at the ground

receiving station, and counter act orbit decay. The satellite will utilize the Variable Specific

Impulse Magnetoplasma Rocket (VASIMR) ion thrusters with Krypton gas fuel for inclination

change, attitude adjustment, and orbit modifications. Helios will utilize Johnson Space Center in

Houston, Texas to serve as the Mission Control Center and will build a facility just south of

Fairbanks, Alaska to serve as energy receiving facility, the Payload Operations Control Center,

and the Spacecraft Operations Control Center. Helios will maintain constant communication with

the ground station during power transmission. The ground station will use a large parabolic high

gain receiver antenna to maintain strong signal strength during downlink. It will also utilize a

single axis rotor for azimuth correction. The satellite will use a side-fed Cassegrain antenna for

its compactness and high gain value. Tracking and Data Relay Satellite System (TDRSS) will be

used to relay uplink commands to Helios during its orbit around perigee. Helios will decidedlybe on a modified Molniya orbit path with a reduced radius of apogee and period. This was done

to reduce the amount of power loss due to transmission distance. Helios will use a Molniya orbit

path, VASMIR ion-thrusters for reaction and attitude control, and the Micorcosm Autonomous

Navigation System (MANS) to gather information on Helios’s orbital parameters. Helios will be

running this software on a Honeywell Generic Very High Speed Integrated Circuit Computers

(GVSC). Overall power generation will come from multiple Brayton cycle solar dynamic turbo

alternators. The turbo alternators will collect the sun’s energy in the form of heat, and will

expand a heated working fluid through a turbine to generate more efficient electrical power over

solar photovoltaic cells. The thermal control system will monitor all the satellite’s critical

components and use passive and active thermal control to keep components within their

operational and survival temperature ranges. The passive control system is comprised of heat

pipes wrapped in MLI blankets transferring heat to radiators, while the active system substitutes

heat pipes for pumped fluid lines to remove heat. The thermal control subsystem is also

responsible for thermal analysis and testing of the payload. These things will be further

developed as the structure is defined. The mission payload includes the microwave transmission

antenna, which will relay electrical power as microwave radiation at a frequency of 2.45 GHz via

a slotted waveguide antenna to the ground station rectenna.

During the time that remains to complete this project, Team Helios plans to make several

specific decisions and further the research and design of the mission. With the launch site



established, a decision will be made as to using the Falcon Heavy, the NASA SLS, or a

combination of both launch vehicles. Included in this decision will be a specific launch time

table. A structural model of the satellite will be created and the truss backbone will be designed

and sized based on the stresses expected to be seen during launch and orbital maneuvers. The

collapsed expandable truss structure will also be designed so that it will fit within selected launch

vehicle payload fairing size and weight constraints. The placement of the VASIMR thrusters on

the satellite and precise amount of fuel stored for use constitutes future work to be done

regarding the propulsion systems. From there, the team will calculate the cost of the fuel and

thrusters based on the total mass of the satellite and the ∆ v required. The new ground station and

microwave receiver will be located just south of Fairbanks, Alaska. Knowing this, the cost of

building and maintaining this facility can be determined. A link budget analysis and bit error rate

(BER) calculation needs to be performed. Interference and noise including polarization, pathloss, free space loss, atmospheric and rain absorption need to be modeled to get a precise gain

value. The mass and size of the communication antennas will be determined after calculating the

gain value. The actual number of 10MW solar dynamic power generation modules and a

preliminary estimate of total power transmission possible within the 21 billion dollar budget will

also be determined. The specific design placement of the VASMIR thrusters must be determined

as well as the power needed on the satellite. The sensor placement to support the MANS must

also be implemented into the structural design. The specific technical data and pricing of the

Honeywell computer must also be obtained and investigated. Additionally, the energy required

for primary battery backup to power the Helios satellite in case of malfunction will be calculated

from essential equipment operating wattages. Selecting the specific components comprising the

passive and active thermal system for thermal control is also part of the team’s future work.

Selecting the precise design for the slotted waveguide array will be accomplished after further

research into antenna design in order to optimize the transmitter’s performance. Other particulars

to be determined include the antenna gain and theoretical values such as beam directivity, beam

power density, and beam radius at the ground station. The final goal of the semester is, with all

of the above decisions made, to create a possible budget estimate and the price per energy that

the satellite system can generate.



Table of Contents

1.0 Introduction ............................................................................................................................... 1

2.0 Mission Architecture ................................................................................................................. 2

3.0 Subsystems ................................................................................................................................ 8

3.1 Structure ................................................................................................................................ 8

3.2 Launch Vehicle ................................................................................................................... 15

3.3 Propulsion............................................................................................................................ 19

3.4 Ground Control ................................................................................................................... 21

3.5 Communications.................................................................................................................. 25

3.6 Guidance, Navigation, and Control ..................................................................................... 28

3.7 Command and Data Handling ............................................................................................. 34

3.8 Power ................................................................................................................................... 35

3.9 Thermal Control System ..................................................................................................... 41

3.10 Payload .............................................................................................................................. 54

4.0 Future Work ............................................................................................................................ 58

5.0 Conclusion .............................................................................................................................. 59

6.0 References ............................................................................................................................... 61

Appendix A ................................................................................................................................... 68

Appendix B ................................................................................................................................... 69

Appendix C ................................................................................................................................... 70

Appendix D ................................................................................................................................... 71



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1.0 Introduction

Solar power is available in limitless quantities, but humans to date have been confined to

collecting what energy reaches the surface of the planet. This energy is far smaller in magnitude

than the solar energy that has not been dissipated by, or reflected from Earth’s atmosphere;therefore, it is desirable to collect this energy in space and transmit it back to Earth in order to

avoid such losses. An endeavor similar to this has never been attempted due to the implications

of wireless power transmission over such great distances. The Helios mission will design, build,

and launch a solar power collection satellite into orbit that will collect solar energy, convert it to

microwave power, and transmit it back to the ground station on the Earth. This energy can then

be converted for use in standard terrestrial power grid with the ultimate goal of establishing a

permanent sustainable energy source.

The overall mission objectives are:

• Minimizing the total mass of the payload and satellite structure

• Minimizing the power consumption of the satellite while maximizing the power

transmitted to earth

• The space vehicle dimensions and mass must be compatible with the selected launch

vehicle

• The cost of the mission shall not exceed 21 billion dollars, including launch costs• Safety must be properly addressed taking into consideration the dangers of beaming

large amount of power to Earth

Based on these criteria, a mission architecture and a conceptual system design for a

space-based solar power system was developed, and a detailed design is currently being

produced. The power system will include one large modular solar power harvesting satellite with

a microwave power transmission payload, and one power receiving ground station. Overall

mission planning is still being finalized, but most critical subsystem operational decisions havebeen completed. Detailed design specifications of all subsystems, including specific components,

mass estimates, dimensional estimates, power estimates, and cost estimates, will be made in the

upcoming semester.



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2.0 Mission Architecture

The Conceptual Design and Architecture for Space-Based Solar Power System (SBSPS)

Mission is based off boosting the amount of available clean energy in the United States. The

main goal of this mission is to create the architecture for a satellite or satellite constellation thatharvests solar power from space and transmits that power to Earth. In addition to the design of

the satellite(s) and the power transmission instrument, the ground station and capture design

must also be included in the ultimate deliverable. The satellite or satellite constellations may be

in any type of Earth orbit and the harvesting satellite(s) may transmit power directly to a ground

station on Earth or can relay energy through a network of satellites on its way to Earth. In

addition to Earth orbits the moon is also an alternative for harvesting solar power.

The Helios satellite will maximize its solar power generation in space and the electricalpower received on the ground through the limiting factor of the 21 billion dollar budget.

Therefore, it can be assumed that the cost estimate for the Helios satellite will remained fixed at

no more and no less than 21 billion dollars. Future work regarding cost estimates will not be

focused on the overall amount (which will be fixed at 21 billion dollars), but in generating a

complete budget outline that shows exactly how the funds will be distributed throughout the

satellite subsystems.

Mission Timeline

The timeline for this mission requires about 25 years of research and development in

several areas that will help to increase the ability to make the satellite even more efficient. This

time period of 25 years includes development of solar power generation efficiency, microwave

power transmission research, and research in structural and thermal space materials. Starting in

2035, manufacturing of the systems must begin and will incorporate what had been developed

over the previous 25 years. From 2035 to 2040, building of the satellite payload, supportingsystems, and power generation modules will take place. These can be built separately as they

will autonomously assemble in space as each piece of the satellite is launched. In this time,

manufacturing of the launch vehicles required is also critical as well as the construction of the

ground station. The ground station will be built to include the Space Operations Control Center

and the Payload Operations Control Center, as well as the rectenna receiver. Starting in 2040 and



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continuing over the next five years, several launches will take up pieces of the satellite starting

with the payload and supporting structure and followed by the power generation modules. These

modules will autonomously assemble to the supporting structure once they have been launched

into space. The estimated life of the satellite is approximately 30 years indicating that end-of-life

for the mission will occur in approximately 2075. Figure 2.1 graphically displays the mission’s

timeframe.

Mission Architecture Drawing

The mission architecture can be seen in Figure 2.2 below. The overall spacecraft design is

large and launching the entire platform at once is not feasible. Helios has been designed to be a

modular system to be assembled over five years. The components will be launched one by oneand assembled in space. First, the main transmitter will be launched into orbit. The transmitter

will serve as the main body of the entire assembly module. In second launch phase, additional

components will be sent up. The entire power generation system will be autonomously

assembled piece by piece in space as all the necessary parts are gathered. Once all the mission

critical components are assembled and harvesting of solar energy can begin and power can be

transmitted back to Earth.

Vehicle Sketch

The current proposed concept space solar power system is made up of one large solar

harvesting satellite. This solar energy collecting, electric power generating, and power

transmission satellite will be comprised of three main components: the solar power generation

array, electric to microwave converter, and a long-distance microwave transmitter, as well as all

other required critical subsystem components. Due to its inherent large size (because of the sheer

amount of power that will be generated and current space power technology), the satellite will bebuilt, launched, and autonomously assembled in space through the use of individual structural

“modules.” The modules will be sized sufficiently to fit within the size and weight requirements

of a heavy lift launch vehicle, and upon entrance to orbit, will autonomously connect and “plug

in” to each other. These power generation modules will be solar dynamic turbo alternators,

which will use an inflatable rigidized structure (currently in technological development) to



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deploy a large solar heat concentrator. Due to the large expanses of space that the modules will

take up once deployed (inherent large size of satellite), the main supporting and connecting

structure of all the modules will be an open metallic truss setup, most likely to be fabricated out

of aluminum. This open truss “backbone” is the most efficient structural support in terms of

highest strength and stiffness at the lowest possible mass. See Figure 2.3 for a conceptual sketch

of the Helios vehicle.

