Kevin E. Witzberger and David Manzella

Glenn Research Center, Cleveland, Ohio

Performance of Solar Electric Powered Deep

Space Missions Using Hall

Thruster Propulsion

NASA/TM—2006-214044

January 2006

AIAA–2005–4268
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Kevin E. Witzberger and David Manzella

Glenn Research Center, Cleveland, Ohio

Performance of Solar Electric Powered Deep

Space Missions Using Hall

Thruster Propulsion

NASA/TM—2006-214044

January 2006

National Aeronautics and

Space Administration

Glenn Research Center

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AIAA–2005–4268
Acknowledgments

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continuous support and feedback.

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Available electronically at http://gltrs.grc.nasa.gov
Performance of Solar Electric Powered Deep Space Missions

Using Hall Thruster Propulsion

Power limited, low-thrust trajectories were assessed for missions to Jupiter, Saturn, and Neptune utilizing 

a single Venus Gravity Assist (VGA) and a primary propulsion system based on either a 3-kW high voltage 

Hall thruster, of the type being developed by the NASA In-Space Propulsion Technology Program, or an 8-

kW variant of this thruster. These Hall thrusters operate with specific impulses below 3,000 seconds. A trade 

study was conducted to examine mission parameters that include: net delivered mass (NDM), beginning-oflife (BOL) solar array power, heliocentric transfer time, required launch vehicle, number of operating 

thrusters and throttle profile. The top performing spacecraft configuration was defined to be the one that 

delivered the highest mass for a range of transfer times. In order to evaluate the potential future benefit of 

using next generation Hall thrusters as the primary propulsion system, comparisons were made with the 

advanced state-of-the-art (ASOA), 7-kW, 4,100 second NASA’s Evolutionary Xenon Thruster (NEXT) for the 

same mission scenarios. For the BOL array powers considered in this study (< 30 kW), the results show that 

the performance of the Hall thrusters, relative to NEXT, is largely dependant on the performance capability 

of the launch vehicle, and that at least a 10% performance gain, equating to at least an additional 200 kg dry 

mass at each target planet, is achieved over the higher specific impulse NEXT when launched on an Atlas 551. 

Nomenclature 

CBE = current best estimate 

V∞ = spacecraft velocity at the sphere of influence with respect to the target planet, km/s 

ΔV = velocity change of spacecraft, km/s 

M0 = launch mass of spacecraft (injected mass), kg 

N = Newton 

mN = milliNewton 

kg = kilogram 

s = second 

kW = kilowatt 

Isp = specific impulse, s 

AU = astronomical unit 

ηsys = system efficiency 

C3 = vis-viva energy, km2

/s2

g = acceleration of gravity on Earth, 9.80665e-3 km/s2

P = Power Processing Unit Input Power, kW 

μE = Earth’s gravitational constant, 398600.485 km3/s2

RE = Earth’s radius, 6378.14 km 

Alt = altitude of spacecraft at escape, km 

R = RE + Alt, km 

V0

2

 = Parking orbit velocity squared, μE/R, km2/s2 

VE

2

 = Earth escape velocity squared, 2 V0

2

 + C3, km2/s2

ΔVL = velocity change of spacecraft provided by the launch vehicle, VE – V0, km/s 

Kevin E. Witzberger and David Manzella 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135

NASA/TM—2006-214044 1
I.  Introduction

rbit raising, repositioning, insertion, and de-orbit functions are a few commonly sited applications of Hall

thrusters.

1,2,3,4 When considering Hall thruster technology for interplanetary missions, past works have focused

on some sort of Earth escape architecture in which Hall thrusters are used just to spiral out from Earth.

2,5 This

architecture enables a smaller launch vehicle to be used, thereby reducing the overall mission costs. A common

theme in past interplanetary mission analysis work is that the optimal specific impulses required for deep space

missions are too high for Hall thrusters to attain, implying that the performance of the Hall thrusters would be

inferior to the higher Isp ion thrusters.

6,7,8 While this may be true for constant power missions, it does not hold true

for certain power-limited deep space missions. Fiehler and Oleson compared Hall and ion technology performance

results for direct Earth-Mars Solar Electric Powered (SEP) missions.

9 Their work showed that a spacecraft with Hall

thrusters, using less array power and fewer thrusters, delivered slightly more payload to Mars than a spacecraft with

ion thrusters. The low thrust trajectory optimization tool used in their study was based on the Direct Trajectory

Optimization Method. This study utilized a calculus-of-variations based optimization code, Solar Electric Propulsion

Trajectory Optimization Program (SEPTOP), developed by Carl Sauer at the Jet Propulsion Laboratory (JPL).

