Dylan M. Pederson and Jeffrey S. Hojnicki

Glenn Research Center, Cleveland, Ohio

Analysis of Roll Steering for Solar Electric

Propulsion Missions

NASA/TM—2012-217683

August 2012

AIAA–2012–3890
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Dylan M. Pederson and Jeffrey S. Hojnicki

Glenn Research Center, Cleveland, Ohio

Analysis of Roll Steering for Solar Electric

Propulsion Missions

NASA/TM—2012-217683

August 2012

AIAA–2012–3890

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

48th Joint Propulsion Conference and Exhibit

cosponsored by AIAA, ASME, SAE, ASEE

Atlanta, Georgia, July 29 to August 1, 2012
Acknowledgments

Dylan Pederson and Jeffrey Hojnicki thank the Power Systems Engineering branch at NASA Glenn Research Center for support 

in this study. Special thanks are due to Thomas Kerslake and Robert Klimek for guidance through this project. Mr. Pederson also

thanks the Undergraduate Student Research Program and the Universities Space Research Association for providing an opportunity 

to work and learn at NASA Glenn Research Center.

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Level of Review: This material has been technically reviewed by technical management. 

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NASA/TM—2012-217683 1 

Analysis of Roll Steering for Solar Electric Propulsion Missions 

Dylan M. Pederson and Jeffrey S. Hojnicki 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Nothing is more vital to a spacecraft than power. Solar Electric Propulsion (SEP) uses that power to 

provide a safe, reliable, and, most importantly, fuel efficient means to propel a spacecraft to its 

destination. The power performance of an SEP vehicle’s solar arrays and electrical power system (EPS) is 

largely influenced by the environment in which the spacecraft is operating. One of the most important 

factors that determines solar array power performance is how directly the arrays are pointed to the sun. To 

get the most power from the solar arrays, the obvious solution is to point them directly at the sun at all 

times. Doing so is not a problem in deep space, as the environment and pointing conditions that a 

spacecraft faces are fairly constant and are easy to accommodate, if necessary. However, large and 

sometimes rapid variations in environmental and pointing conditions are experienced by Earth orbiting 

spacecraft. SEP spacecraft also have the additional constraint of needing to keep the thrust vector aligned 

with the velocity vector. Thus, it is important to analyze solar array power performance for any vehicle 

that spends an extended amount of time orbiting the Earth, and to determine how much off-pointing can 

be tolerated to produce the required power for a given spacecraft. This paper documents the benefits and 

drawbacks of “perfectly pointing” the solar arrays of an SEP spacecraft spiraling from Earth orbit, and 

how this might be accomplished. Benefits and drawbacks are defined in terms of vehicle mass, power, 

volume, complexity, and cost. This paper will also look at the application of various solar array pointing 

methods to future missions. One such pointing method of interest is called “roll steering”. Roll steering 

involves rolling the entire vehicle twice each orbit. Roll steering, combined with solar array gimbal 

tracking, is used to point the solar arrays perfectly towards the sun at all points in the orbit, while keeping 

the vehicle thrusters aligned in the velocity direction. Roll steering is particularly attractive for a recently 

proposed mission that involves a spiral trajectory from low Earth orbit (LEO) to the Earth-Moon 

Lagrange Point 1 (E-M L1). During the spiral, the spacecraft will spend over 300 days experiencing the 

full spectrum of near-earth environments and solar array pointing conditions. An extensive study of the 

application of SEP (and roll steering) to this spiral mission is included, highlighting the ultimate goal of 

reduced vehicle cost and mass. Tools used for this analysis include the Systems Power Analysis for 

Capability Evaluation (Refs. 1 and 2) (SPACE) electrical power systems code, and SEP trajectory 

simulation tools developed at NASA Glenn Research Center. 

Nomenclature 

 angle between the vector from Earth to orbit noon and the Earth-Sun vector 

SEP Solar Electric Propulsion 

EPS Electrical Power System 

HET Hall Effect Thruster 

RAAN Right Ascension of the Ascending Node 

ISS International Space Station 

CMG Control Moment Gyroscope 

E-M L1 Earth-Moon Lagrange Point 1 

SPACE System Power Analysis for Capability Evaluation 
NASA/TM—2012-217683 2 

Introduction 

Spacecraft that employ solar electric propulsion (SEP) for primary propulsion are already flying in 

interplanetary space (e.g., Deep Space 1, Dawn) and have proven that SEP is a key enabling technology 

for some solar system exploration missions. SEP uses solar arrays to generate power for electric thrusters. 