Future Work

Future work relating to the mission architecture consists of determined exactly how and

when the modular design of the Helios satellite will be launched into orbit. The modular truss

structure of the Helios satellite will be designed and sized in order to efficiently fit and be packedwithin the launch vehicle. The solar dynamic modules are a fixed size (10 x 3 m) and a

packaging orientation that allows for the most efficient launch cost of these 14.1 ton modules

will also be determined. The timing between launches and the exact launch dates, including the

specific number of launches needed, will be determined based upon initial structural mass

estimates of the Helios truss backbone and how many solar dynamic modules are allowed based

on the mission budget.



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2012 – 2035Research and Development• Solar power generation

efficiency increase• Microwave power

transmission research• Structural and thermal

space materialsdevelopment

2035 – 2040Systems Manufacturing • Build satellite payload

and supporting systems• Start building power

generation modules• Manufacture launch

vehicles• Build ground station

2040 – 2045Launch Satellite • Begin launching

satellite components• Launch payload• Supporting structure• Power generation

modules• Autonomous space

assembly

2075Estimated End of Life• Power generation

system decay • Orbital decay

2012 2040 2045 2072035

Radiator nc entrator

BraytonConversion

Receiver

Ground Station Rectenna

Figure 2.1 – Helios Mission Timeline



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Figure 2.2 – Helios Mission Architecture



7

Figure 2.3 – Helios Vehicle Sketch



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3.0 Subsystems

3.1 Structure

The main requirement of the structures subsystem is to support, enclose, and protect all

other subsystems aboard the solar power spacecraft during mission operation as well as provide a

structural interface during assembly, transportation, and launch to orbit. The final structural

design must be able to accommodate all equipment from all other subsystems and protect it from

the space environment once in orbit, while first being strong enough to withstand the severe

dynamic, shock, and vibrational loads of the launch vehicle. The structural members of the

spacecraft must provide the mating and attachment points for all other subsystem components

such as thrusters, propellant tanks, plumbing, communication modules, data electronics modules,

power generation, wiring, and so on. In addition to surviving launch loads and protecting internalcomponents, the structure must also sustain the stresses and loads of perigee and apogee firings

to get into its final orbit, and the deployment of the large solar concentrators and radiators on

each Brayton cycle solar turbo alternator once in orbit. Finally, the main microwave power

transmitter will be deployed and the structure must be able to survive station keeping and attitude

adjustment firings in order to keep the transmitter aligned with the power receiving ground

station in Alaska.

The initial conceptual design of the Helios space solar power satellite structure wasconducted with the following defined requirements 51:

• Launch Loads and Space Environment (Design Strength & Stiffness Criteria)

• Mass Properties (Overall Weight, CG Location, Material Selection)

• Accommodations and Performance (Interfaces, Layout, Design Efficiency)

Loading and environmental properties are normally considered the main design driver for

a spacecraft structure. Specifically, ground and launch vehicle loads, shocks, and vibrations are

the highest that will be seen and protection from the space environment, including radiation anddebris/micrometeoroid shielding are important design factors. The launch vehicle loads are yet to

be determined, but they will be set as the limit load (or assumed maximum loading) for the

structure, and will be multiplied by a factor of safety of 1.5 in order to find the ultimate load that

the final structure will be designed to withstand without permanent deformation 5. In addition, the



9

(3.1)

fundamental frequency of the overall spacecraft and deployable power systems will be

determined using Equation 3.1 below, and compared to that of the launch vehicle to make sure

the resonant frequencies do not match and cause amplified loading.

= 12 3

The spacecraft will be modeled as a cantilever bean, constrained rigidly at the launch

vehicle interface, where f n is the fundamental frequency in Hertz, E is the modulus of elasticity

of the main structural material, I is the area moment of inertia, M is the mass of the craft, and L

is the length from the interface to the spacecraft CG. The frequency found above corresponds

directly to the stiffness of the craft and can largely determine the stresses experienced by the

spacecraft during launch. The structure will also be designed to protect sensitive components

from other subsystems from the space environment. Electronic components will be shielded from

radiation through an enclosed aluminum containment unit of sufficient thickness to block

radiation of to the individual component’s hardness. Additionally, all electronic equipment will

be protected from debris impact by the same aluminum enclosure as well as a Whipple shield of

Kevlar and Nextel wrap. 51 Gas propellant tanks for the station’s many VASMIR control and

boosting rockets will be fabricated from filament wound carbon-fiber composite, and any other

critical rocket components (pumps, fuel lines, valves) will be protected by exterior aluminum

cases and a Whipple shield to prevent space debris penetration as well.

There are two main types of overall satellite super-structures, body mounted and open

truss. 44 Due to the main mission requirement of large scale power generation, the overall

structure of Helios will be of a very large scale and will need to support a repeating configuration

of solar dynamic turbo alternator modules (SDMs) across a large expanse of space (163 linear

meters per 10MW solar turbo alternator). 44 Due to this, a large open truss was selected as the

main backbone of the satellite, with sensitive equipment being housed in enclosed solid

aluminum equipment boxes. A truss design offers the lightest weight and stiffest structure when

a large expense needs to be covered without the need for many subsystem features needing to be

attached. 51 This works perfectly for Helios as the large SDMs will be the only things attached

along the length of the truss backbone, while the microwave transmitter payload and other

subsystem components including communication and data handling will be enclosed in a body



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structure near one end of the satellite and mounted to the truss backbone using aluminum

honeycomb panels. The truss backbone can be of various cross sections (triangular, circular, or

rectangular), but a rectangular cross section was selected for Helios in order to give ample side

area for the SDM power generation modules to be attached, as well as propulsion thrusters for

control and attitude adjustment along the length of the satellite. In addition, a rectangular cross

section also allows for easier attachment of equipment boxes and will be able to accommodate

the microwave transmitter payload efficiently. The material selected for the equipment

enclosures and truss backbone must meet the design ultimate strength, specific stiffness, and

must also be easily available, lightweight, and easy to fabricate (machine, cast, and weld) into

large structures. Aluminum, Magnesium, Titanium, and Beryllium are the major lightweight

materials that make up most spacecraft structures. 44 Aluminum has been selected as the main

structural material for Helios due to its high strength-to-weight ratio, low cost, and ease of fabrication over the other materials. This will significantly reduce costs due to the large size of

the Helios satellite and the amount of fabrication that will be needed. Composite materials were

also considered, but due to the extreme cost of their fabrication, and the selection of an open

truss backbone, composite materials are not an efficient solution. In addition to just strength,

weight, and fabrication concerns, thought was also put into using the spacecraft structure for

other roles such as a heat sink and electrical ground. 44 This gives yet another reason why a

metallic structure was chosen over a composite structure. The power generation system is likely

to produce a large amount of thermal energy as well as electrical power, and the structure may be

used to sink some of that heat if needed. If more heat is required to be absorbed than Aluminum

can handle with its relatively low 400 degree Fahrenheit limit temperature, Titanium will be used

to replace the aluminum (limit temperature 1200 degrees Fahrenheit) at an increased cost and

fabrication difficulty. 44 Magnesium is much weaker than Aluminum; while Beryllium is strong

and very lightweight, but is toxic and extremely brittle, meaning both are not suitable for large

scale fabrication and a truss structure. 44

The overall conceptual layout of the Helios satellite was designed with efficiency and

mass balance in mind. The overall design is based upon a large truss backbone, from which all

the SDMs will be attached in a balanced configuration, alternating one on each side. At one end

of the truss backbone, will be a system of cross trusses (the main “body” of the spacecraft), on

which aluminum honeycomb panels will be mounted in between for the main equipment of all



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other subsystems to be attached within shielded equipment boxes. The main payload, a slotted

waveguide microwave power transmitter, will also be attached on this end, as will any and all

communication dishes, power converters, data handling electronics, and all other critical control

instruments. The large truss backbone that extends away from the collection of subsystem

equipment boxes on this main body will support only the SDMs, mounted on 1-axis rotary

gimbals in order to track the sun. Due to the size required of Helios in order to produce power on

the scale of megawatts, this large truss will be of a modular design, with each unit fabricated as

an expanding truss section, with a location at its center to which the SDMs will attach. The truss

will be an up scaled and increased strength version of the FAST Mast (fully articulated square

truss) currently in use upon the International Space Station. 58 The currently in use FAST truss

system on the ISS can be seen in Figure 3.1.

Figure 3.1 – Space Station FAST Expandable Truss Structure 58

The collapsible aluminum truss will be designed to the maximum strength and minimum

size possible in order to withstand both the stresses of launch, and also to support the stresses

involved with maneuvering the Helios satellite, which will grow to several kilometers in length

as SDMs are added in its linear configuration. A FAST truss system is stowed for launch in a

collapsed configuration, and deploys in orbit through the use of a motorized rotating nut. 58 As

each square truss batten (individual bay) is deployed, structural cross members that were buckledwhile stowed snap straight, and this stored energy stabilizes and rigidizes the entire structure by

tensioning the diagonal face cables in each bay. 58 A diagram of a FAST truss deploying can be

seen in Figure 3.2.



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Figure 3.2 – FAST Expandable Truss Structure Deployment 58

This system creates an extremely strong, yet extremely lightweight truss structure.

Additionally, the FAST truss design does NOT rotate as it deploys, meaning electrical wiring can

easily be run along its length and coiled within the collapsed configuration. This design

configuration also allows the structure to be sized and packaged according to the selected launch

vehicle. The truss square cross section will be sized so that it can be fit into the cross-sectional

area of the launch vehicle payload bay. The individual truss members themselves will be

designed with cross sectional areas proportional to the rigidity needed to support the Helios

satellite during launch and orbital maneuvers. The mass per length of the large FAST structureneeded for Helios is unknown at this time, but is estimated to be much smaller in magnitude than

the mass of the SDM power generation modules themselves. As many launches will be required

to get a solar power generation capacity of sufficient size into space, this modular truss design

allows the solar capacity to be increased incrementally along with the power produced, as truss

and power system modules would be launched as needed and then would autonomously attach

and “plug in” to the existing large truss end of the satellite until the mission budget is exhausted.

Please see the Mission Architecture and Overview sections of this report to view sketches of the

conceptual vehicle design in Figure 2.2 and Figure 2.3.

