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Advanced Satellite Bus Technology for Near-Term Solar Sail Missions

Vaios J. Lappas1

Surrey Space Centre, Guildford, Surrey, GU2 7XH, UK

Bong Wie2

Arizona State University, Tempe, AZ, AZ 85287-6106, USA

Sailcraft come with the significant promise to propel spacecraft using photons reflecting on large reflective surface. Paramount to the success of these spacecraft is the need to miniaturize subsystems as much as possible in order to maximize the solar sail’s performance characteristics. The paper examines the availability of advanced small satellite technology and presents an analysis on utilizing miniature small satellite bus technology for near-term solar sail missions. A satellite bus Attitude Control System is presented and the problem of angular momentum dumping due to the constant solar radiation disturbance is analysed. A magnetic control logic is presented and simulation results show the ability to successfully control excessive angular momentum build up to a small band of values using small satellite size magnetorquers, thus making substantial mass savings to the sailcraft design.

I. IntroductionEVELOPMENTS in microelectronics, solar sail technologies and small satellite missions have brought new levels of confidence that near-term solar sail missions can be achieved with significant niche science returns.

The paper examines the technical requirements and engineering solutions for near term solar sail missions. Due to the need to develop the challenging and critical solar sail technologies in the past, miniaturisation of the satellite bus, which is critical to achieve the required solar sail acceleration, has been underestimated. Using the mathematical description for the solar sail acceleration:

P

m

PA

m

Fac max (1)

Where P = 4.536 x 10-6 N/m2 is the SRP constant of is the thrust coefficient, Fmax = PA is the solar sail maximum thrust, = m/A is the areal density and m is the total mass of the solar sail including the mass of the bus (mbus). Clearly Eq.1 indicates the importance of having the smallest mass for the spacecraft bus. Solar sail studies mostly focus on the solar sail aspect of the mission, however even with the availability of small satellite platforms there is a need for mass and systems design optimisation. References [1-3] present the fundamentals of solar sails as well as the mission design of some proposed solar sail missions. Wie [4] presents the fundamentals for the attitude control and dynamics of various solar sail control schemes and designs for recent NASA solar sail proposals. Reference [5] presents the failed Cosmos solar sail design. Other studies [6-12] also present various solar sail/platform designs for near term and longer-term solar sail missions. There is a recent emphasis from US and European institutions to develop mission concepts for a solar sail demonstration mission in the near term. Wie [13] details a systems design for a proposed solar sail mission for NASA’s ST-9 program, which will be used as the benchmark to propose small satellite bus technologies, which can enhance the mission further. The paper will examine some of the system design issues and trades for two different and diverse near term solar missions (i) a 40 x 40 m solar sail demonstration mission in a 1600 km sun synchronous orbit (ii) a 160 x 160 m Solar Polar Imager mission being analysed for a 2012 launch which will travel to an orbit close to 0.48 AU about the Sun. The goal is to derive system mass, power and cost estimates for the proposed two solar sail missions mentioned above as well as any potential critical design issues that might arise with respect to the spacecraft bus that might have an effect on the

1 Assistant Professor, Surrey Space Centre, University of Surrey, Guildford, GU2 7XH, United Kingdom2 Professor, Arizona State University, Tempe, AZ 85287-6106, USA

D

AIAA/AAS Astrodynamics Specialist Conference and Exhibit21 - 24 August 2006, Keystone, Colorado

AIAA 2006-6179

Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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overall solar sail design. A focal point of the paper will be to examine the trade of angular momentum unloading of the reaction wheels used for solar sail control via magnetic torque-rods or/and thrusters.

II. Mission RequirementsReferences [13, 14] detail the two candidate solar sail mission requirements and are summarised in Table 1.

Requirements ST-9Lifetime 3 months ( + 3 option)

Orbit SSO (1000 km)Payload (power,

mass, data)82 W average, 152W peak, mass 69 kg, data rate: 27.5 kbps average, 2.65 Mbps

peakAOCS Pointing control: ~ 0.5°

Pointing stability: ~ 0.5°Pointing knowledge: ~ 0.1°

Ix = 8000 kgm2, Iy = Iz = 4000 kgm2

Propulsion 400-650 m/s hydrazineLaunch 2009

Table 1: Solar Sail Mission Requirements

Table 2 presents a preliminary solar sail design for different sail configurations based on ATK’s scalable sailcraft designs [8, 13]. The 40 m sail design fits the ST-9 mission requirements and the 160 m sail fits the SPI requirements. Table 2 will be the main source for the solar sail parameters for a system engineering analysis on the small satellite bus to be used on both sail mission studies and also on the bus ACS design. For the bus design using a conventional design philosophy 150 kg and 250 kg bus is needed for the ST-9 and SPI missions whereas using a more aggressive design philosophy 50 kg and 100 kg values are presented. Although these values are still conceptual ST-9 and SPI have very diverse requirements which lead to subsystem requirements which become the key drivers for mass and power to be analysed in the following sections.