These thrusters, which could be Hall effect thrusters (HETs) or ion thrusters, expel ionized xenon gas to 

generate thrust. Present HET capability provides a very small amount of thrust (0.5N – 3N) which slowly 

accelerates the spacecraft to high speeds. Although HETs are very efficient, their low thrust means it 

takes a long time to reach sufficient speeds. NASA is now investigating applications that use the HET 

thrusters to propel a spacecraft from low Earth orbit (LEO) and escape Earth’s gravitational pull. The low 

thrust of the SEP system necessitates a trajectory which spirals slowly from LEO out to deeper space. 

Because of the extended trip time and increasing altitudes during the spiral, the spacecraft making the trip 

will experience the full spectrum of space environments from low to high Earth orbit. 

Some of the most significant challenges faced during the spiral involve variations in solar flux, 

radiation dose, atomic oxygen, atmospheric density, temperature, and  angle. Variations in these 

parameters have a great impact on the power production of the spacecraft’s solar arrays. This, in turn can 

have a significant impact on the HET specific impulse and thrust, and therefore, on the total spiral trip 

time. One variable of particular interest to this analysis is the solar  angle (Fig. 1). The  angle varies 

throughout the year because of the Earth’s movement around the Sun and from precession of the 

longitude of the ascending node of the spacecraft’s orbit. Additionally, for spiral trajectories, the rate of 

change of the  angle changes as the altitude increases. This gives rise to an interesting  angle profile 

throughout the trajectory. The  angle profile depends on many different variables, but a typical profile 

for a spiral trajectory is shown in Figure 2. 

The  profile depends on the right ascension of the ascending node (RAAN) of the spacecraft’s orbit 

at the beginning of the spiral. After the spacecraft reaches a high enough altitude, the RAAN no longer 

varies from Earth’s oblateness. Depending on when the RAAN precession rate becomes negligible, the 

spacecraft could experience large oscillations in  (50°) or smaller oscillations (or even no oscillations at 

all). Some possible  profiles for this spiral mission are shown in Figure 3. 

The  angle is important to spacecraft power because it affects the amount of off-pointing a solar 

array will experience, if left uncompensated. A typical spacecraft will consist of a bus, with two solar 

array wings mounted collinear with one of the spacecraft’s principle axes. The wings typically are pointed 

by single axis gimbals that are used to compensate for once-per-orbit revolution of the spacecraft around 

the Earth (i.e., so-called alpha pointing). Compensation for the  angle is either not performed (i.e., accept 

off-pointing), or handled by varying the attitude of the spacecraft. Since an SEP spacecraft must also 

maintain the vehicle’s thrust axis in a specific direction, there are limited attitudes that can be flown. 

 
NASA/TM—2012-217683 3 

Therefore, a solar array with a single axis gimbal, barring any other corrections, will experience an offpointing equivalent to the  angle (e.g., for 20°  angle, the solar array will suffer 20° off-pointing; for 

30° , 30° off-pointing, and so on). When the solar array is off-pointed, it loses power because it is 

absorbing a smaller area of sunlight. The solar array is said to suffer from “cosine power loss”. This is 

because the power with respect to  is closely approximated by 

P  P0cos (1)

where P(0) is the power produced at a  angle of 0°. In other words, maximum power is achieved at 

 = 0°. Clearly, if there is a way to eliminate off-pointing completely, it could provide a boost in power 

production for such a spiral mission which would decrease the trip time through the spiral. 

Power Benefit of Perfect Pointing 

A solar array is said to be perfectly pointing when it is adjusted to point directly towards the sun at all 

times (i.e., total elimination of off-pointing, within the accuracy that the attitude control, structures and 

mechanisms allow). For any mission that will experience a wide range of  angles, it is important to know 

exactly how much perfect pointing might improve power generation compared to allowing off-pointing. 

To determine the effect that perfect pointing has on spacecraft power, analyses were performed with the 

SPACE (Refs. 1 and 2) power model to study orbits with perfectly pointed solar arrays and orbits with 

off-pointed solar arrays. Relevant variables for the spiral mission power are orbital altitude and  angle. 

Altitude is important because at higher altitudes, the time spent in insolation during an orbit increases, 

which improves overall energy production. The  angle is important for reasons mentioned earlier. 