The Brayton cycle solar dynamic turbo alternator power generation modules will each

concentrate the suns heat energy using very large deployable thin film rigidized inflatable

Fresnel solar concentrators (163 meters in diameter) and then dissipate excess heat through very

large deployable rigidized inflatable radiators (1394 square meters of surface area). 44 The



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reasoning behind the selection of inflatable structures for the solar concentrators and radiators is

due to their extremely large size. The cost of launching such massive structures into space would

be immense, and then assembling them in space would be yet an additional large cost. Rigidized

inflatable space structures are low-cost, have a very high mechanical packing efficiency (very

low volume when packaged in launch vehicle), very low weight, high reliability, and are

currently able to be manufactured with a high surface precision. 28 NASA demonstrated the use of

an inflatable antenna structure during the 1996 Inflatable Antenna Experiment (IAE) on board

the STS-77 Space Shuttle mission. 28 The IAE proved that inflatable structures are an efficient

way to form large scale structures in space, and a picture of the IAE can be seen in Figure 3.3. It

can be seen that the 14 meter diameter inflatable antenna structure is comprised of two basic

elements, the inflatable thin film concentrator assembly and torus/strut inflatable supporting

structure surrounding the concentrator and connecting the entire structure to the spacecraft. 27

Figure 3.3 – NASA IAE Inflatable Structure 28

This setup is identical to how the Fresnel solar concentrators on each SDM will be

designed, but they will instead focus sunlight instead of radio waves onto a receiver. The

inflatable Fresnel concentrator on Helios will be made of aluminized mylar (just as the IAE) and



14

will be stressed to 1200 psi by a very low inflation pressure of 10 -6 psi in the surrounding torus

(which will be designed and sized accordingly to provide this stress to a 163 meter diameter

concentrator), giving an overall surface precision of less than 1 mm. 27, 28 The torus and support

struts will be made of Kevlar and Nextel (similar to a Whipple shield), in order to resist puncture

by space debris and retain their inflated dimensions. Additionally, near-term advances in space

inflatable technology will allow for further rigidization of the structure after deployment inflation

through a resin impregnated exterior fabric, greatly minimizing the effect of any pressure losses

due to leaks and puncture. 34 The liquid resin (similar to epoxy) maintains a high fabric flexibility

for a dense, low volume, packaging aboard the launch vehicle, but cures solid when deployed,

inflated, and exposed to ultraviolet light in space (even in the extremely low temperatures of

orbit). This solidifying effect greatly rigidizes the entire inflatable support structure, allowing an

increased lifespan and greater load capacity. 34 Both the solar concentrators and radiators will beinflated in orbit and pressure regulated using a solid subliming powder, inserted inside the

inflatable structure during manufacturing. The solid powder will sublime into a gas when

exposed to the space environment and is able to produce the required pressure of 10 -6 psi within

the structure. Subliming powders are non-toxic, low cost (no need for compressed gas tanks,

valves, and plumbing), non-corrosive, and are able to self-pressure regulate (excess powder

within the inflatable structure will only sublime into make-up gas as existing gas leaks out and

the pressure drops). 27

Future work for the structures subsystem includes finalizing overall design dimensions

and getting estimates of the mass and cost of the Helios satellite structure. This is one of the last

steps in the spacecraft design process, as sizes and mass requirements for all other subsystem

components will need to be determined before the overall size of the spacecraft structure can be

found as the structure will need to support and enclose this equipment. The loads and stresses

from the launch vehicle as well as in-orbit deployment will be finalized and the structure will be

designed to those loads plus a 1.5 factor of safety. In order to make sure strength and stiffness

design requirements are met, a 3D model of the overall vehicle will be developed in order to run

finite element analysis. This model will also be used to accurately determine mass estimates and

the center of gravity for use in launch vehicle analysis.



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3.2 Launch Vehicle

When considering a launch vehicle that would be suitable for the mission, the three most

important factors are cost of launch, getting the payload into the desired orbit, launch

mass/dimensions, and frequency of launch. With a mission budget of 21 billion dollars, the costof launch must be minimized. The payload reaching the desired orbit is crucial to the mission as

failure to do so would negate the mission’s initial purpose. Also, the desire to optimize each

launch to carry as much as possible, the launch mass and dimensions will be an important factor

in determining which launch vehicle works best for this mission. Finally, the selected launch

vehicle must be able to launch frequently enough to be able to transport all of the satellite

components within five years.

Other factors that will impact the final decision include launch vehicle reliability,

location of launch, and loads and vibrations placed on the payload at launch. Because the launch

vehicle is limited to be from the country within the satellite is launched, the choice of a launch

vehicle will narrow the launch location options. While loads and vibrations must be taken into

account, the launch vehicle must be able to handle the mass and dimension of the load first and

then structural requirements can be set. Finally, vehicle reliability is a factor in the launch

vehicle decision; however, it is not as crucial of a factor as the cost and the launch mass and

dimensions will be. A launch vehicle that has a reliability greater than 0.9 is the most desirable

(meaning 90% of the launches attempted were successful).

The desire to use a modified Molniya orbit would limit the options for launch locations.

Launching from Plesetsk, Russia would not require an inclination change. 40 There is also the

option to launch out of Vandenberg Air Force Base (VAFB) or Kennedy Space Flight Center

(KSCF) however, an inclination change would be necessary to reach the desired orbit. A

Molniya orbit has an inclination of 64.3°. KSFC can launch with an inclination between 39° and

57° depending on the launch azimuth while VAFB can launch between an inclination of 70° and

104°. 39 However, changing the inclination is not the only ∆v required to achieve the desired

orbit.

The ∆v required is the sum of three orbital transfers to achieve our desired orbit. The first

transfer is a Hohmann transfer from the manufacturer’s circular orbit in LEO to a circular orbit

with the same radius as our desired orbit’s radius at perigee, 7378.14 km. The second transfer is



16

(3.2)

(3.3)

(3.4)

(3.5)

a pure inclination change from the manufacturer’s launch inclination to our desired inclination of

63.4°. The third transfer is from the circular orbit to the Molniya orbit with an eccentricity of

0.605216. The total ∆v can be c alculated by using Equations 3.2, 3.3, 3.4, and 3.5 and summing

∆v1, ∆v 2, and ∆v 3. (Notes: Subscript 1 indicates the first circular orbit as specified by the

manufacturer. Subscript 2 indicates the second circular orbit with manufacturer’s inclination and

radius equal to 7378.14 km. Subscript 3 indicates final desired orbit)

=

∆1= 2+1−1 + 1−1 2+1

∆2= 2√ sin ∆2

∆3= � 2 + −� 2 + + � 2 + −� 2 +

The easiest way to reduce the ∆v required is to slightly reduce the payload mass to

increase the launch inclination. Seeing as the pure inclination change is the largest contributor to

the ∆v, reducing the change in angle would be the easiest solution in reducing the overall ∆v.

Of the Russian-made launch vehicles available, ones with a reliability factor greater than0.9, include the Proton, Rockot, Soyuz, and Zenit. The Proton has the lowest reliability factor

(0.922), followed by the Zenit (0.933), the Rockot (0.941), and the Soyuz (0.944) which has the

highest reliability factor as of December 31, 2010. The Soyuz and the Proton have had the most

launches of the four systems. 41 Of these four systems with a reliability factor about 0.9, only the

Soyuz and the Rockot are able to launch from the Plesetsk launch site effectively eliminating the

Proton and Zenit as viable vehicles.

Also under consideration are American launch vehicles including the Atlas V, the DeltaII and IV Heavy, the Falcon 9, the Falcon Heavy, the Minotaur I and IV, and the NASA SLS

Block II. Each of these launch vehicles, with the exception of the Delta II (0.988), the Falcon

Heavy (has not been launched), and the NASA SLS Block II (has not been launched), has a

reliability factor of a 1.000 meaning that all of their launches has been successful. 41



17

Following the requirements stated, a trade study was conducted between many of the

potential launch vehicles described above. Of those launch vehicles, the Rockot, Minotaur I, and

Minotaur IV were all eliminated as they could only carry payloads of 1,950 kg, 580 kg, and

1,735 kg respectively into Low Earth Orbit (LEO). 49, 50, 54 In Table 3.1 below, these requirements

are stated and directly compared. The “Cost per Launch” is given in millions of US Dollars and

it is noted what year the stated price was quoted for each launch system. The “Payload Capacity”

is given in kilograms and would be the capacity to each manufacturer’s specific orbit altitude and

inclination in LEO. Under “Payload Dimensions”, the listed height of the payload fairing for

each system is the overall length of the fairing and the diameter is the outer diameter of the

fairing with all measurements in meters. The “∆v Required” column lists a maximum and

minimum value – the maximum indicating the ∆v required to move from the manufacturer’s

achievable orbit per the payload mas s and the minimum indicating the ∆v required using alaunch inclination of 57° as is the closest possible to our desired inclination from the launch site,

KSFC. This ∆v will be reduced by slightly decreasing the payload mass in order to launch at an

inclination closer to that of our desired Molniya orbit. Finally, the “Reliability Factor” is listed

for each system to indicate the system’s historical success if applicable.





19

launch multiple times in one year could be desirable. This would mean either using only the

Falcon Heavy or both the Falcon Heavy and the SLS.

3.3 Propulsion

The primary concerns during the course of this mission will be ∆ v required for insertion

into the Molniya orbit, attitude control and orbital decay. The antenna must remain pointed at the

ground station, and the solar energy collection system must point at the sun in order to minimize

cosine losses and maximize energy collection. Based on the levels of radiation the system will

experience over the course of its lifetime, the satellite’s expected total lifetime is about 30 years

due to solar panel degradation and radiation exposure. The selected modified Molniya orbit

poses inherent difficulties related to orbit decay including atmospheric drag and J2 effects. Ionthrusters are very efficient and have a very long lifespan due to their having very few moving

parts. These propulsion systems will draw the necessary power from the solar energy collected

on board the satellite. Both of these systems eliminate the need to carry a large amount of

explosive gas (such as oxygen) on board. With the potential for micrometeorite damage,

carrying a large amount of pressurized fuel for attitude thrusters is undesirable due to the danger

of catastrophic failure. The high efficiency and long lifespan of ion and plasma thrusters is the

best option for the mission.

A diagram of a basic ion thruster is included in Figure 3.4. The choices of actual thruster

design was determined by trade studies conducted based on each thruster’s specifications

including type of fuel, service lifespan, required power, thrust, and specific impulse. Major

contenders included NASA’s NSTAR, VASMIR (Variable Specific Impulse Magnetoplasma

Rocket), and the experimental MPDT (Magnetoplasmadynamic Thruster).