The work in this paper goes beyond that of Fiehler and Oleson and assesses the performance of next generation

Hall thrusters designed for deep space missions for which the performance of ion thrusters, such as NEXT, have

been previously assessed.

10,11 The main objectives of the mission trade studies were to:

(1) Assess the performance capability of a spacecraft using next generation Hall thrusters for its primary

propulsion system by identifying the top performing NEXT and Hall propelled spacecraft to each target

planet and comparing the results;

(2) Gain insight into the importance of the combination of the launch vehicle and BOL array power in

determining the performance of the next generation Hall thrusters;

(3) Identify thruster performance modifications that would enhance the overall performance on a mission.

These classes of missions are to the outer solar system and typically require a Hohmann-type ΔV of 6 – 8 km/s.

A VGA is used to provide increased heliocentric orbital energy, thereby reducing the launch energy (C3)

requirement provided by the launch vehicle. Three planets are targeted: Jupiter, Saturn, and Neptune. The NDM at

each target planet’s sphere of influence (SOI) is the dry mass of all elements of the spacecraft, including the science

payload, less the SEP module and propellant. Capturing into a final parking orbit was not a requirement for this

study. Parametric analysis is presented in the form of NDM versus trip time for a given BOL array power

(referenced at 1 AU) and launch vehicle for both Hall thrusters and NEXT. The trade-off between the number of

operating thrusters and throttling profile are assessed for each destination. The top performing spacecraft

configuration for Hall and NEXT is identified and a relative comparison is made.

At this point it is important to point out that although comparisons are made between Hall and NEXT

technology, they are at different stages of development. NEXT is the more mature technology; it is an advancement

of the current state-of-the-art (SOA) ion thruster, the 2.6-kW, 3,100 second NASA Solar Electric Propulsion

Technology Application Readiness thruster (NSTAR) that successfully served as the primary propulsion system on

Deep Space 1. Notable technological advancements include: increased throttling range (from 5:1 to about 10:1),

about a 30% increase in maximum specific impulse, and nearly a 300% increase in maximum operating power. A

breadboard test has been conducted, and NEXT is on track to be at a NASA technology readiness level (TRL) 6 by

the end of 2006. TRL 6 is defined as system/subsystem model or prototype demonstration in a relevant

environment.

12 In-contrast, the 3-kW Hall thruster and its 8-kW variant is currently at TRL 4, which is defined as

analytical and experimental critical function and/or characteristic proof-of-concept.

12

This paper begins with section II, an overview and status of Hall thruster technology development at Glenn

Research Center. A recent overview of NEXT can be found in reference 10. Mission analysis assumptions and

results are covered in section III. Finally, overall conclusions about this study are presented in section IV.

II.  Hall Thruster Technology Overview and Status

Hall thruster technology has been utilized during the past several decades for Earth orbital station-keeping

applications at power levels on the order of one kilowatt (kW) per thruster due to its ability to reduce launch mass

and increase satellite operational lifetimes.

13 This success has lead to two Earth orbiting missions utilizing Hall

thruster technology for primary propulsion: the United States National Reconnaissance Office’s Science Technology

O

NASA/TM—2006-214044 2
Experiments spacecraft (STEX) launched in 1998 and the European Space Agency SMART-1 Lunar probe launched

in 2003. These spacecraft including their Hall thruster propulsion systems are shown in Fig. 1.

Extending the use of Hall thruster technology to interplanetary missions requires higher powers, increased

specific impulse, and for solar electric applications, the ability to operate efficiently over a range of input powers.

In response to this, in 1999 NASA initiated a program to develop a Hall thruster that could provide this capability by

increasing the specific impulses of existing SOA thrusters from 1500 seconds (sec) to 3000 sec and higher.

14 The

result of this program was the nominally 8-kW NASA-173M thruster, which demonstrated efficient operation at

specific impulses from less than 1500 sec to nearly 3400 sec at power levels from less than 1 kW to nearly 10 kW.

15

In 2003 the NASA In-Space Propulsion Program initiated the High Voltage Hall Accelerator (HIVHAC)

development project, an effort to further develop this technology to support future robotic space-science exploration

missions. While this program initially focused on the 8-kW power level demonstrated by the NASA-173M, in 2004

the power level of interest was reduced to 2.8 kW (with throttling down to 300 Watts) to optimize the technology for

NASA Discovery-Class mission applications. The present study considers the mission benefit of Hall thruster

systems utilizing thrusters at both these power levels for representative deep space missions.