III. Spacecraft Bus Design

Miniaturisation and the availability of commercial off the shelf (COTS) components have brought a revolution on spacecraft design. Smaller and more capable platforms can be used to complement large conventional space missions. As mentioned before a small and compact small satellite bus is instrumental on many interplanetary/science missions for launch and cost purposes. But miniaturisation of the satellite bus is more critical for solar sail missions due to the total specific acceleration needed. Considering the diverse solar sail mission requirements developing a small satellite bus is not trivial. Driven subsystem drivers (propulsion, ACS, thermal) will lead to different designs.

A. Space Technology (ST)-9 Solar Sail Bus Design

Surrey has developed in synergy with its commercial enterprise a successful family of spacecraft platforms for a variety of space missions with a significant in-orbit track record of 25 spacecraft launched in orbit to date. The ST-9 bus requirements listed in Table 1 are those of a classical micro-satellite mission with the exception of the propulsion requirement of a hydrazine 500-650 km/s system for orbit circularisation due to the launcher selection.For the ST-9 mission, the Pegasus XL is the base-lined launch vehicle. The ST-9 Solar Sail requires a liquid fuel (mono-propellant hydrazine) propulsion system capable of providing a total V of ~ 400 m/s (for a 220 x 1000 km initial orbit), or ~ 650 m/s (for a 220 x 1500 km initial orbit). The liquid propulsion system will be used for launch vehicle dispersion correction maneuver, reduction of tip-off rates, perigee raising and orbit circularization, backup attitude control, backup momentum dumping, and end-of-mission disposal via perigee lowering to 450 km. The proposed platform for ST-9 is based on the successful UoSAT-12 minisatellite. UoSAT-12 is a multi-purpose bus launched in 1998 and is currently still fully operational in orbit. With a mass of 320 kg SSTL’s minisatellite bus provides a robust, off the shelf bus that can be tailored for most small satellite space missions.

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a0.25% of the overall sail size is assumed for a nominally worst case.cA nominally worst, maximum disturbance torque for untrimmed sailcraft.d50% of the pitch/yaw solar torque is assumed for the windmill disturbance torque.

Table 2: Mass properties and solar disturbance torques of ATK’s scalable sailcraft [8, 13]

Figure 1: UoSAT-12 Minisatellite

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B. Propulsion Subsystem

It is clear that the traditional SSTL propulsion systems, such as nitrogen, butane and xenon systems will not be suitable for the V required. They all have specific impulse figures much lower than the ammonia resistojet and hence would have propellant mass fractions substantially higher than the 29% of ammonia. Additionally these propellants typically have low storage densities and hence would also take up unacceptable volumes. Thus a hydrazine system needs to be implemented.

Table 3: Hydrazine Propulsion Trade-Off

Figure 2 - Propulsion System Schematic

The hydrazine propellant is stored in 2 propellant tanks. The propellant is fed from either tank through latching valves into redundant branches of thrusters. The thrusters are located around the spacecraft at 8 locations, with prime and redundant units located together in the same module. The thrusters are equipped with series redundant valves. This protects against loss of hydrazine due to valve leakage. It also ensures that there are three mechanical inhibits between the hydrazine and the thrusters, which is a safety requirement at many launch sites. The system is controlled using two independent propulsion controllers, the same design as is used on Jiove-A (ESA Galileo Testbed built by SSTL for ESA). One controller will control the A branch thrusters and the two latch valves which

Type of propulsion system

Advantage Disadvantage

Hydrazine blowdown Relatively simple system Relatively low cost Significant heritage Multiple qualified equipment

suppliers

Hydrazine is toxic, hence increased safety issues

Thrust levels reduce through life

Hydrazine regulated 9.2kg propellant saving over blowdown mode

Consistent thrust throughout life Multiple qualified equipment

suppliers

Additional high pressure tank and regulation equipment.