According to SPACE analyses, although the magnitude of the power gain from perfect pointing 

(compared to off-pointing) does depend on orbit altitude, the percentage gain does not. That is, the 

benefits of perfect pointing are similar at all altitudes. This makes sense intuitively, as the cosine power 

rule does not depend on orbital altitude. Therefore, just one altitude was used for further examination. 

As Figure 4 indicates, perfect pointing is most beneficial at higher  angles. Perfect pointing gives 

just 1 percent power gain on orbit average (compared to off-pointed power generation) at a 10°  angle. 

On the other hand, at large  angles perfect pointing can give well over 50 percent more power than offpointing. The  profile for this mission shows that the spacecraft will experience  angles from –52° up to 

+52°. Thus, a significant amount of power can be gained from perfect pointing throughout the spiral. 
NASA/TM—2012-217683 4 

Options to Compensate for Off-Pointing? 

Spacecraft have been able to achieve perfect pointing in the past using a variety of methods. For 

example, the International Space Station (ISS) solar structure has two gimbals for each solar array. The 

two gimbal rotation axes are orthogonal to each other, allowing for perfect pointing at all times (there are 

some complications to this that will be discussed later). Other methods are available, and each has its 

drawbacks. This study will consider the following methods. 

Larger Solar Arrays 

Perhaps the most obvious way to gain more power throughout a mission (and therefore speed up the 

trip) has nothing to do with pointing, but instead is to make the solar arrays bigger. The approach is 

simple: by making the solar arrays 10 percent bigger, then 10 percent more power is available at all 

angles. The obvious downside is that the solar array still suffers from cosine power loss, but now on a 

larger scale! Also, the electrical power system (EPS), thermal control system (TCS), and spacecraft 

structure must be sized accordingly for a larger solar array. In general, this means that spacecraft as a 

whole will grow in mass. In most cases, a spacecraft’s mass is limited by the lift capabilities of its launch 

vehicle. If the spacecraft outgrows the launch vehicle, then the mission may need to consider a larger (and 

more expensive) launch vehicle to accommodate the spacecraft. 

In addition to mass concerns, making a solar array larger is expensive. Solar cell production and 

installation achieves limited “economies of scale”. This is because, presently, solar cell production is 

mostly a batch process and the installation of the cells onto the solar array wing involves manual 

processing. Larger solar array sizes can also be more difficult (and costly) to build, test and integrate with 

the spacecraft and to fit inside the launch vehicle fairing. So, in general, making a solar array larger is 

undesirable. 
NASA/TM—2012-217683 5 

Two-Axis Tracking 

The largest and most powerful solar array wings ever flown in space are on the ISS. As mentioned 

earlier, the ISS solar array structure provides two gimbals for each solar array. This method of solar array 

pointing is called two-axis tracking, because the solar arrays can rotate along two orthogonal axes. 

Case I—ISS Model 

The ISS solar arrays all have one rotation axis in common, controlled by the solar alpha rotary joint 

(SARJ). Additionally, every solar array has a second gimbal, known as a beta gimbal that rotates the array 

along its individual axis perpendicular to the alpha rotation axis. These two gimbals allow each array to 

point perfectly (within tolerance) towards the sun. The weight of each Beta gimbal is approximately 

323 kg, while the weight of the SARJ is 1161 kg. This arrangement is considered mass prohibitive for an 

SEP spiral mission, and the challenge of packaging such a structure into a single launch vehicle may be 

even more challenging. 

Case II—Dumbo Model 

An alternative method of two-axis tracking is to have two adjacent gimbals for each wing. One 

gimbal would be stacked upon the other, where the wing meets the vehicle body. This model is named 

after its Dumbo-like “flapping” of the solar array wings. This would provide perfect pointing without the 

need for a large alpha truss structure. One concern with this method is that when the wings are tracking 

the sun, it is possible for the center of mass of the vehicle to shift, causing the vehicle to pitch while the 

HETs are running. This would necessitate the ability to gimbal the thrusters continuously as the solar 

arrays are tracking, leading to propulsion performance losses and hence, longer spiral mission trip times. 

Additionally, while the arrays are tracking, array shadowing may be more likely, plus the vehicle 

designers will have to be conscious to prevent the wings from bumping into the vehicle body. Such 

complexity in spacecraft operation is undesirable. Further attention should be given to the feasibility of 

this method, but that is not the purpose of this study. 