While thruster development work is ongoing, the Hall thruster systems used in this analysis were the author’s

projections based on the current SOA for station-keeping systems and near-term potential capabilities. The

technology requirements and remaining technical challenges for each of the components of the Hall thruster system

are discussed below.

A. Thruster

The thrusters used for these analyses were based on the NASA-170M, a flight-like development model of the

lab-model NASA-173M thruster, and the NASA-77M, a low power variant being developed by the HIVHAC

project. Both these development model thrusters were designed to operate over a wide range of operating voltages

and input powers. The performance of the 8-kW thruster was estimated based on experimental data obtained with

the laboratory model thruster. The performance of the 2.8-kW thruster was estimated by de-rating the 8-kW

performance to account for the greater influence of cathode flow rate and magnet power on the efficiency of the

smaller lower power thruster. The In-Space Propulsion Program will verify these estimates through experiments

planned for early 2005. Both these thrusters have a TRL of 4. The thruster mass used in these analyses was

estimated from the solid models of the mechanical designs of each thruster. Maximum allowable propellant

throughput was estimated at 800 and 300 kg of xenon based on operational lifetimes of 30,000 hours for the 8-kW

and 2.8-kW thrusters respectively. 30,000 hours was selected based on the demonstrated lifetime of the required

hollow cathode.

16 These lifetimes have not been demonstrated and are in excess of the lifetimes required and

demonstrated for prior Earth orbital applications. Both thrusters are shown in Fig. 2.

a) STEX b) SMART-1

Figure 1. Spacecraft that utilize Hall thruster primary propulsion.

NASA/TM—2006-214044 3
B. Power Processing Unit

The power processing units (PPUs) used for these analyses were based on mass and electrical efficiency of the flight

units developed for Earth orbital applications. These include the PPU used on the STEX experiment and the PPU

undergoing qualification as part of 4.5 kW Hall Thruster Propulsion System (HTPS) developed for high power

geostationary communication satellites.

17 Each of these units is shown in Fig. 3. All the spacecraft control and

telemetry functions are incorporated into the PPU due to its simplicity relative to the PPU of a gridded ion thruster

electric propulsion system eliminating the need for a stand alone digital control interface unit (DCIU).

C. Propellant Feed System and Xenon Storage

The feed system and xenon storage tank used for this analysis were based on TRL 6 components previously

developed. The feed system uses an industry developed proportional flow control valve from the HTPS system and

metal-lined, composite-overwrapped propellant tank as used by the xenon fueled gridded ion thruster system used

a) 8-kW NASA-170M b) 2.8-kW NASA-77M

Figure 2. NASA Hall thrusters.

a) 1-kW STEX b) 4.5-kW HTPS

Figure 3. Hall Power Processing Units

NASA/TM—2006-214044 4
on the NASA Deep Space 1 spacecraft.

18 The propellant tank was sized as necessary, assuming a constant tankage

fraction, depending on the required propellant load.

III. Mission Analysis

A. Thruster Modeling

SEPTOP, an indirect two-body trajectory optimization tool, was used for this analysis. It has the ability to

optimize such parameters as launch and fly-by dates and the hyperbolic departure velocity. SEPTOP also optimizes

the discrete number of operating thrusters and can insert coast periods. It has been used for many years for many

SEP mission analyses across various NASA centers. One important limitation of SEPTOP is that it can only model

the thrust or mass flow rate of an individual electric thruster as a function of a single variable: power into the PPU.

As a consequence of this limitation, the extreme regions of the thruster performance data are typically modeled, by

two polynomial curve fits (one for the thrust and one for the mass flow rate), creating a high thrust and high Isp

“envelope” around the performance data. In reality, the thruster may be able to physically throttle in a region that is

anywhere inside the high thrust and high Isp envelope. To better model the thruster, the thrust and mass flow rate

would be a function of two variables: PPU input power and a “control” type variable that would be optimized inside

the region bounded by high thrust and high Isp envelope.

A fourth order polynomial is the maximum order polynomial permitted by SEPTOP. The form of the polynomial

is given in Eq. (1), where P is the PPU input power in kilowatts and y represents either thrust or mass flow rate.

!

y = a + bP + cP2 + dP3 + eP4

(1)

Table 1 provides the operating range of each thruster and

the assumed (except for NEXT) PPU efficiencies.