Extra mass over blowdown system compensates for propellant saving

Higher volume than blowdown More expensive than blowdown

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feed the branch. This ensures that the propulsion controllers are functionally redundant as both hydrazine tanks can be depleted through a single thruster branch. Two 35 l tanks will be used with a 530mm diameter, which are COTS components available and with heritage (Giove-A SSTL/ESA). Three suppliers of thruster have been identified. Each has a nominal thrust level of 1N.

The potential suppliers are: The MONARC-1 thruster from AMPAC-ISP (formerly ARC-UK), 285 of which have been built and tested to

date The 1N thruster from EADS-ST, Germany, as used on Globalstar and similar programmes The 1N thruster from Aerojet, as used on Skynet 4

Table 4 summarises the key components for the proposed ST-9 propulsion subsystem.

Propulsion Subsystem SummarySubsystem DescriptionPropulsion Hydrazine monoprop. system

2 x ~35 litre tanksDesigned for dispersion maneuvers, perigee

raising, attitude manoeuvres and wheel/momentum offloading.

Thrusters 8 x 1N thrusters

Table 4: Propulsion Subsystem Summary

Figure 3: SSTL Proposed ST-9 Propulsion Module

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C. RF Subsystem

The S-band TT&C system comprises two transmitters and receivers with associated antennas. The transmitters are cold redundant and have their own default frequency allocation at switch on, so they can work autonomously. The set-up has omni-directional coverage and includes provision for safety & back-door commanding for emergency operations.

Figure 4 - S-Band TTC Communications Subsystem

D. Power

The ST-9 bus power system uses heritage from previous SSTL power systems, mainly the DMC-1 spacecraft. The power system uses a battery bus topology, and delivers an unregulated 28V bus to the platform and payloads. The solar arrays are connected to the bus via Battery Charge Regulators (BCRs), providing a versatile and efficient method of maximising panel power during sunlight by adapting to array characteristic variations with changing temperature. Direct connection between the battery and the bus ensures maximum efficiency during battery discharge in eclipse. The power system design is modular and can be scaled to meet varying power demands for different missions.

Solar Panels

Developments in solar cell technologies have increased the standard solar cell conversion efficiency to 27%. This has the knock on effect of requiring less solar array area (and hence mass) for the same power level, helping reduce the overall system mass. For this reason, SSTL propose that this GaAs multi-junction cell technology be used on the solar arrays. To meet the OAP requirements the spacecraft has been equipped with eight body mounted solar panels. Lifetime of 18 months was taken into account in the sizing of the solar panels.

Batteries

The selected battery for the mission is from AEA Technology Space. This battery incorporates the 1.5Ah SONY HC18650 hard carbon cell. For ST-9 the battery uses 8 cells per string (in series) and 10 strings in parallel. The battery has internal redundancy by construction. If one of the 1500mAh cells in a string fails, then only one string of cells in the battery is lost. The loss of 2 strings is accounted for in the initial sizing of the battery for EOL.

9.6 kbps

9.6 kbpsS-Band

Rx 0

S-BandRx 1

Direct TC

OBC Software TC

CANbus

Direct TC

TM/TC

S-BandTx 0

S-BandTx 1

38.4 kbps

38.4 kbpsTM

TM

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OVShunt

Battery ChargeMonitor

PDM

Battery

Battery Bus (+28V ±6V)BCR2

BCR3

BCR12

BCR1

SolarArray

Sections

PCMBPCMA +5V

ActivationCircuit

Figure 5: Proposed ST-9 Power System Block Diagram

E. OBDH

The Data Handling system is based on two heritage computers, with a third redundant one which is an upgrade from the heritage. This allows redundancy and high performance whilst preserving heritage. The OBDH is divided into platform and payload sides each of which has its own dual redundant CAN busses. A high-speed bus is provided between the redundant processor module and the two Data Recorders. The main goal of the platform side is to do AOCS control, whilst the payload side will also handle spacecraft housekeeping. Finally, the Dual Redundant receivers provide access to the payload can bus and direct access to Platform OBCs and Data Recorders. For Data storage, Data will be stored in SDRAM devices that have technology heritage on SSTL satellites. Configuration via the on board processor will provide a means of controlling the storage. This will include a mode that provides protection of captured data against SEU. These devices (Hardware Data Recorder or HWDR) are based on SSTL’s previous SSDR units, one of which is provided for additional backup.