20

Figure 3.4 – Basic Ion Thruster Design

Comparison was conducted between each choice across traits of specific impulse, power

requirement, lifespan, size, complexity, and total thrust. The wild card in the whole trade-study

was that the VASIMR system possesses the inherent 'variable thrust' capability65

, whereas noneof the others do. Given a relatively low-to-moderate rate of power consumption (200 kW total),

long lifespan, high durability, relatively moderate-to-high thrust capability, simplicity of design

compared to other ion thruster types (VASIMR has very few moving parts and no electrodes),

compact size, and high specific impulse (~5000 s) VASIMR was selected to be the most suitable

choice for our mission. 65 NASA has tested the system on the ISS, and intends to replace the

current method of chemical rocket boosting to counteract the effects of atmospheric drag. 65 This

is significant, as our satellite will experience very similar effects. The estimated savings amount

to $210 million per year vice chemical rockets. 65

Ad Astra Rocket Company’s Variable Specific Impulse Magnetoplasma Rocket

(VASIMR – shown in Figure 3.5) will be the specific thruster used in the mission. It has a

specific impulse of 5000 seconds, which provides a thrust of approximately 5 Newtons. 42 The

thrust can be varied by throttling the input power up or down; power will be drawn directly from



21

the solar turbo alternators. The large power requirements of this system will not be an issue, as

the turbo alternators will provide power in the megawatt range, while the thrusters require pulsed

power in the range of kilowatts. The thruster will utilize Krypton gas propellant, which has

yielded a total system efficiency of just over 60%. 23 The ∆ v required to enter into the Molniya

orbit at an inclination of 63.4 ° of from the initial inclination of 28.5 ° is 4.8906 km/s. This

includes the transition from circular to elliptical orbit and the inclination change. This system

will also be responsible for augmenting the radius from 200 km to 1000 km altitude. The fuel

will not require replenishing during the lifetime of the satellite, due to the high efficiency of the

VASIMR thruster.

Figure 3.5 – VASIMR Ion Thruster Design65

3.4 Ground Control

The ground control system will be a series of facilities that will control different aspects

of the mission. These facilities are known as the Spacecraft Operations Control Center (SOCC),



22

the Payload Operations Control Center (POCC), and the Mission Control Center (MCC). The

SOCC and the POCC will be co-located in Alaska and the MCC will be at Johnson Space Center

(JSC) in Houston, Texas. Each launch will take place out of Kennedy Space Center (KSC) in

Cape Canaveral, Florida. KSC will have control of the procedures preceding the launch. Once

the launch vehicle has cleared the tower, JSC will take over responsibilities for the mission. The

SOCC will be required to command the payload through the orbital transfers to achieve the

desired orbit. It will also monitor and send commands to the satellite that are mission specific

such as when to make orbital transfers, when to boost the satellite back into the desired orbit, and

when to autonomously connect another component of the satellite. The POCC will be

responsible for all that is related to the microwave transmitter; for example, the direction it is

pointing for transmission.

The specific location of the ground control station is an important issue to consider – it

must be within the northern hemisphere due to the selection of a Molniya orbit. Additionally,

there must be enough ground space to build an estimated 3- to 5-kilmoeter microwave receiver

dish as well as both the SOCC/POCC building and the power storage/distribution facility. The

location must be reasonable in cost to transport building materials and construct the facilities.

The ability of the location to transmit the power received from space to people effectively and

without further significant energy loss is also crucial. With these criteria in mind, the locations

being considered are northern Canada, Greenland, Alaska, and northern Russia.

If the ground control antenna is placed above 53 ° latitude, the satellite's elevation relative

to the ground will be between -80 ° and -70 ° for most of the orbit (see Appendix A). A satellite

with an altitude over 2000 km will typically require an elevation of at least 10 degrees of

freedom while in line of sight with the ground station 38; thus, placing a ground station

somewhere in northern Canada or Greenland would eliminate the need for elevation-tracking.

Each of these locations has satisfied the first requirement – they are all far north in the

northern hemisphere. This will allow the satellite the maximum amount of transmission time due

to the characteristics of the selected Molniya orbit. The second requirement of ground space will

be difficult in some areas due to protected land, however, each of these locations also satisfy the

need for a large amount of land. This land must also be flat enough to be able to build a large

receiver dish which, again, each of these locations satisfies as a requirement. The last



23

requirement being the cost of building the facilities is still being investigated. The movement of

materials, the building and maintenance of the facilities, and the safety and care of the personnel

involved are extremely important.

Finally, it is desired that the location be able to disseminate the power it receivedeffectively and with minimal further energy losses. This desire encourages the choice of Alaska,

Canada, or Russia. Further analysis of energy transmission loss over a distance remains as

research for the future but will also have an impact on the final decision of the location of the

ground control station.

With all of these factors in mind with a focus on consideration for the workforce the

ground facilities will require, the location of the POCC and SOCC ground stations as well as the

receiver dish will be located just southwest of Fairbanks, Alaska shown in Figures 3.6 and 3.7.

This is optimal for several reasons. This location provides the required terrain and land area for

building a large receiver dish as well as the buildings to support the mission. It is close to a major

city in Alaska that includes highways and an international airport, which will be useful in

reducing building costs associated with transporting materials. Also, this location specifically

grants almost 76% access time (measured over two weeks) to the satellite for transmission of

power. Shown below in Figure 3.8 is a sample set of access time to the selected location

highlighting approximately two days. Finally, while Fairbanks itself it not part of the United

States power grid, this would be an opportunity for them to provide continuous power to their

city as well as become connected to the power grid. 16 With this connection to the power grid, the

power can be transmitted, received, and distributed to the United States population with the

minimum amount of power loss.



24

Figures 3.6 and 3.7 – Images of Alaska and Area South of Fairbanks, Alaska

Figure 3.8 – Satellite Access Times in Fairbanks, Alaska

Future work requires analysis into the cost of the materials to build, the construction, the

maintenance, and the operation of the ground facility. With the location of the POCC/SOCC

facilities and the receiver dish established, the specific costs of this effort will be further



25

investigated. Other future work requires discussion about the staff of the ground station including

their pay, benefits, and hours. This also inherently incurs a cost that will be a part of the budget

through the duration of the mission.

3.5 Communications

Helios needs to be closely monitored during its operation cycles due to the large amount

of solar and electric energy being stored and transmitted. Therefore, it is important to maintain a

constant and stable communication with the ground station during periods of access. Access

times are shown in Figure 3.8 in the previous section. The ground station will have two antennas,

one for reception and one for transmission. The receiver antenna will retrieve telemetry, station

keeping and operation data from the satellite’s GNC, thermal and power subsystems duringaccess to Helios; while the transmitting antenna will handle sending commands directly to the

satellite.

The ground station will both receive and transmit data to Helios. The architecture is

based on a design by Gerald Maral in Satellite Communication Systems 43; it is shown in Figure

3.9. The data transfer will be performed by directly linking the ground station and the satellite

during line of sight access, which accounts for most of Helios’s orbit.

Figure 3.9 – Design architecture for ground segment of communication system



26

(3.6)

(3.7)

Antenna pointing is another important design factor for the ground segment. While most

elliptical orbits require a rotating azimuth-elevation tracking antenna in order to align with the

satellite properly. Helios’s modified Molniya orbits spend over 80% of its orbit over the northern

hemisphere in low elevation. And with the ground station placed in Alaska, the satellite's

elevation relative to the ground will be between 80 ° and 70 ° for most of the access time (see

Appendix A). This eliminates the need for elevation tracking when combined with a large sized

parabolic reflector antenna. 68

Helios’s orbit has a very large apogee radius of 30000 km (see Appendix B), the

transmission power is a proportional to antenna gain and inversely proportional to transmission

distance as seen in Equation 3.6. 26 So in order to maintain good signal strength, a high gain

antenna is necessary to offset the large transmission distance r . A rough estimation of antenna

gain can be calculated using Equation 3.7 (assuming 80% efficiency). With an antenna diameter

of 30m and frequency of 4000 to 6000 MHz, gain is estimated to be on the order of 50 to 100

dB.

= (λ , ) ( , )

(4πr)

= 10 10(4 πr) 2

λ 2

Helios will not be in line of sight of the ground control as it orbits around perigee. The

ground station will use the ephemeris data collected from GNC system to predict Helios’s orbit

around perigee. The United States Space Surveillance Network (SPACETRACK) will be used as

a backup tracking network in cases of equipment failure. During this operational period, Helios

will not be transmitting energy back to Earth, so there is minimal amount of downlink between

the onboard satellite systems and the ground station. Helios’ perigee attitude of 1000km also

means there is negligible amount of atmospheric drag; this eliminates the need for frequent

uplink commands for attitude correction around perigee. In cases of unexpected orbit

perturbations, the ground station will use existing geostationary communication satellites to relay

uplink commands to Helios.



27

(3.8)

Launch vehicle imposes a constraint on the size of antennas onboard the satellite. So

packaging is also an important factor in consideration. A side-fed Cassegrain antenna

configuration is chosen for its excellent signal strength and compactness inside the launch

vehicle. 11 Its aperture diagram is shown in Figure 3.10.

Figure 3.10 – Side-fed Cassegrain Antenna Aperture Configuration 11

Helios’s communication system will operate at C-band, which has downlink frequencyaround 4 GHz and uplink frequency around 6 GHz. 57 C-band was chosen because of its ability to

operate during adverse weather conditions. There is also a large pool of Russian heritage

missions that operate in the C-band to study from. 37

Helios will spend a significant portion of its orbit outside the Van Allen belt, so radiation

shielding for communication hardware onboard the satellite will be important. This could affect

the radiation efficiency of the antenna gain as well 8: since gain is the product of directivity and

radiation efficiency. Directivity is given by Equation 3.8. Lastly, radiation shielding forcommunication hardware will also add to the mass of the satellite itself.

( , ) = (4)⁄

A link budget analysis needs to be performed (possibly using STK). A simple link budget

can be estimated using Equation 3.9. Additional terms such as free space loss, path loss, and

polarization loss can also been added into Equation 3.9. A feedline system needs to be

determined for the ground and the satellite receiving antennas. 9 Velocity factor and antenna

impedance will need to be taken into account to optimize antenna gain. 8 Specifications for

transmitter and receiver antenna dishes need to be determined in conjunction with launch





29

Figure 3.11 – Power-Generation Satellite with Two Relay Satellites in Orbit

The two relay satellites are offset at 180 degrees in true anomaly and 50 degrees in

RAAN to ensure there is a near constant access to the power generation satellite. An access time

graph can be seen in Figure 3.12 below.

Figure 3.12 – Relay to Harvesting Satellite Access Times

The ground station will eventually move out of line of sight from the relay satellites due

to Earth’s spinning axis. This means that several ground stations needed to be constructed

maintain near constant transmission to the relay satellites. An access report to a ground station



30

placed in Northern United States over a period of one day can be seen in Appendix B. The report

shows that the ground station only has access for only 11460 seconds or about 3 hours per day. A

circular non-GEO orbit will always result in poor access time to a fixed ground station.

Using Molniya orbits presented itself as a possible solution. Molniya orbit is unique inthe sense that it will have high access time to any points in the northern hemisphere. Molniya

orbit also has an inclination of 63.4°, which guarantees near constant sunlight exposure. 37 Lastly,

Russians have been using Molniya orbits for communication satellites since the 1960s. There is a

large pool of Russian heritage launches for the team to study. 37 Those properties make Molniya

very appealing to the project.