Roll Steering 

A solar array pointing method known as “roll steering” is achieved by rolling the spacecraft about the 

velocity vector to achieve the same pointing effectiveness as a second gimbal. This requires rolling the 

spacecraft twice per orbit while maintaining the thrust vector closely collinear with the velocity vector. 

This allows the solar arrays to perfectly point with just one gimbal. To roll a large vehicle, control 

moment gyroscopes (CMGs) are preferred over reaction wheels or magnetic torquers because CMGs 

provide the large torque and momentum storage capability needed for such a maneuver. Roll steering 

provides perfect pointing, and thus reaps the power benefits outlined earlier. 

The number of CMGs required to implement full roll steering will depend on the total vehicle 

moment of inertia about its roll axis and the torque capability of the CMGs. Torque (as opposed to 

momentum) is the limiting factor in determining how many CMGs are needed to fully roll steer, because 

at low  angles the vehicle is required to roll at a large rate that is sometimes referred to as a “snap” roll 

(as opposed to a gradual roll). These snap rolls can require a large torque to complete. The drawbacks of 

full roll steering include a potentially large mass penalty for the CMGs and increased operational 

complexity. 
NASA/TM—2012-217683 6 

Roll Switch 

Roll switching is a variation of roll steering that tolerates a certain amount of off-pointing. In other 

words, the vehicle only rolls when the  angle is greater than a specified angle. This is done to take 

advantage of the large power benefit from rolling at higher  angles, and to avoid the extra mass of CMGs 

required to roll at lower  angles. The cross-over  angle (between roll steering and not) will depend on 

the moment of inertia of a specified vehicle and its CMG capabilities. 

Application to an Asteroid Mission 

To compare the alternatives for solar array pointing, each method was separately applied to a 

proposed mission vehicle. The vehicle specifications came from the Human Exploration Framework 

Team (HEFT) and the Collaborative Modeling for Parametric Assessment of Space Systems 

(COMPASS) team at NASA Glenn Research Center. COMPASS and HEFT have laid out and analyzed a 

proposed human mission to the asteroid 2008-EV5 (henceforth referred to as the EV5 mission) that 

heavily uses SEP technology (Refs. 3 to 5). This mission involves a spiral trajectory carrying cargo from 

low-Earth orbit to the Earth-Moon Lagrange point 1 (E-M L1) followed by a heliocentric trajectory 

carrying crew to the asteroid, both using a 300-kW-class SEP vehicle. This analysis uses the EV5 SEP 

vehicle as a baseline for comparison. 

To fairly evaluate roll steering and its alternatives, it is necessary to compare to the base case of not 

rolling the vehicle and accepting off-pointing. A quantitative comparison was performed with an SEP low 

thrust trajectory simulation tool when possible. Otherwise, qualitative comparisons were performed where 

the relative performance of the various options was estimated, since a detailed assessment of those factors 

could not be performed for this analysis. The following information is summarized in Table 1. 

Larger Solar Arrays 

Scaling the 300 kW solar arrays on the EV5 vehicle up by 10 percent would add around 300 kg of 

mass to the vehicle, taking into account resizing of the EPS, TCS, and vehicle structure. The real problem 

with the larger arrays is that they add cost to the most expensive part of the vehicle. Figure 6 shows that 

the most expensive part of an SEP vehicle is the solar array. This is why it is impractical to make larger 

solar arrays. 

TABLE 1.—COMPARISON OF ALL METHODS TO THE NO ROLL STEERING CASE 

Pointing 

method 

Vehicle 

mass, 

kg 

Array 

powera, 

W 

Flight 

hardware 

volume 

Flight software 

development 

Systems 

complexity 

Vehicle 

cost 

Est. trip time 

to E-M L1, 

days 

No roll 

steering 

77424 ---- Low Low Low $$ 321 

Roll steering 79056 +14% Med High Med $$$ 295 

Upsize solar 

arrays 

10 percent 

77700 +10% High N/A High $$$$$ 285 

Two axis 

tracking—

ISS model 

79000+ +14% High High High $$$$ 290 

Roll switch at 

 = 21° 

78240 +13% Low High Med $$ 294 

a

All power numbers are in comparison to the case of no roll steering.
NASA/TM—2012-217683 7 

Two-Axis Tracking 

To two-axis track, this 300 kW class SEP vehicle would require eight beta gimbals and two structures 

equivalent to the ISS SARJ. The solar arrays on the 300 kW vehicle are smaller than the ISS solar arrays, 

so the required gimbal size would be smaller. Each beta gimbal is estimated to be around 215 kg (Ref. 6), 

totaling to 1720 kg (not including the alpha gimbal and structure)!