PPC OBC386 0 386 1

HWDR 0 HWDR 1

LRTx

Rx 0 Rx 1

Power System

PlatformModules

PayloadModules

HRTx

Imager

LRTx

Platform CAN Bus Payload CAN Bus

High Speed Bus

Figure 6 - On-board Data Handling subsystem

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F. ST-9 Bus Mass/Power Breakdown

Table 5 summarise the mass and power breakdown for the ST-9 satellite bus. As expected the dominating subsystems are those of the propulsion and payload (provided) subsystems. The propulsion subsystem brings a substantial overhead on the power and structure mass figures due to the high V for the Pegasus XL launcher and the requirement for orbit circularisation. This is the main reason for exceeding the 180 kg satellite bus mass target.

Subsystem Mass (kg) Av. Power (W) Peak (W)

Payload 69 82 153

AOCS 16 12 18Propulsion 85 22 35

OBDH/Harness 19 13 15RF 3 4 6

Power 17 1 3Structure 20

Margin (15%) 264 154.1 264.5

Table 5: Proposed ST-9 Bus Characteristics

IV. Satellite Bus based Sail ACS

Reference [13] describes a Solar-Sail ACS design for a sail flight validation mission in a sun-synchronous orbit with similar requirements to the NASA ST-9 sailcraft mission study. The proposed sailcraft ACS design and ACS requirements summarized in Table 6.

ACSRequirements

Pointing control: ~ 0.5°Pointing stability: ~ 0.5° Pointing knowledge: ~ 0.1°Ix = 8000 kgm2, Iy = Iz = 4000 kgm2

Table 6: DDSS Sailcraft ACS Requirements [13]

Attitude stabilization of a fully deployed sailcraft by a conventional ACS of typical small spacecraft is not usually aviable option. Small reaction wheels and a conventional propulsion subsystem of a typical 100-kg class bus areunsuitable and/or inefficient for a fully deployed sailcraft because of its large moment of inertia and its large solarradiation-pressure (SRP) disturbance torque. For example, a 40 x 40 m sailcraft with a nominal solar thrust force of 10 mN and a center-of-mass to center-of-pressure (cm/cp) offset of ±0.1 m has an SRP disturbance torque of ±1mN-m, which is about 100 times larger than that of large geosynchronous communications satellites. A conventionalACS would require large reaction wheels and/or a prohibitively large amount of propellant to counter such a majordisturbance torque acting on a large sailcraft for extended mission lifetimes. Most conventional control systems ofsmall satellites can generate sufficient control torques for solar sail applications, but their momentum storage and total impulse capabilities are to be investigated for solar sail applications [4, 13].

Reference [13] investigates an interesting sailcraft ACS trade-off of implementing a sailcraft bus ACS with 10 m-Nm reaction wheels (RW) and magnetorquers (MT). The focus of the remaining paper is investigating the feasibility of implementing a RW and MT ACS bus based design for the ST-9 and SSFV [13] mission studies.

A. SSFV ACS with RW

References [4, 13] describe the mathematical and theoretical background of a single and three axis dynamics model of a sailcraft implementing RWs. For the following simulations the attitude control model used is [4, 13]:

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(1)

where (Tx, Ty, Tz) are the control torques generated by reaction wheels. The feedback logic used is [4, 13]:

(2)

where e = ψ − ψc, and Kp and Kd are, respectively, the attitude and attitude rate gains to be properly determined. The saturation function is defined as:

(3)

With },2min{ maxeaKL r (4)

Figure 7: A -90 deg yaw maneuver of an untrimmed 40-m sailcraft (with cm/cp offset of 0.1 m) using a reaction wheel for transition to the standby full-thrust mode

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Figure 7 is a -90 degree yaw maneuver of an untrimmed sailcraft with a cm/cp offset of 0.1 m) using a reaction wheel for transition to the standby full-thrust mode. As it can be seen in the figure, the RW angular momentum immediately builds up to an excessive value of -5 Nms in 2 hours due to the constant SRP disturbance of 1 mNm, which requires immediate momentum dumping. For the SSFV mission study the use of magnetorquers is possible due to its 1000 km, 98 degree inclination orbit. The following section describes the implementation of an angular momentum dumping logic for the SSFV [13] sailcraft.