Molniya orbits have good access times to anywhere in the Northern Hemisphere. This

orbit also eliminated the need for relay satellites altogether. The main concern for utilizing

Molniya orbit is the transmission loss due to its extreme distance at apogee. Power transmission

loss is affected by transmission distance, solar, and magnetic interference. The altitude at apogee

can be as high as 40,000 km, the transmission loss at that distance will be severe; the satellite

will also be outside the Van Allen belt at that distance, so radiation exposure to hardware and

magnetic interference is also a concern.

In order to address the transmission loss issue, reducing the Molniya orbit’s semi-major

axis was necessary (thus changing its eccentricity and period as a result). However, shortening its

semi-major axis also means a reduction in access time to the ground station because the satellite

will spent less time near the northern hemisphere. A tradeoff study was done to compare the

different orbits. A detailed comparison between different Molniya orbits is shown in Appendix C

and D. Eccentricity and orbit periods are both dependent values of the semimajor axis. Below is

Figure 3.13 of different access times as a function of orbit period.



31

Figure 3.13 – Access percent as a function of orbit period

As seen from Figure 3.13, access time decreases with a shorter orbit period. The original

Molniya orbit is designed with a period of half a sidereal day to avoid J2 effects. The modifiedorbits also bring the perigee closer to earth, so the satellite will experience additional

atmospheric drag. See Appendix B for the specific parameters of the selected orbit.

The primary concern during the course of this mission will be attitude control and orbital

decay. The antenna must remain pointed at the ground station, and the solar energy collection

system must point at the sun in order to minimize cosine losses and maximize energy collection.

Based on the levels of radiation the system will experience over the course of its lifetime, the

satellite’s expected total lifetime is about 30 years due to solar panel degradation and radiationexposure. The selected modified Molniya orbit poses inherent difficulties related to orbit decay

including atmospheric drag and J2 effects. Also, due to radiation and decay of solar panels, a

new satellite or a new set of panels will have to be sent up periodically. Using the VASMIR

propulsion system modifications to orbit path and attitude will be able to be made.



32

Power transmission is a critical part of the Helios mission. With current technologies

there is a lot of power lost in transmission, so ensuring that the transmitter is always aligned with

the ground station is imperative. By running a system that can quickly processes the Helios

satellite’s current and future attitude and direction, the on board computer should be able to

direct the angles of the transmitter should be pointing at all times to reach the ground station.

Additionally, due to imperfections in orbit determination and the real possibility of being

knocked out the desired orbit from space debris it is absolutely vital that this system be

constantly running throughout the duration of Helios’s mission life (i.e. autonomous). The only

navigation method that best meets these demands is the MANS system.

The Micorcosm Autonomous Navigation System (MANS) can be utilized to supply

accurate ephemeris data that include Helios’s position, velocity, attitude, attitude rate, ground

look point, and lighting conditions with inexpensive fully autonomous navigation software. This

satisfies more than all of the needs the Helios Mission requires to achieve maximum productivity

and operability. The MANS system has been flight tested and been proven to work reliably. 33

The accuracy of the MANS system tested met the requirements of determining the spacecraft’s:

position error with a three sigma error of 400 meters, velocity with a three sigma error of 0.4

meters per second, attitude (the angle about each axis) with a three sigma error of 0.05 degrees,

and attitude rate (the angular rate about each axis) with a three sigma error of 0.005 degrees per

second. Accuracy is predicted to be able to increase with the addition of more sensors or moreaccurate sensors that should be available since this test was performed in March 1994. The

MANS system can also incorporate Global Position System receivers (which are known to be

accurate from 15-100meters), gyroscopes, accelerometers, and star sensors to also improve

accuracy. However, use of extra sensors have been deemed unnecessary for completing the

mission and would only draw our maximum power output. Figure 3.14 is a photograph of two

scanners and their electronics units. Table 3.2 describes their weight, dimensions and power

usage.



33

Figure 3.14 – Photograph of the Dual Cone Scanners and their Electronic Units

Table 3.2 – Cone Scanner and Electronic Unit

Scanner Electronics Unit

Weight (lbs) 3.4 8.8

Dimensions (inches) 3.875 x 3.875 x 9.15 11.2 x 5.1 x 4.3

Power (watts) 14 14

In conclusion, all of the data the MANS system obtains can be utilized towards

maximizing sun exposure, keeping Helios on the desired orbit path, and providing accurate

enough information that the power transmission remains in constant contact with the ground

station to avoid losing any of the directed power. Estimates given by the power and

communications subsystem indicate the ground receiver’s size could be somewhere between 1-

10 kilometers, so optimally the core MANS system with an extra GPS receiver, the worst case

scenario will be off 15 meters giving 0.15-1.5% error. In addition the MANS system does this all

autonomously so time is not lost sending, waiting, and receiving information from the ground

station.



34

The Guidance Navigation and Control subsystem has decided on a Molniya orbit path,

using VASMIR ion-thrusters for reaction and attitude control, and using the Micorcosm

Autonomous Navigation System (MANS) to gather information on Helios’s orbital parameters.

This system will be able to autonomously correct Helios’s attitude and orbit path using the

variable thrusters. Future work is required in a number of areas within the GNC subsystem. First,

future work in the reaction control system includes determination of the specific design

placement of the VASMIR thrusters. Second, calculations need to be made to determine the

power needed on the satellite, as well as the time required between potential fuel resupply

missions, to support the VASMIR system. Finally, future work of the MANS entails

implementing the sensors into the structural design.

3.7 Command and Data Handling

The Command and Data Handling subsystem is in control of coordinating power

transmission and monitoring the status of Helios. In order to maximize the efficiency of the

power transfer and maintain operational capabilities, keeping and storing data on power output,

orbital attitude, orbital trajectory, and thermal status of electronic components. This subsystem is

closely related to the Communications subsystem and the Guidance and Navigation Control

subsystem.

Computers are important for computing the data from other subsystems, mainly the

Guidance and Navigation Control subsystem in Helios. For the Helios mission an onboard

computer that can support the MANS system is needed to be reliable and accurate as proposed

by the Guidance Navigation and Control subsystem.

The computer used in the original MANS system flight test was a Honeywell Generic

Very High Speed Integrated Circuit Spaceborne Computer (GVSC). Table 3.3 describes the

specifications of the GVSC used in the original flight test. In accordance with Moore’s law,exponential advances in computer technology could provide a GVSC today that is 18 times as

fast and has 18 times as much memory. Today’s GVSC would also be lighter and cheaper; an

exact price could not be found for the original Honeywell GVSC, but an unreferenced estimate

placed the pricing between 300-500 thousand dollars.



35

Table 3.3 –Honeywell Space Computer 33

Honeywell Space Computer

Name and Function Honeywell Space Computer

Company Honeywell Inc., Space SystemsWeight (lbs) 10.6 lbs

Dimensions (mm) 203 x 262 x 70

Power (W) 5.0

Processor MIL-STD-1750 A

RAM Size 512 k words SRAM, I/O buffer with 8k words SRAM

Data Interfaces I/O: MIL STD 1553 B, 16 bit parallel, serial I/O (Duplex)

Radiation Hardness 100 kRad

Cost Range $300K to $500 K

Other than the MANS system other computational loads that will be supported include

the Thermal subsystem thermal regulation data, power efficiency and generation, and the

autonomous addition of power modules. As there is no science subsystem, the Helios Mission’s

computational needs will be well exceeded by today’s advanced computer technology.

In conclusion, a modern a Honeywell GVSC will well exceed the computational needs of

all the subsystems and will be well protected from the environment. To avoid computational

faults, active redundancy methods, such as Standby Sparing and watchdog timers, will be

implemented into the system. The computer will also require battery power for the short period

of time when Helios is out of the sunlight in its Molniya orbit. Future work in this subsystem is

focused on obtaining specific technical data and pricing from Honeywell’s current computer

products whose specs are only given to potential customers.

3.8 Power

The power subsystem has the main requirements of generating, storing, regulating, and

distributing electrical power to all of the various instruments and equipment in all other

subsystems aboard the spacecraft. Helios is a special case, as its primary mission requirement is



36

to generate and transmit as much power as possible back to ground on earth, therefore certain

calculations that deal with individual components within the power subsystem have become

obsolete. Tasks such as making sure the spacecraft is able to supply energy to all onboard

operational components are not needed due to the immense amount of power Helios will be

generating. The requirements of the power subsystem can be broken down into the following

definitions. 51

• Power Source (Solar Photovoltaic or Solar Dynamic)

• Energy Storage (Primary and Secondary Batteries)

• Power Regulation and Control (AC-DC converters, Voltage Range, Peak Power)

• Power Distribution (Wiring, Load Switches, Fault Protection)

The power generation source has been narrowed down to a solar dynamic system(Brayton cycle solar turbo alternator), which focuses the suns energy in the form of heat and

converts it to electrical power. The main purpose of the Helios mission is to generate as much

clean, renewable energy from the sun as possible, and it will now utilize several individual large

scale Brayton cycle solar turbo alternator power “modules” that will plug into each other and the

rest of the Helios satellite as they are individually launched into orbit. Due to the solar renewable

energy requirement, and due to the large amount of power needed, radioisotope thermoelectric

generators (RTGs), nuclear reactors, and fuel cells are not options for power generation on

Helios. Solar photovoltaic cells, which absorb sunlight and convert it directly to DC electric

current, are also no longer a suitable option for Helios. The most efficient Triple Junction

Gallium Arsenide solar cells in production today have a lab efficiency of 33.8%, meaning they

can at maximum produce 462 watts per square meter in orbit around earth. 51 These triple

junction GaAs cells weigh in at approximately 2.8 kilograms per square meter, giving them a

specific power of only 165 watts per kilogram. These cells are the best cells available and are the

only ones that were considered for Helios, as other cell materials like Silicon and Indium

Phosphide have significantly reduced efficiencies. To generate electric power for transmit toearth in the multiple megawatt range, thousands of square meters of solar cells will be needed,

the largest arrays ever built, and generating 10 MW of power using GaAs cells will require over

60600 kg (66.8 tons) of solar panels to be launched into orbit. This is not cost-efficient compared

to a solar dynamic system and currently, the world production of solar cells is inadequate to



37

supply them at sufficient amounts. 12 Due to the very low specific power (power per mass) of

solar photovoltaic cells compared to a solar dynamic power generation system, solar cells were

eliminated as a possible option.

A solar dynamic system, more precisely, a closed Brayton cycle solar turbo alternator,uses the same Brayton heat cycle as in all gas-turbine power plants and generators on the surface

today, as well as in numerous other applications such as aircraft jet engines. It is therefore a very

mature technology, with current efficiencies nearing 41.8%, and efficiencies at the onset of the

Helios mission being launched could exceed 47.0%. 44 With this large of an industry already

existing, the infrastructure and knowledge available to produce these solar turbo alternators at

high volume is available and is already far above that of solar cells. 12 The solar turbo alternators

would be constructed and launched to orbit in individual self-contained solar dynamic modules

(SDMs) consisting of a deployable solar concentrator (thin film rigidized inflatable Fresnel lens

structure), a solar heat receiver, the turbo alternator for heat conversion to AC electrical power,

and deployable heat radiators (rigidized inflatable). A schematic of an SDM module can be seen

in Figure 3.15.