Full Roll Steering 

This study assumes that the same CMGs as are used on the ISS will be used for the EV5 vehicle to 

perform full roll steering, because the ISS CMGs are flight proven, and are approximately the required 

capability needed for the SEP mission. The ISS CMGs can provide 4760 Nms of momentum storage and 

258 Nm of torque output. The torque and momentum required for roll steering the 300-kW-class SEP 

vehicle are shown in Figure 5. For this vehicle to roll at almost all  angles (5° and above), a total of 

6 ISS CMGs are necessary (5 for rolling, and 1 for redundancy). The CMGs add 1632 kg of mass to the 

vehicle, and draw approximately 900 W of power (averaged over one orbit). 

Roll Switching 

As mentioned earlier, the optimal cross-over  depends on the vehicle requirements. For this 300 kW 

SEP vehicle, the cross-over  will be 21° so that just 3 CMGs are needed (2 for rolling, and 1 for 

redundancy). This  was found by numerically optimizing the balance between extra mass from CMGs 

and extra power from rolling. Compared to full roll steering, this method cuts the weight and power load 

from the CMGs in half to 816 kg and 450 W. It is fairly evident that full roll steering and two-axis 

tracking are the worst two alternatives overall. In fact, from a quick glance at Figure 6, no roll steering 

may seem like the best option. But, the “best” option really depends on what the mission is trying to 

accomplish. If the spacecraft is trying to reach its destination as quickly as possible (as a communications 

satellite might do, to make profits as soon as possible) then it may be best to simply have larger arrays. 

 
NASA/TM—2012-217683 8 

TABLE 2.—DIRECT COMPARISON OF NO ROLL STEERING AND ROLL SWITCHING WITH RESIZED ARRAYS 

(FOR A SPIRAL WITH AN INITIAL RAAN OF 110° AND A START DATE OF JANUARY 1) 

Pointing 

method 

Vehicle 

mass, 

kg 

Array 

power 

Flight 

hardware 

volume 

Flight software 

development 

Systems 

complexity 

Vehicle 

cost 

Est. trip 

time to 

E-M L1, 

days 

No roll 

steering 

77424 300 kW 

EOL 

High Low Low $$$$ 321 

Resized 

arrays with 

roll switch 

at  = 21° 

78102 270 kW 

EOL 

Lower High Med $ 319 

However, “best” in the context of an SEP spacecraft for the given mission is defined as the lowest cost 

and mass. The spiral mission is not concerned with trip time as long as it is within the mission time frame. 

In fact, if the spacecraft gets to E-M L1 quickly then it will potentially have to use more propellant to 

maintain its “halo” orbit for a longer period of time while waiting to rendezvous with other mission 

elements. Thus, upsizing the solar arrays is really not a good option for this mission. 

So far, three options can be eliminated; larger solar arrays are not good because they do not fit the 

definition of “best” as we define it here, and two-axis tracking and full roll steering are trumped by the 

lighter weight and lower cost of roll switching. To get a fair, head-to-head comparison of the remaining 

options (roll switch and no roll), the factors that are unimportant to this mission must be evened out. In 

other words, the trip time of the roll switch case and the no roll case must be the same to get a fair 

comparison of cost and mass. The solar array power levels and propellant loads for each of the two 

options were adjusted until trajectory analyses showed the trip times of each case matched to 3201 days. 

The resulting effect on vehicle cost and mass are shown in Table 2. 

From this perspective, roll switching looks more attractive, as it saves cost by allowing for smaller 

solar arrays. The mass of the vehicle is heavier with roll switch, but still stays within the launch vehicle 

constraint of 80 metric tons. This example was the average case (with an initial RAAN of 110°), but 

depending on initial orbital conditions, the power savings can vary from 10 to 50 kW. Even in the worst 

case, reducing the power required by 10 kW can still save millions of dollars on the solar arrays. In 

addition, the size of the TCS and EPS can probably be downsized because of the smaller arrays, saving 

additional mass and cost. 
NASA/TM—2012-217683 9 

Conclusion 

The authors conclude that roll switching can offer one of two benefits, depending on your objective. For 

the EV5 asteroid mission, roll switching can be used to allow smaller solar arrays, and thus save money on 

the spacecraft. Additionally, roll switching can be used to get to a destination more quickly, simply as a 

result of more efficient collection of solar power. Larger arrays can also get a spacecraft to its destination 

faster, but it is more cost effective to pay for relatively cheap CMGs than to pay for larger arrays. 