B. SSFV Sailcraft Bus RW Angular Momentum Dumping

Magnetic control has been used in many spacecraft missions. The simplicity, inexpensive hardware and reasonably good attitude control (0.5o to 5o in all axes) makes magnetic control very attractive to use, especially for small satellite angular momentum dumping. 3-Axis magnetic torquers are primarily used for attitude control and momentum dumping of reaction/momentum wheels [15-18].

Interaction between a magnetic moment, M, generated by a spacecraft with the Earth’s magnetic field, B, produces a control torque NM acting on the spacecraft:

BMNM (5)

The direction of M can be controlled on average by a proper sequence of magnetic torquers firings, but the B field vector is dependent on the orbital location. As a result, the torque NM which always is orthogonal to B and M is not necessarily favourable for control of the attitude of a specific spacecraft axis, in certain regions of the orbit. Another drawback of magnetic torquers is that it is possible that a desirable control torque for a certain attitude axis (pitch, roll, yaw), might generate undesirable disturbance torques for the other axes.

The Earth’s magnetic is predominately a magnetic dipole. The magnetic field can be expressed mathematically by a spherical harmonic model, the so-called IGRF (International Geomagnetic Reference Field) model [18]. For purposes of simulation a first order dipole model is utilised in order to represent the geomagnetic field vector. This dipole vector is expressed as:

TeeT

3RR1MMR

B

3s

3s RR

(6)

where,

is the vector gradient operator

Rs is the length of the geocentric position vector

R is the unit geocentric position vector

Me is the geomagnetic strength of the dipole vector

1 is the identity matrix

In orbital coordinates, the model is expressed as:

sinsin2

cos

cossin

i

i

i

R

M

B

B

B

3s

e

oz

oy

ox

oB (7)

where,

i is the orbit inclination and is the orbit angle measured from the ascending node.The model then has to be referred to the inertial frame (XI, YI, ZI):

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cossinsincossin2

cos

sinsin2cossin 22

ii

i

ii

R

M3s

eiB

Having described the geomagnetic field, the relation between the vector dipole moment M from the magnetic

torquers and the torque required to dump the required angular momentum needs to be made. M is given by:

2

c

B

NBM

(8)

where,

M is the dipole moment vector

B is the geomagnetic field vector measured in body coordinates

Nc is the torque required to dump the required angular momentum

The momentum dumping control logic is defined as:

Nc = k(h-href) (9)

where k =1 and h is angular momentum of reaction wheel and href is the required angular momentum (zero).

The magnetic control angular momentum dumping logic is implemented for the SSFV sailcraft problem using the following case studies:

i) Single axis -90 degree yaw reorientation maneuver without the SRP disturbance and magnetic control angular momentum dumping logic (10 A-m2 moment)

ii) Single axis -90 degree yaw reorientation maneuver with the SRP disturbance and without the magnetic control angular momentum dumping logic (Figure 7)

iii) Single axis -90 degree yaw reorientation maneuver with the SRP disturbance and magnetic control angular momentum dumping logic (100 A-m2 moment)

iv) 55 degree yaw maneuver in a full 3-axis coupled dynamic model for orbit raising without the magnetic control angular momentum dumping logic

v) 55 degree yaw maneuver in a full 3-axis coupled dynamic model for orbit raising with the magnetic control angular momentum dumping logic (100 A-m2 moment)

Table 7 shows the simulation parameters used.

Parameter ValueSize of sail 40 m

Moments of Inertia [Ix, Iy, Iz] [4340, 2171, 2171] kg-m2Solar radiation pressure P 4.563e-6 N/m2

Sail thrust coefficient 1.3Cm/cp distance 0.1 m

Reaction wheel torque saturation 10 m-NmAltitude of orbit 1000 km

Inclination of orbit 98 degAttitude feedback gain Kr 300

Table 7: Simulation Parameters

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0 2 4 6-100

-50

0

time(hours)

Yaw

Ang

le

(de

g)

0 2 4 6-0.06

-0.04

-0.02

0

0.02

time(hours)

Yaw

Rat

e d

/dt

(deg

/s)

0 2 4 6-2

-1

0

1

2

time(hours)

RW

Mom

entu

m(N

-m-s

)

0 2 4 6-10

-5

0

5

10

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 6-2000

-1000

0

1000

2000

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-0.2

-0.1

0

0.1

0.2

time(hours)

Mag

netic

Tor

que(

mN

-m)

0 2 4 6-4

-2

0

2

4x 10

-5

time(hours)

B (

Tes

la)

0 2 4 6-10

-5

0

5

10

time(hours)