Figure 3.15 – Solar Dynamic Power Module Layout44



38

The Brayton cycle conversion of solar heat to electrical power utilizes an inert noble

working gas that is heated and expanded through a turbine, which then turns an alternator to

generate AC electric current. 44 The hot gas is then re-pressurized in a compressor and cooled

through a heat exchanger that radiates the excess waste heat into space before it returns to the

solar heat receiver. 44 A schematic of the Brayton cycle is shown in Figure 3.16.

Figure 3.16 – Brayton Cycle Schematic 12

The thin film rigidized inflatable Fresnel solar concentrators will be coated in Silver to

increase their focal efficiency to nearly 92%. 44 The solar heat receiver at the focal point of the

Fresnel sunlight concentrator will consist of multiple Lithium heat pipes arranged in an inverted

conical pattern to dissipate any reflection and coated with high-temp Graphite to improve their

heat absorption. 36 The turbo alternator itself will consist of a ceramic bladed turbine with a

working fluid inlet temperature of 2000k connected to an AC electrical alternator generating

10kV, and then to a compressor with a temperature and pressure ratio of 4.5 and 2.6

respectively. 44 And finally, the waste heat radiator will consist of rigidized inflatable Carbon-

Carbon panels coated in graphite and will need to be an active pumped loop system. 44 A cut-

away model of a Brayton cycle turbo alternator module (seen without a solar concentrator or

radiator) can be seen in Figure 3.17, this 10MW generating module is 3 meters in diameter by 10meters long and weighs 14.1 tons (including solar concentrator and radiator). 36, 44



39

Figure 3.17 – Brayton Cycle Turbo Alternator Model 36

The higher conversion efficiency of SDMs when compared to solar cells means that the

overall solar collection area can be much smaller for the same power output, and the large

inflatable thin film concentrators have a very low mass per square meter, significantly reducing

the weight of the entire structure and increasing its specific power. Current technology permits

for SDMs in the power output range of 10 MW per unit, with a deployed solar concentrator size

of 163 meters in diameter, a deployed radiator surface area of 1394 square meters, and a mass of

12700 kg (14.1 tons). 44 This gives a specific power of 787 watts per kilogram, compared to the

low 165 W/kg with GaAs solar cells. Using currently produced smaller power output ground

based SDMs as a reference, and scaling the power generation and design changes for a space

application, an estimated cost of each 10MW SDM module is around $50 million in 2011 36.

Using this initial estimate, the cost of generating 200MW of solar power would be $1 billion plus

launch costs for getting 255826 kg (282 tons) (20 SDM modules) into orbit. This cost is not

unreasonable and is much cheaper than using solar photovoltaic power generation. Additionally,



40

the SDMs would have much longer life components due to their inherent space environment

tolerance (no radiation effects). Space debris will still play a role, as damage to the large solar

concentrator and radiators will reduce efficiency. However, due to mission requirements of

generating the maximum amount of power possible on a $21 billion budget, the beginning-of-life

and end-of-life power calculations of the SDMs are not required, as the immense power

generated by Helios will always be able to supply all other critical on-board subsystem

components. This efficiency degradation of the SDM concentrators and radiators will only cause

power transmitted to ground to decrease as mission years pass. Additionally, the solar turbo

alternators will supply power to the spacecraft’s power bus at very high voltage (10,000 V) using

three-phase alternating current (AC) at 1 kHz. 12 This high voltage AC current source will

significantly reduce resistive losses and improve overall power system efficiency across the

spacecraft’s very long length (93% efficient), also allowing for simpler wiring systems and veryreduced cable mass when compared to direct current (DC) relatively low voltage solar cells. 44

Energy storage on most spacecraft is extremely important, however due to the nature of

Helios being purely a power generation satellite that will create power nearly 100 percent of the

time (it rarely sees solar eclipse due to a modified Molniya orbit), energy storage will be limited

to just the power that is needed for critical subsystem equipment to operate for 24 hours in case

of solar eclipse, or a major malfunction in the power generation system. Also, Helios will not be

transmitting its generated power to earth at all times due to ground station time accessrestrictions. However, it is extremely impractical to store power on the magnitude of megawatts

within batteries, as the battery mass required to put in orbit would be immense. Therefore, the

Helios will radiate and transmit its excess power away into space, or to any other possible future

third party satellites and space stations in Earth’s orbit, when not in view of the Alaska ground

station. Lithium ion batteries were selected due to their energy density being the highest of all

space-rated batteries at 125 W-hr/kg. 51 The exact mass of batteries needed to keep the satellite

operational (Sensors, Comms, and on-board data handling computer) in case of eclipse or

malfunction will be calculated at a future date once all other subsystems have selected their exact

components and the power load they require. Currently, the Honeywell on-board data handling

computer and dual sensor arrays require 32W-hr, requiring 6.15 kg of battery mass to keep

powered for up to a day. Additionally, a thermal energy storage (TES) system will be added to

the SDMs. The TES system will consist of Lithium Fluoride-Calcium Difluoride (LiF-Ca2F)



41

phase change material contained in canisters surrounding the turbine inlet gas flow tubes. 44 The

LiF-Ca2F will remain hot as it transitions from liquid to solid, transferring heat to the working

fluid during any solar eclipse in order to allow continuous power generation. 44 The Li-Ion

batteries will serve as a redundant power backup system in case of malfunction in the Helios

PMAD (power management and distribution) system.

Future work for the power subsystem includes determining the exact power distribution,

regulation, and control units that will be needed in order for all other spacecraft subsystem

components to operate off of the SDM supplied 10 kV AC current at 1 kHz (voltage

transformers, AC-DC converters, etc). This includes determining the exact voltages and currents

the slotted waveguide microwave transmitter needs supplied to it for the most efficient power

transmission to earth, as well as what voltages and currents other critical spacecraft equipment

needs to operate on. Additionally, the exact power load that all critical satellite operational

equipment requires will need to be determined, in order to calculate the mass of Li-Ion batteries

that will need to be on board for a 24 hour long redundant power backup in case of eclipse or

malfunction.

3.9 Thermal Control System

The thermal control subsystem is responsible for generating accurate thermal models of the spacecraft, determining operating temperatures for specific equipment, and making decisions

regarding the implementation of passive and active thermal control on board the spacecraft.

Additionally, the subsystem is responsible for testing the payload under the expected

environmental conditions to confirm that the electrical components composing the payload stay

well within their operational and survival temperature limits during the mission. Taking into

consideration the mission architecture and objectives as well as the harsh environmental

conditions the payloads will experience in the modified Molniya orbit, requirements for the

thermal control subsystem were generated. These requirements are listed in Table 3.4. The

requirements remain unchanged since the original generation of requirements for the mission.



42

Table 3.4 – Thermal Control System Requirements

Requirement Number Short Name Description

THM 1 Monitoring All critical components shallbe monitored by C&DH viathermal sensors.

THM 2 Passive Control The thermal system shallregulate the temperatures of critical components viapassive thermal controlwhenever possible tominimize the power consumedby the thermal system.

THM 3 Active Control The thermal system shallregulate the temperatures of critical components via activethermal control when passivecontrol is deemed ineffectivefor regulating.

THM 4 Modeling The thermal subsystem shallgenerate accurate thermalmodels of the payload in orderto justify decisions made bythe subsystem.

THM 5 Testing The thermal subsystem shalltest the payload at thesubsystem and system level toensure survivability of thepayload in orbit.

THM 6 Gradient The thermal subsystem shallattempt to minimize thethermal gradient across thespacecrafts whenever possible.

The following subsection of the report details the work that has been completed since thelast report towards satisfying the requirements described above of the thermal control subsystem.

It also identifies the future work that needs to be completed by the thermal control subsystem.





44

Table 3.5 – Temperature Sensors

Type of Temperature Sensor Model Number(s)

Standard Surface Capsules 29230, 29309

Miniature Surface Capsules 29222, 29223Stikon RTD 22810

Strapon RTD 22391, 22392, 22393

Wide Range Thin Film Strapon 22525, 22526

Wide Range Wire Wound Strapon RTDSensor

22488, 22489

Low Cost Capsule 29399

Stikon Thermocouples 26721, 26722, 26723, 26724, 26893, 26894,

26895, 26896Foil Thermocouples 2010

Strapon Thermocouple 26391, 26392, 26393

Measurement Specialties Inc. produces five different types of temperature sensors. They

are the 44900 Series Thermistors, ESCC Epoxy Coated Thermistor Series, ESCC Surface Sensor

Series, Glass 55000 Series, and the Thin Film Sensor. 56 A trade study will take place featuring

the fifteen sensor types discussed above to select the type of sensor used onboard Helios. The

sensor types will be scored against the metrics listed below in Figure 3.18. The metrics have

been scored against each other in a pair-wise comparison matrix in order to determine their

effective weights. The weighting took into consideration the objectives of the mission and the

constraints the team has identified for the satellite. The final weights of the metrics are also

depicted in Figure 3.18. In the figure, a score of 0 indicates that the metric in the row is less

desirable than the metric in the column, a score of 1 indicates that the metrics are of equal

importance, and a score of 2 indicates that the metric in the row is more desirable than the metric

in the column.





46

determining if those limits will be approached during the duration of the mission via thermal

simulations. Determining which sensors will be used onboard includes finishing the trade study

using information from the providers’ websites and component data sheets. Both tasks are

ongoing.

THM 2

Passive thermal control is commonly used on satellite missions to cool components and

to transfer heat to other places on the satellite. Passive thermal control includes tools like

coatings or surface finishes, insulation or isolators, heat pipes, and conduction bars. 67 The

modified Molniya orbit that will be utilized in the mission places the satellites in direct sunlight

for the entire duration of the orbit. It is expected that passive thermal control will beimplemented on the payloads to attempt keep components below their maximum operational and

survival limits. The general passive thermal control system has been finalized. Heat will be

pulled from heat dissipating components via heat pipes. Conduction bars were previously also

being considered for this task; however, their lack of heritage in aerospace applications has

precluded them from being used onboard Helios. The heat pipe components will be wrapped in

multi-layer insulation (MLI) blankets in order to minimize the loss of heat to the system as the

heat from the components are transferred to the radiators. Figure 3.19 illustrates this design.

Figure 3.19 – Passive Thermal Control Design



47

In addition to this general passive control system, thermal coatings will be applied to all

applicable surfaces to minimize the heat absorbed by the payload, but maximize the heat radiated

away from the satellite. Table 3.6 lists the companies that are being considered for supplying

specific portions of the thermal control system. Each of the companies specializes in the type of

thermal control they would supply to Helios and have produced that specific component for the

aerospace industry in previous space missions.