Several versions of the NEXT throttle table exist and

have been used for recent mission analyses.

10 For this

analysis a recent NEXT throttle table that provides lower

power limits and increased thruster efficiency was used.

Figure 4 shows the curve fits for the 3-kW Hall, 8-

kW Hall, and 7-kW NEXT. Table 2 provides the

corresponding polynomial coefficients.

Table 1 Minimum and maximum thruster and

PPU input powers

Power into thruster, 

kW Power into PPU, kW

Pmin Pmax Pmin Pmax

3-kW Hall 0.312 2.856 95 0.328 3.006

8-kW Hall 1.000 8.000 96 1.042 8.333

7-kW NEXT 0.743 6.864 91-95 0.817 7.225

Thruster PPU efficiency, 

%

Table 2 4th order polynomial curve fit coefficients

Mass flow 

rate, kg/s Thrust, N

Mass flow 

rate, kg/s Thrust, N

Mass flow 

rate, kg/s Thrust, N

Mass flow 

rate, kg/s Thrust, N

a 4.972E-07 3.079E-03 5.475E-08 1.352E-03 4.170E-07 9.261E-03 2.766E-06 1.744E-02 3.767E-06 3.091E-02

b 2.952E-06 5.542E-02 1.266E-06 3.310E-02 1.389E-06 2.755E-02 -7.854E-07 2.948E-02 -3.118E-06 -3.991E-03

c -1.166E-06 -1.066E-02 4.115E-08 1.247E-03 4.858E-07 1.103E-02 2.879E-07 -3.420E-03 1.856E-06 1.990E-02

d 3.296E-07 2.852E-03 6.897E-09 2.258E-04 -7.905E-08 -1.696E-03 -1.514E-08 1.022E-03 -3.184E-07 -3.461E-03

e -3.766E-08 -3.178E-04 -9.746E-10 -2.856E-05 3.951E-09 8.420E-05 -2.446E-10 -7.393E-05 1.752E-08 1.839E-04

Coefficient

7-kW NEXT

High Isp High Thrust

Mass flow 

rate, kg/s Thrust, N

High Isp High Thrust

3-kW Hall 8-kW Hall

NASA/TM—2006-214044 5
From these curve fits, substantial differences in thrust and mass flow rates can be seen. Comparing thrust at 6-

kW, the 3-kW Hall (2), 8-kW Hall, and NEXT generate thrust of roughly 250 mN, 250 or 310 mN, and 200 or 215

mN, respectively. So, for 6-kW of PPU power, the 3-kW Hall can generate 16 or 25% more thrust than NEXT

(depending on which NEXT throttling profile is simulated), and the 8-kW Hall thruster can generate 16 or 55%

more thrust (again, depending on which throttling profile is simulated for each thruster). The differences in mass

flow rates are even more substantial. At 6-kW of PPU power, the 3-kW Hall (2) and 8-kW Hall have mass flow rates

that are 65 or 95% and 65 or 295%, respectively, higher than NEXT. These differences in thrust and mass flow rates

are even larger at full power, and are due to the substantial differences in Isp that exist between the NEXT and Hall

thrusters and are shown in Fig. 5. The peak Isp for the Hall thruster is less than 3,000 seconds, while the Isps for

NEXT are greater than 3,000 seconds for most of the operating range. Substantial differences in efficiency also

exist and are shown in Fig. 6. Note that the system efficiency for NEXT is greater than the 3-kW Hall thruster for

power above roughly 1.5 kW and greater than the 8-kW Hall thruster for essentially the entire operating range.

Despite these differences in specific impulse and efficiency, there are limited regions where the 8-kW Hall

thruster and 7-kW NEXT produce the nearly the same thrust. For example, the NEXT high thrust throttling profile

and the 8-kW Hall high Isp throttling profile generate nearly the same thrust for power less than 4-kW, but the mass

flow rate is higher for the 8-kW Hall thruster when the power is greater than 2-kW.

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

0 1 2 3 4 5 6 7 8 9 10 11 12

Power, kW

Thrust, N

3-kW Hall

8-kW Hall High Isp

8-kW Hall High Thrust

7-kW NEXT High Thrust

7-kW NEXT High Isp

0.0E+00

2.0E-06

4.0E-06

6.0E-06

8.0E-06

1.0E-05

1.2E-05

1.4E-05

1.6E-05

1.8E-05

2.0E-05

0 1 2 3 4 5 6 7 8 9 10 11 12

Power, kW

Mass Flow Rate, kg/s

8-kW Hall High Isp

7-kW NEXT High Thrust

7-kW NEXT High Isp

8-kW Hall High Thrust

3-kW Hall

a) Thrust b) Mass flow rate

Figure 4. 3-kW Hall, 8-kW Hall, and 7-kW NEXT thruster data and curve fits.