M (

A-m

2 )

x

y

z

Figure 8: Single axis -90 degree yaw reorientation maneuver without the SRP disturbance and magnetic control angular momentum dumping logic (10 A-m2 moment)

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0 2 4 6-100

-50

0

time(hours)

Yaw

Ang

le

(de

g)

0 2 4 6-0.06

-0.04

-0.02

0

0.02

time(hours)

Yaw

Rat

e d

/dt

(deg

/s)

0 2 4 6-3

-2

-1

0

1

2

time(hours)

RW

Mom

entu

m(N

-m-s

)

0 2 4 6-10

-5

0

5

10

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 6-4000

-2000

0

2000

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-2

-1

0

1

2

time(hours)

Mag

netic

Tor

que(

mN

-m)

0 2 4 6-4

-2

0

2

4x 10

-5

time(hours)

B (

Tes

la)

x

y

z

0 2 4 6-100

-50

0

50

100

time(hours)

M (

A-m

2 )

Figure 9: Single axis -90 degree yaw reorientation maneuver with the SRP disturbance and magnetic control angular momentum dumping logic (100 A-m2 moment)

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0 1 2 3 4 5 6-1

-0.9

-0.8

-0.7

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0

time(hours)

Sol

arsa

il T

orqu

e(m

N-m

)

Figure 10: SRP Disturbance

Figure 8 shows the single axis -90 degree yaw reorientation maneuver without the SRP disturbance and magnetic control angular momentum dumping logic (10 A-m2 moment). This simulation validates the magnetic control angular momentum dumping logic using 10 A-m2 moment which is a reasonable choice for a 100-kg small satellite magnetorquers. This simulation also confirms that if the SRP disturbance is canceled via sailcraft trimming or ballast control the angular momentum dumping process will be easier. Figure 9 illustrates the sailcraft problem of excessive angular momentum build up due to the constant SRP disturbance and the need to dump the added angular momentum. As mentioned the angular momentum build up is due to solar torque (approx. 1mNm) as depicted in Figure 10. The RW momentum keeps up building in the negative direction even after the completion of the maneuver in order to compensate for the SRP disturbance unless it is dumped via magnetic torque. The 100 A-m2

magnetic-moment is used to give a peak value of 2mNm periodic magnetic torque (see Fig. 9) to counter the 1mNm secular solar torque. When the magnetic field B becomes weak then the RW compensates solar torque and results in a small momentum build up. Meanwhile, when the B-field becomes strong and MT starts dumping momentum a small build up of momentum in the RW is generated and compensates for the SRP disturbance. This explains the angular momentum oscillations in Figure 9 which occur due to the magnetic torque becoming weak for certain parts of the SSFV orbit and due to the persistent SRP disturbance of 1 mNm.

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0 2 4 6-100

-50

time(hours)

Yaw

Ang

le

(de

g)

0 2 4 6-0.1

0

0.1

time(hours)Yaw

Rat

e d

/dt

(deg

/s)

0 2 4 6-10

0

10

time(hours)RW

Mom

entu

m(N

-m-s

)

0 2 4 6-10

0

10

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 6-10000

-5000

0

5000

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-1

0

1

time(hours)Mag

netic

Tor

que(

mN

-m)

Figure 11: 55 degree Yaw maneuver in a full 3-axis coupled dynamic model without momentum dumping

0 2 4 6-0.1

0

0.1

time(hours)

Rol

l Ang

le

(de

g)

0 2 4 6-2

0

2x 10

-4

time(hours)Rol

l Rat

e d

/dt

(deg

/s)

0 2 4 60

2

4

time(hours)RW

Mom

entu

m(N

-m-s

)

0 2 4 6-2

0

2

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 60

2000

4000

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-1

0

1

time(hours)Mag

netic

Tor

que(

mN

-m)

Figure 12: Roll Axis, 55 degree Yaw maneuver in a full 3-axis coupled dynamic model without momentum dumping

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0 2 4 60

0.05

time(hours)

Pitc

h A

ngle

(

deg)

0 2 4 6-2

0

2x 10

-4

time(hours)Pitc

h R

ate

d /d

t (d

eg/s

)

0 2 4 60

10

20

time(hours)RW

Mom

entu

m(N

-m-s

)

0 2 4 6-1

-0.5

0

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 60

5000

10000

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-1

0

1

time(hours)Mag

netic

Tor

que(

mN

-m)