In general, it is not worth the extra CMG mass to fully roll steer at all  angles. “Roll switching” is 

the better alternative to save money and mass on the vehicle. The overall effectiveness of roll switching 

will depend on the mission  profile. For the example mission, even in the most favorable case (for solar 

array pointing), the spacecraft can reduce the size of its solar arrays by 3 percent with roll switching. 

References 

1. Hojnicki, J.S., et al., “Space Station Freedom Electrical Performance Model,” NASA TM-106395, 

28th Intersociety Energy Conversion Engineering Conference, 1993. 

2. Jannette, A.G., et al., “Validation of International Space Station Electrical Performance Model via 

On-Orbit Telemetry,” IECEC-2002-20007, 37th Intersociety Energy Conversion Engineering 

Conference, 2002. 

3. Brophy, J.R., et al., “300-kW Solar Electric Propulsion System Configuration for Human Exploration 

of Near-Earth Asteroids,” AIAA-2011-5514, 47th AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference & Exhibit, San Diego, California, July 31, 2011. 

4. Hoffman, D.J., et al., “Concept Design of High Power Solar Electric Propulsion Vehicles for a 

Human Exploration,” IAC-11-D2.3.5, 62nd International Astronomical Conference, Cape Town, 

South Africa, October 3–7, 2011. 

5. Capadona, L.A., et al., “Feasibility of Large High-Powered Solar Electric Propulsion Vehicles: Issues 

and Solutions,” AIAA-2011-7251, Space 2011 Conference & Exposition, Long Beach, California, 

September 27–29, 2011. 

6. Mercer, C.R., et al., “Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to 

an Asteroid,” AIAA-2011-7252, Space 2011 Conference & Exposition, Long Beach, California, 

September 27–29, 2011. 
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13. SUPPLEMENTARY NOTES

14. ABSTRACT

Nothing is more vital to a spacecraft than power. Solar Electric Propulsion (SEP) uses that power to provide a safe, reliable, and, most importantly, fuel efficient means to 

propel a spacecraft to its destination. The power performance of an SEP vehicle’s solar arrays and electrical power system (EPS) is largely influenced by the environment in 

which the spacecraft is operating. One of the most important factors that determines solar array power performance is how directly the arrays are pointed to the sun. To get the 

most power from the solar arrays, the obvious solution is to point them directly at the sun at all times. Doing so is not a problem in deep space, as the environment and pointing 

conditions that a spacecraft faces are fairly constant and are easy to accommodate, if necessary. However, large and sometimes rapid variations in environmental and pointing 

conditions are experienced by Earth orbiting spacecraft. SEP spacecraft also have the additional constraint of needing to keep the thrust vector aligned with the velocity vector. 

Thus, it is important to analyze solar array power performance for any vehicle that spends an extended amount of time orbiting the Earth, and to determine how much offpointing can be tolerated to produce the required power for a given spacecraft. This paper documents the benefits and drawbacks of “perfectly pointing” the solar arrays of an 

SEP spacecraft spiraling from Earth orbit, and how this might be accomplished. Benefits and drawbacks are defined in terms of vehicle mass, power, volume, complexity, and 

cost. This paper will also look at the application of various solar array pointing methods to future missions. One such pointing method of interest is called “roll steering”. Roll 

steering involves rolling the entire vehicle twice each orbit. Roll steering, combined with solar array gimbal tracking, is used to point the solar arrays perfectly towards the sun 

at all points in the orbit, while keeping the vehicle thrusters aligned in the velocity direction. Roll steering is particularly attractive for a recently proposed mission that involves 

a spiral trajectory from low Earth orbit (LEO) to the Earth-Moon Lagrange Point 1 (E-M L1). During the spiral, the spacecraft will spend over 300 days experiencing the full 

spectrum of near-earth environments and solar array pointing conditions. An extensive study of the application of SEP (and roll steering) to this spiral mission is included, 

highlighting the ultimate goal of reduced vehicle cost and mass. Tools used for this analysis include the Systems Power Analysis for Capability Evaluation (Refs. 1 and 2) 

(SPACE) electrical power systems code, and SEP trajectory simulation tools developed at NASA Glenn Research Center.

15. SUBJECT TERMS

Solar electric propulsion; Attitude control; Solar arrays 

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STI Help Desk (email:help@sti.nasa.gov) 

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