1000

1500

2000

2500

3000

3500

4000

4500

0 1 2 3 4 5 6 7 8 9

PPU Power, kW

Specific Impulse, Seconds

3-kW Hall

8-kW Hall, High Thrust

8-kW Hall, High Isp

7-kW NEXT, High Thrust

7-kW NEXT, High Isp

Figure 5. Specific impulse comparison

across each thruster’s operating range.

0.3

0.35

0.4

0.45

0.5

0.55

0.6

0.65

0.7

0 1 2 3 4 5 6 7 8 9

PPU Power, kW

System Efficiency

3-kW Hall

7-kW NEXT, High Isp

7-kW NEXT, High Thrust

8-kW Hall, High Thrust

8-kW Hall, High Isp

Figure 6. System efficiency comparison

across each thruster’s operating range.

NASA/TM—2006-214044 6
Just by looking over Fig. 5 and Fig. 6, one might be inclined to conclude that the performance, measured by

NDM to target for a given trip time, of a higher specific impulse thruster, such as NEXT, would/should be superior

to that of a lower specific impulse thruster, like Hall thrusters, for deep space missions. One might logically assert

that the wet mass of the spacecraft using Hall thrusters would be greater than the wet mass of the spacecraft using

NEXT, therefore the ratio of propellant mass to injected mass would be smaller for NEXT compared to Hall. While

this statement is true, it does not necessarily mean that the NDM to target would be less for a higher thrust thruster.

If the difference in injected mass is large enough between the two propulsion systems, this may overcome the

difference in wet mass and the end result could be a higher final (dry) mass. Intuitively, one would expect the

spacecraft that can produce more thrust throughout the entire trajectory to rely less on the capability of the launch

vehicle for initial escape velocity. The question really becomes whether or not the difference in injected mass is

enough to overcome the difference in the amount of propellant required.

B. SEP Module Modeling

Once the thruster performance has been modeled using two fourth order polynomials, a generic mass model of

the entire SEP module was utilized. This mass model was based on the JPL Titan Orbiter Team X Study.

19 A

SEPTOP refinement enables a simplified model of the system masses to be input into SEPTOP via a series of

coefficients. This refinement also enables a parametric analysis of payload and trip time to be accomplished

efficiently. A top-level equation for the launch mass, MLaunch, was previously developed for input into SEPTOP.

11

!

MLaunch ="power# P + MSEP$Fixed + Mpayload # (1+ kstructure) + Mpropellant # (1+ ktankage)

(2)

Stated in terms of the SEPTOP input variables this equation can be written as:

!

m0 = alfa(1) " P0 + alfa(2) + alfa(7) " (1+ alfa(8)) + mp" (1+ kt). (3)

αpower, kstructure, and ktankage are parameters that model the rate at which their respective systems grow. MSEP-Fixed is the

fixed mass of the SEP module and includes the masses of the propulsion system, structure and cabling for the

propulsion system, and thermal system. It is independent of the array size and propellant load. This parameter

changes only when adding or removing a thruster (or PPU). Table 3 provides the input values for the above

parameters for each electric propulsion system. The exact value of alfa(7), Mpayload, that is input into SEPTOP is

unimportant. What is important is that a solution be found. After that, a time sweep is performed (Mpayload is now a

free variable) and Mpayload is simply solved for in a simple spreadsheet calculation. Table 4 provides the masses of

the major components of the propulsion system. Examples of the complete SEP module mass breakdown are shown

for each target planet in the results section, once the propellant load is determined.

An additional SEPTOP refinement includes a modified model of the propulsion duty cycle. Typically, a 90-95%

propulsion duty cycle is simulated in the trajectory to account for very short coast periods for navigation and

trajectory updates. This duty cycle reduces the available thrust and mass flow rate of the thruster. The modification

that was made decoupled the thrust and mass flow rate duty cycle by adding 5% back to the mass flow rate duty

cycle. This 5% of additional propellant is unavailable for thrust and simulates navigation and trajectory modeling

errors and propellant feed system leakage that takes place throughout the entire thrusting portion of the trajectory.