Figure 13: Pitch Axis, 55 degree Yaw maneuver in a full 3-axis coupled dynamic model without momentum dumping

Figures 11-13 show the sailcraft 55 degree yaw maneuver in a full 3-axis coupled dynamic model without momentum dumping. Without sail trim and with a 0.1 m cm/cp offset the angular momentum on all 3 axes builds up excessively up to 12 Nms in 5-6 hours in the yaw and pitch axes and up to 4 Nms in roll due to the 3-axis coupling. The momentum build up due to the SRP disturbance will have to be cancelled through thruster firings however their excessive values will bring a large mass overhead to the bus design. Therefore for the SSFV mission the MT can save a substantial amount of mass if the can be used to control the angular momentum to smaller values or within a band of values (e.g. ± 1-2 Nms) where small thruster firings can completely cancel the angular momentum.

Figure 14-16 show the sailcraft 55 degree yaw maneuver in a full 3-axis coupled dynamic model with momentum dumping. An increase in magnetic moment of 100 A-m2 is needed in order to generate substantial magnetic torque to cancel the angular momentum required. Figure 14 shows the required maneuver completed and the angular momentum controlled within a band of ± 1.8 Nms, a significant finding for the SRP disturbance of 1 mNm. A larger magnetorquers can be used however a trade-off with a thruster design needs to be done in order to showcase possible mass savings for the platform. The weak magnetic torque for aprts of the orbit, again, which is a known constraint of magnetic control, makes the oscillations in angular momentum and torque dominate the simulation results, however the angular momentum is completely dumped when the torque is maximized. Angular momentum oscillations are not so intense on the roll axis due to the sailcraft 3-axis coupling though yaw and pitch axis momentum oscillations are similar in maximum/minimum values (band of ± 1.8 Nms).

American Institute of Aeronautics and Astronautics17

0 2 4 6-100

-50

time(hours)

Yaw

Ang

le

(de

g)

0 2 4 6-0.1

0

0.1

time(hours)Yaw

Rat

e d

/dt

(deg

/s)

0 2 4 6-2

0

2

time(hours)RW

Mom

entu

m(N

-m-s

)

0 2 4 6-10

0

10

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 6-2000

-1000

0

1000

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-5

0

5

time(hours)Mag

netic

Tor

que(

mN

-m)

Figure 14: 55 degree Yaw maneuver in a full 3-axis coupled dynamic model with momentum dumping

0 2 4 6-0.1

0

0.1

time(hours)

Rol

l Ang

le

(de

g)

0 2 4 6-5

0

5x 10

-4

time(hours)Rol

l Rat

e d

/dt

(deg

/s)

0 2 4 6-2

0

2

time(hours)RW

Mom

entu

m(N

-m-s

)

0 2 4 6-5

0

5

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 6-500

0

500

1000

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-5

0

5

time(hours)Mag

netic

Tor

que(

mN

-m)

Figure 15: Roll Axis, 55 degree Yaw maneuver in a full 3-axis coupled dynamic model with momentum dumping

American Institute of Aeronautics and Astronautics18

0 2 4 6-0.1

0

0.1

time(hours)

Pitc

h A

ngle

(

deg)

0 2 4 6-1

0

1x 10

-3

time(hours)Pitc

h R

ate

d /d

t (d

eg/s

)

0 2 4 6-2

0

2

time(hours)RW

Mom

entu

m(N

-m-s

)

0 2 4 6-5

0

5

time(hours)

RW

Tor

que

(mN

-m)

0 2 4 6-1000

-500

0

500

time(hours)

Req

uire

d T

orqu

e(m

N-m

)

0 2 4 6-5

0

5

time(hours)Mag

netic

Tor

que(

mN

-m)

Figure 16: Pitch Axis, 55 degree Yaw maneuver in a full 3-axis coupled dynamic model with momentum dumping

V. ConclusionA preliminary analysis on the design of advances satellite bus for solar sail mission is presented for different

solar sail mission studies. Small satellite technology is fundamental for solar sail mission success as achieving the sailcraft mission parameters depends heavily on the sailcraft bus miniaturisation and functionality. Solar Sail bus technologies are explored in this paper with an emphasis on solar sail ACS based on the sailcraft bus. Reaction wheels can be used for sailcraft attitude control and for LEO missions they can be practical for near term implementation. The largest disadvantage of using reaction wheels is the excessive build up in angular momentum due to the SRP disturbance for untrimmed solar sails (with cm/cp offsets) which is a fixed disturbance that needs to be counteracted. An angular momentum dumping logic using magnetic control is presented which can control the excessive (> 10 Nms) angular momentum to a small band of max/min values (1.8 Nms) using small satellite sized magnetorquers and thus providing substantial mass savings.