Table 3.6 – Passive Thermal Controls Available

Type of Passive Thermal Control Company

Heat Pipes Thermacore 32, Advanced CoolingTechnologies 3, ATK Space Systems 5

Multi-Layer Insulation Dunmore Corporation 20, AerospaceFabrication, RUAG Space 55, ATK SpaceSystems 6

Radiators Thermacore 35, Vanguard SpaceTechnologies 19, Lockheed Martin 1, ATK SpaceSystems 4

Coatings ATK Space Systems 7, Tiodize Technologies 22

The most desirable options from each company will be put through a trade study using a

set of predefined metrics. The metrics and respective weights for heat pipe, MLI, radiator, andcoating selections are displayed in Figure 3.20, Figure 3.21, Figure 3.22, and Figure 3.23,

respectively. Weighting of the metrics was completed in the same fashion as the temperature

sensor trade study metrics. The determination of which components will feature passive thermal

control is ongoing as other subsystems finalize their major component selection and when

preliminary thermal models of the satellite are generated. The specific design of the passive

thermal control system will be finalized once the structure is defined.



48

Figure 3.20 – Heat Pipe Metrics and Weight Criteria

Figure 3.21 – MLI Metrics and Weight Criteria



49

Figure 3.22 – Radiator Metrics and Weight Criteria

Figure 3.23 – Coating Metrics and Weight Criteria

The solar turbo alternators’ design features a radiator with a total surface area of 3000

square meters. The radiator will be of the same type decided using the radiator trade study. The

pumped fluid loop of the alternator design will also be decided by a similar trade study as a part

of requirement THM 4. Because of the need for operational thrusters onboard the satellite andbecause of the irreversible damage to the thrusters if they overheat, they will feature a thermal

coating determined by the coating trade study. The addition of other passive thermal control

methods or active control methods for the thrusters will be considered when a thermal model of

the payload is completed.



50

Outstanding work related to this requirement includes determining which components

will feature passvie thermal control and finalizing the passive thermal control design.

Determining the components with passive control includes determining which componenets will

overheat during the mission via thermal simulations as the components are selected. Finalizing

the design entails completing trade studies for the selection of the system heat pipes, MLI,

radiators, and thermal coatings.

THM 3

Active thermal control is used when passive thermal control cannot effectively regulate

the temperature of components. As active thermal control uses power and could potentially place

a burden on the flight computer, the thermal subsystem desires to minimize its presence on thesatellites; however, it will most likely be needed. The extended period of sun exposure of the

modified Molniya orbit will make active thermal control a necessity. That does not diminish the

need for passive control, however. Passive control will act as a redundancy to the thermal system

if a central computer or power failure was to occur. Previously, it was discussed that control can

occur via the flight computer or via independent operating systems. The thermal subsystem has

decided to control active cooling via the flight computer. The flight computer chosen for Helios

has enough computing power to handle the thermal systems as well as all other critical functions.

By using one central computer the thermal subsystem utilizes less mass than what would be

required of independent operating systems. With this architecture thermal sensor data will be fed

directly to the flight computer and all decisions will be made by the software.

Active thermal control includes items like heaters, louvers, and cryocoolers. The orbit

selected for this mission places the satellites in direct sunlight for an extended period of time. For

this reason the thermal subsystem will not use heaters to regulate temperature. Additionally,

cryocoolers will not be used to cool components as no components will need to be cooled to cryo

temperatures. However, if heat loads are expected to exceed the heat transfer capabilities of

passive thermal control, closed-loop chiller lines will be used to accelerate heat transfer to the

radiators on the satellites. 2 Figure 3.24 30 displays a simplified schematic of how the fluid lines

would work onboard Helios. The proposed passive control system would work in a similar

fashion. Crane Aerospace & Electronics has been selected to develop the pumped fluid lines of



51

the active thermal control system. Research has shown that of the limited number of companies

that produce these lines for aerospace applications, Crane has the heritage desired by the Helios

team. Their fluid lines have flown on various space missions including every Mars rover

mission. Crane currently produces twelve different fluid solutions. 17 The twelve different

methods will be compared to determine which solution is most applicable to aerospace payloads

and the Helios mission specifically. As with the passive control system, determinations about

which components or systems will feature any type of active thermal control will be made as

other subsystems choose components and thermal simulations are generated.

Figure 3.24 – Example Fluid Lines Schematic

Outstanding work related to this requirement includes determining which components

will feature active thermal control and finalizing the active thermal control design. Determining

the components with active control includes determining which componenets will drastically

overheat during the mission via thermal simulations. Finalizing the design entails completing

trades between the fluid solutions offered by Crane using research from the company website

and data sheets.

THM 4

As discussed above, accurate thermal models are needed to provide justifications and

make decisions regarding the thermal control system. Radiant energy heat balance calculations

are good for making preliminary decisions regarding the control system; however, refined

models will guarantee mission success by thermal standards. As the structure is defined



52

simulations will be generated using STK SEET or COMSOL Multiphysics to not only make

thermal control system choices but to also define the thermal testing profile the payload will

ultimately be subject to. As there is currently no decision regarding the size or complexity of the

payload, the decision on whether to use STK or COMSOL for simulation purposes cannot be

made at this time. The decision rests solely on the complexity of the model and its ability to be

defeatured. STK SEET assumes that the spacecraft is one node and the program calculates the

temperature of that node in orbit based on the orbital parameters. COMSOL generates a multi-

node, time-dependent solution. The COMSOL simulation would be desired; however, COMSOL

cannot accept incredibly complicated geometries. Thus, the decision between the two programs

will be heavily influenced by the structure’s complexity.

The modified Molniya orbit Helios will be in gives the satellite a perigee in low Earth

orbit (LEO) and an apogee in medium Earth orbit (MEO), the altitude range between the end of

LEO and the geosynchronous orbit (GEO) altitude (36 kilometers). There are three different

types of heating Helios will experience in this type of orbit. They are direct solar radiation, Earth

IR emissions, and radiation energy from the Sun reflected by the Earth (Albedo). The range of

direct solar radiation incident on any surface of Helios during any point of its orbit is constant at

1414 W/m 2 to 1323 W/m 2. This is due to the fact that the change in altitude of the orbit is

insignificant compared to the distance of the Earth from the Sun. The Earth IR and Earth albedo

incident on the surface of the satellite in orbit, though decreases with increasing altitude. At LEOaltitudes the range of Earth IR and Earth albedo experienced is 241 W/m 2 to 227 W/m 2 and 438

W/m 2 to 384 W/m 2, respectively. At GEO altitudes the range of Earth albedo experienced is 7.19

W/m 2 to 2.54 W/m 2. Earth IR is constant at 5.52 W/m 2 at GEO altitudes 21. Taking into

consideration this information, the worst case hot and worst case cold conditions occur at perigee

and apogee, respectively. Table 3.7 summarizes the heat contributions to Helios at these points.



53

Table 3.7 – Heat Contributions

Worst-Case HotPerigee (Alt: 1000 km)

Worst-Case ColdApogee (Alt: 23622 km)

Direct Solar Radiation 1414 W/m2

1323 W/m2

Earth IR Emissions 241 W/m 2 < 5.52 W/m 2

Earth albedo 438 W/m 2 < 2.54 W/m 2

Outstanding work related to this requirement includes determining which software

package to complete simulation work with, generating a defeatured model of the spacecraft, and

generating a thermal model taking into consideration the environment influences.

THM 5

Thermal testing is needed in order to verify choices made regarding the passive and

active control systems. Additionally, it is needed to verify the payload’s survivability during the

full duration of the mission. The thermal subsystem will implement thermal-cycling and thermal

burn-in testing on electrical components utilized on the satellites. Additionally, the subsystem

will implement thermal-vacuum testing and thermal balance testing on the full system. The

duration of the tests and requirements of testing typically follow a military standard foraerospace payloads. The Helios mission will use military specification document, MIL-STD-

1540B, to generate all testing requirements. Aspects of the temperature profile will be

determined based off of the thermal simulations’ and the expected temperature ranges to be

experienced in orbit.

Outstanding work related to this requirment includes generating testing requirements for

the payload based off the thermal simulations projected temperature ranges and military testing

document, MIL-STD-1540B.

THM 6

The goal of any thermal subsystem is to keep the active electrical components working

and to additionally minimize the thermal gradient across the satellite As the thermal subsystem is



54

satisfying THM requirements 1 thru 4, it was also look to minimize the thermal gradient across

the sun facing and non-sun facing sides of the payload whenever possible. The refined thermal

simulations will be the main indicator of success of this goal.

There are several decisions that need to be made in the future regarding the thermalsubsystem; however, the framework to make these choices, illustrated above, will ensure speedy

selections of key aspects relating to the thermal management of the payload.

3.10 Payload

The payload subsystem is responsible for satisfying the primary objective of the mission.

The subsystem has to develop the mechanism or method for transmitting energy to the Earth.

The payload subsystem has derived the following requirements related to power transmission:

• The payload shall transmit energy captured by the satellites to the specified ground

station on Earth.

• The payload subsystem shall design the payload to fit the specifications of the launch

vehicle and structure as defined by the structures subsystem.

• The payload subsystem shall select the method that maximizes the amount of energy

received on Earth.

• The method of energy transfer shall be safe and shall not negatively impact the

surrounding communities to the ground station or the surrounding environment to the

ground station.

Preliminary research shows that there are very few ways to transmit energy from space to

Earth wirelessly. Additionally, research shows that microwave technology is the only feasible

way of satisfying the requirements above.

Laser Transmission

One way of transmitting power from space wirelessly is through laser power beaming.

Laser power beaming works essentially like solar cells on spacecraft. When the sun shines on a

solar array, photovoltaic cells convert the sunlight into energy. With laser power beaming



55

photovoltaic cells convert laser light into energy. Benefits of wireless power beaming include

compact size, zero radio frequency interference, and greater energy concentration at long

distances. Laser power beaming technology has been studied conceptually by NASA and the

Department of Energy since the 1970s. 38 Over time NASA looked to make the idea of space-

based laser transmission more feasible; however, the main limiting factor has been the cost of

laser diodes. Studies suggest that at their current price the gains of a space based power system

would be lost to the cost of the diodes. Cost in itself is what creates the issues associated with a

space-based laser power beaming method of transmitting power. Because the cost is so high,

laser beaming designs have not moved from a conceptual idea to a realistic, tangible system over

time; thus, the subsystem believes that the technology will not be ready to be implemented

during the development phase of the Helios satellite constellation. Another issue with laser

beaming technology is that is hasn’t been tested over long distances. A 2009 NASA competitionthat had design teams develop systems for beaming power via a laser only resulted in a max

distance of transmission of 1 km. Professional development has had little success in increasing

that distance. 38 For those reasons power transmission by means of laser beaming will not be

implemented on the satellite constellation. The payload subsystem believes that the issues

explained above regarding laser beaming will not satisfy the requirements designated at the

beginning of this subsection.

Microwave Transmission

The optimal method of wireless energy transfer from orbit is focused microwave

radiation beamed to a rectenna station on the Earth. The energy collected by the system is

converted to microwaves and beamed via a transmission antenna. Various frequencies of

microwave radiation have been utilized in wireless power transmission, each having different

benefits and levels of efficiency. The performance of these frequencies is outlined in Table

3.8.47 The efficiency of the receiving station rectenna also depends on the frequency used, the

primary factor being the type of diode used to receive and convert the microwaves. These

efficiencies are displayed in Table 3.9. 47



56

Table 3.8 – Performance of Printed Rectenna

Type of Rectenna Operating Frequency (GHz) Measured Peak ConversionEfficiency (%)

Printed Dipole 2.45 85

Circular Patch 2.45 81

Printed Dual Rhombic 5.6 78

Square Patch 8.51 66

Table 3.9 – Rectenna Efficiency for Various Diodes at Different Frequencies

Frequency (GHz) Schottky Diode Measured Efficiency (%) Calculated Efficiency (%)

2.45 GaAs-W 92.5 90.55.8 Si 82 78.3

8.51 GaAs 62.5 66.2

Based on the efficiencies and performance levels, a microwave frequency of 2.45 GHz

was determined to offer the optimum performance and transmit the most energy. Inherent

drawbacks to using a smaller wavelength include losses and dispersion that occur due to

atmospheric conditions such as clouds and rain. Choosing the 2.45 GHz frequency minimizes

this issue while maximizing performance; studies have demonstrated that atmospheric loss at this

frequency is less than 1%. 45

There are three potential antenna types that could be used: a parabolic dish antenna,

miscrostrip patch antenna, and slotted waveguide antenna. The best option for mass energy

transmission is the slotted waveguide antenna due to its relatively low cost, high aperture

efficiency (~95%), and high power handling capacity. 15 A sample slotted waveguide antenna is

seen in Figure 3.25.



57

Figure 3.25 – Slotted Waveguide Antenna

A basic block diagram of the transmitter-rectenna system is displayed in Figure 3.26. 47

Figure 3.26 – Transmitter-Rectenna System Block Diagram

Future developments for the payload subsystem include selection of individual

components of the microwave transmitter, including the coax-waveguide adapter, waveguide

circulator, and tuner/directional coupler. Trade studies will be conducted to determine these



58

components based on cost, efficiency, and availability.

Environmental Effect and Safety

OSHA standards for microwave exposure are set at 10 mW/cm2

. Approximating thepower transmission levels at 23 mW/cm 2 at the center of the rectenna, and assuming receiver

with a diameter not less than a few kilometers is used, concentrations outside the edge of the

receiver dish would be on the order of 1 mW/cm 2. This complies with the OSHA standard, and

also the ANSI and IEEE standards. Concentrations of this magnitude are on the same level of

magnitude as cell phone radiation, and pose no threat to the environment even over prolonged

periods. Aircraft passing through the beam would be protected by the aluminum shell (Faraday

Cage Effect), and airspace control would prevent smaller craft and balloons from encountering

the beam. Fencing would prevent access to the facility. The fear of being “cooked” by the beam

is almost entirely irrational and could not happen unless one was to lie at the center of the

rectenna for a very long time.

4.0 Future Work

During the next semester the major budgets and estimates of the mission will be

established. The power estimate will be generated by summing the power requirements of all thespecific components selected by each subsystem. Because this mission is a power generation

mission, the team does not anticipate the need to limit the amount of power supplied to each

subsystem. The cost of the mission is fixed at 21 billion dollars; however, next semester the team

will allocate those funds to different parts of the mission including launch operations and flight

operations. Additionally, funds will be allocated to individual subsystems as well. Finally, a

mass budget will be created, which will allocate the total mass allowable by the selected launch

vehicle to each subsystem. The subsystem will decide how they use that allocated mass and a

mass equipment log will be generated featuring the breakdown of how the total mass is being

used.



59

5.0 Conclusion

The Helios team will design and construct a space-based power system (SBPS) capable

of gathering energy from the sun’s rays and wirelessly transmitting the collected power back to

Earth as a sustainable form of power production. This system will include a satellite launchedinto a modified Molniya orbit at an inclination of 63.4° and a ground receiver rectenna station in

Alaska. The Molniya orbit will allow for maximum exposure to the sun for energy collection and

greatest window for power transmission back to the ground station, while being closer in

distance than GEO. The large Helios satellite will be modular in design, and each module will

be launched from a high latitude ground station to facilitate injection into the Molniya orbit.

This ground station will be located just south of Fairbanks, Alaska, while the satellite

components will be launched from Cape Canaveral, Florida. The first module launched will

immediately begin power transmission, and as subsequent modules are sent up, they will

autonomously dock, plug-in, and begin power generation. The overall satellite structure will be

an expandable FAST truss system, made primarily from Aluminum. The satellite will utilize the

Microcosm Autonomous Navigation System (MANS) to provide guidance and navigation. The

system will run all software on a Honeywell Generic Very High Speed Integrated Circuit

Computer (GVSC), which will also monitor all subsystems and power transmission and

communicate with the ground control station. High gain parabolic antennas are used to transfer

uplink and downlink data directly from Helios to the ground station. The data transmission will

operate at C-band with uplink frequency of 6 GHz and downlink frequency of 4 GHz. The

Tracking and Data Relay Satellite System (TDRSS) will be used to relay uplink commands to

Helios during its orbit around perigee. VASIMR thrusters along the entire length of the satellite

will be utilized to insert the satellite into the modified Molniya orbit following launch, maintain

the orbit, make attitude adjustments, and counteract J2 effects, solar radiation pressures, and

atmospheric drag. Power will be generated by multiple Brayton cycle solar dynamic turbo

alternator modules, each generating 10MW of AC power at 10kV. Thermal control will stemfrom a combination of active and passive systems. Helios’ payload consists of a large slotted

waveguide antenna, which will transmit power in the form of microwaves at a frequency of 2.45

GHz. These microwaves will be of low enough intensity that they will have a minimal impact on

the environment and humans. The ground-based rectenna station will receive the power beamed

from the satellite’s antenna and convert it into usable AC electricity that can be delivered to the



60

ground based power grid. This system will operate over a lifetime of approximately 30 years,

providing sustainable power with minimal environmental effects, and reducing reliance on non-

renewable sources of power generation.



61

6.0 References

1. "Absolute Freedom." Lockheed Martin Heat Rejection Radiators (HRS) . Lockheed Martin

Corp. Web. 28 Nov. 2012.

<http://www.lockheedmartin.com/us/products/HeatRejectionRadiators.html>.2. "Active Thermal Control System." Active Thermal Control. The Boeing Company, 2012.

Web. 7 Oct. 2012.

3. "Advanced Cooling Technologies, Inc." ACT . Advanced Cooling Technologies, Inc., 2012.

Web. 28 Nov. 2012. <http://www.1-act.com/markets/spacecraft-thermal-control.php>.

4. "Alliant Techsystems Inc." ATK . Alliant Techsystems, Inc., 2012. Web. 28 Nov. 2012.

<http://www.atk.com/products-services/heat-pipe-radiators-and-panels/>.


<http://www.atk.com/products-services/heat-pipes/>.


<http://www.atk.com/products-services/thermal-blankets/>.


<http://www.atk.com/products-services/thermal-coatings/>.

8. Antenna Efficiency and Antenna Gain. (n.d.). www.Antenna-Theory.com . Retrieved from

http://www.antenna-theory.com/basics/gain.php

9. ANTENNAS AND FEEDLINE. (n.d.). Welcome to QSL.NET :: Web Services for Ham

Radio operators . Retrieved from http://www.qsl.net/ng3p/haminfo/ham-

tutor/antennas_and_feedline.htm

10. Bergin, Chris. "SpaceX: Falcon Heavy, Falcon 9 Tag Team Set to Share 20 Launches a

Year." NASASpaceFlight.com . NASASpaceFlight.com, 5 Apr. 2011. Web. 25 Nov. 2012.

<http://www.nasaspaceflight.com/2011/04/spacex-falcon-heavy-tag-team-share-20-launches-

year/>.

11. Chandler, C., Hoey, L., Hixon, D., Smigla, T., Peebles, A., & Em, M. (n.d.). Ka-BandCommunication Satellite Antenna Technology. American Institute of Aeronautics and

Astronautics , 729 (154), 3. Retrieved from

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12. Chapman, Philip K. An Isointerial Solar Dynamic Sunsat . Tech. N.p.: n.p., 2008. Print.



62

13. Chapman, Philip K. "Deploying Sunsats." Online Journal of Space Communication 16

(2010): n. page. Web. <http://spacejournal.ohio.edu/issue16/chapman.html>.

14. Clark, Stephen. "Lockheed Martin to Select Delta 4 Rocket for Orion Test." Spaceflight Now .

9 Nov. 2011. Web. 25 Nov. 2012.

<http://www.spaceflightnow.com/news/n1111/10orioneft1/>.

15. Coburn, William, Marc Litz, et al. United States of America. Department of Defense.

Slotted-Waveguide Array for High-Power Microwave Transmission . Adelphi, MD: Army

Research Laboratory, 2001. Web. <www.dtic.mil/cgi-bin/GetTRDoc?AD=ADA387308>.

16. Conway, Edmund. "World's Biggest Battery Switched on in Alaska." The Telegraph .

Telegraph Media Group, 28 Aug. 2003. Web. 25 Nov. 2012.

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68

Appendix A

Elevation Angles of Molniya Relative to Ground station



69

Appendix B

Helios Orbital Elements

Orbit Selected

Approximate overall access time(taken over two weeks)

75.88%

Semimajor Axis 18689.1 km

Eccentricity 0.605216

Period 25426.8 sec (7.063 hours)

Inclination 63.4 °

Argument of Perigee 270 °

RAAN145.86 °

Radius of Apogee 30000 km

Radius of Perigee 7378.14



70

Appendix C

Access Graph of Modified Molniya Orbits



71

Appendix D

Comparison of Modified Molniya Orbits

Semi Major Axis (km) Period (hours) Eccentricity Overall access percent13800 4.4815 0.534 60.00%

15800 5.4902 0.590 66.36%

16400 5.8059 0.628 68.35%

17500 6.3997 0.628 69.89%

18100 6.7320 0.647 71.51%

19100 7.2973 0.666 73.25%

20200 7.9370 0.685 75.24%

21500 8.7149 0.703 77.10%

22900 9.5801 0.722 78.84%

26553 11.961 0.740 78.45%

Documents

Space Based Solar Power Mission