.

References[1] McInnes, C. R., Solar Sailing: Technology, Dynamics and Mission Applications, Springer Praxis Publishing,

1999.

[2] Friedman, L., Star Sailing: Solar Sails and Interstellar Travel, John Wiley & Sons, New York, 1988.

[3] Wright, J. L. Space Sailing, Gordon and Breach Science Publishers, 1992.

[4] Wie, B., “Solar Sail Attitude Control and Dynamics: Parts 1 and 2,” Journal of Guidance, Control, and Dynamics, Vol. 27, No. 4, pp. 526–544

[5] Cosmos 1 Solar Sail Mission, http://www.planetary.org/solarsail, accessed July 18, 2005

American Institute of Aeronautics and Astronautics19

[6] Lappas, V., Wie, B., McInnes, C., Tarabini, L., Gomes, L., Wallace, K., ‘A Solar Kite Mission to Study the Earth’s Magneto-tail’, Journal of the British Interplanetary Society, Vol. 58, No. 1/2, Jan/Feb 2005

[7] Leipold, M.; Garner, C.E.; Freeland, R.; Herrmann, A.; Noca, M.; Pagel, G.; Seboldt, W.; Sprague, G.; Unckenbold, W., ‘ODISSEE - A Proposal for Demonstration of a Solar Sail in Earth Orbit’, Acta Astronautica, Vol. 45, Nos. 4-9, pp. 557-566, 1999

[8] Murphy, D. M., Murphey, T. W., and Gierow, P. A., “Scalable Solar-Sail Subsystem Design Concept,” AIAA Journal of Spacecraft and Rockets, Vol. 40, No. 4, 2003, pp. 539-547.

[9] Wie, B. Murphy, D., Paluszek, M., Thomas, S., “Robust Attitude Control Systems Design for Solar Sail Spacecraft: Parts 1 and 2,” AIAA 2004-5010 and 2004-5011, AIAA Guidance, Navigation, and Control Conference, Providence, RI, August 18-19, 2004

[10] McInnes, C.R., Macdonald, M., Angelopolous, V., and Alexander, D.: ‘GEOSAIL: Exploring the Geomagnetic Tail Using a Small Solar Sail’, Journal of Spacecraft and Rockets, Vol. 38, No. 4, pp. 622-629, 2001.

[11] Leipold, M., Lappas, V, Lyngvi, A., Falkner, P., Fichtner, H., Kraft, S., “Interstellar Heliopause Probe: System Design of a Solar Sail Mission to 200 AU”, AIAA Guidance, Navigation, and Control Conference, San Fransisco, CA, August 13-18, 2005

[12] Wie, B., ‘Solar Sailing Kinetic Energy Interceptor (KEI) Mission for Impacting/Deflecting Near-Earth Asteroids’, 41st AIAA Joint Propulsion Conference, Tuscon, AZ, July 10-13, 2005

[13] Wie, B., Murphy, D., ‘Solar-Sail ACS Development for a Sail Flight Validation Mission in a Sun-Synchronous Orbit’, Submitted to Journal of Spacecraft and Rockets, 2006

[14] NASA GSFC RFI For the Spacecraft Bus for the ST-9 Solar Sail, http://www.spaceref.com/news/viewsr.html?pid=18119, Accessed November 24, 2005.

[15] Steyn WH., Hashida Y., Lappas V., “An Attitude Control System and Commissioning Results of the SNAP-1 Nanosatellite”, 14th AIAA/USU Small Satellite Conference proceedings, SSC00-VIII-8, August 2000

[16] Steyn, W.H., Hashida Y., “An Attitude Control System for a Low-Cost Earth Observation Satellite with Orbit Maintenance Capability”, Proceedings of the 13th Annual AIAA/USU Conference on Small Satellites, Utah State University, Logan, Utah, August 1999

[17] Steyn, WH, A Multi-mode Attitude Determination and Control System for Small Satelites, PhD Thesis, Stellenbosch, 1995

[18] Wertz, J. R., “Spacecraft Attitude Determination and Control”, Reidel Publishing Company, Dordrecht, Holland, 1978