Table 3 SEPTOP input parameters

Spacecraft 

Assumptions [SEPTOP 

Inputs] 3-kW Hall 8-kW Hall 7-kW NEXT

Fixed mass [alfa(2)] {3 

thrusters/3 PPUs, 4 

thrusters/4 PPUs, 

etc.}, kg

n/a, 376, 423, 

470, 516

519, 625, 731, 

n/a, n/a

511, 627, 742, 

n/a, n/a

Payload [alfa(7)], kg

Payload Structure 

Fraction [alfa(8)] 0.085 0.085 0.085

Tankage (includes 

structure & 5% 

residual, 3% for NEXT) 

[kt]

0.175 0.175 0.137 (3% 

residuals)

Power System Specific 

Mass [alfa(1)], kg/kW 9.9 9.9 9.7

850 - 2,200 depending on target planet and LV

Table 4 Major propulsion system masses (kg)

3-kW Hall 8-kW Hall 7-kW NEXT

CBE This Study CBE This Study CBE This Study

Thruster 3.6 3.9 10.0 10.0 12.4 13.3

PPU 8.4 9.0 25.0 25.0 26.0 26.0

DCIU incorporated into PPU incorporated into PPU/added 

3 kg to total PPU mass 5.65 5.2

Gimbal 2.7 4.64 4.0 3.3 4.64 3.3

Feed System 

Fixed 4.0 4.0 4.0 3.5 2.2 2.2

Feed System Per 

Thruster 1.0 1.0 1.0 2.2 4.1 4.5

Componet

NASA/TM—2006-214044 7
Simulating the trajectory in this manner affects how the propellant contingency mass is book kept. Instead of

adding an 8-10% propellant contingency factor, which is typical for these types of simulations, to the total mass of

the deterministic propellant (to account for leaks, restarts, filling error, residuals, etc.) only a 3-5% propellant

contingency factor is added to the mass of the deterministic propellant; the other 5% is assessed by SEPTOP as

described above. The rest of the pertinent modeling assumptions are listed below.

• Mission Assumptions:

Jupiter epoch: 24 July 2009

Saturn epoch: 30 December 2010

Neptune epoch: 31 March 2010

Venus gravity assist

• Launch Vehicle Assumptions:

Delta IV 4450

Atlas V 551

10% launch mass contingency

• SEP Stage Assumptions:

Single thruster operation permitted

Single thruster redundancy

5% Xenon propellant contingency for flight errors (assessed by SEPTOP)

5% Xenon propellant contingency for reserves (NEXT: 3%)

90% propulsion duty cycle (90% thrust & 95% mass flow rate)

170 W/kg solar arrays (NEXT: 174 W/kg)

5% mass added to solar arrays for power margin

1/R2 array power

2% per year array degradation

250 Watts supplied for housekeeping activities

Xenon tank mass fraction: 4.5% (NEXT: 3.4%)

Power and Thermal Structure: 16%

Stage Cabling (not PPU): 6% of power mass

Tankage Mass Fraction includes structure (26% dry) and Xenon propellant contingency for reserves

Spacecraft Adapter: 42 kg

30% dry mass contingency

C. Launch Vehicle Modeling

Performance data for the Delta IV 4450 and the Atlas V 551 was obtained from the Kennedy Space Center web

site.

20 This performance is modeled with an exponential curve fit and the coefficients and the reference altitude are

input into SEPTOP. The form of the equation is given in Eq. (4) and assumes that the coefficients (A, B, and C) are

all positive and that ΔVL is less than 100 km/s.

!

MLaunch = m0 = Ae

"#VL

C

" B

(4)

The coefficients that were used in this study are given in Table 5; the launch vehicle performance curve fits are

shown in Fig. 7.

NASA/TM—2006-214044 8
D. Trade Studies

Any spacecraft sent on an interplanetary mission must first escape Earth’s gravity well. For this analysis, an

expendable launch vehicle provides the required velocity change to reach a hyperbolic orbit relative to Earth. The

spacecraft builds up its heliocentric orbital energy by raising its aphelion. Thrust from the Hall thrusters efficiently

lowers perihelion for a Venus flyby (modeled as an instantaneous event in SEPTOP). The spacecraft’s velocity

vector is redirected and its kinetic energy is substantially increased as a result of the flyby. This dramatic increase in

kinetic energy as a result of the flyby is largely responsible for increasing the spacecraft’s orbital energy, and relives

the burden that would be placed on the propulsion system without the gravity assist. Once the required ΔV is

obtained, the thrusters are turned off and the SEP module is jettisoned. The spacecraft then starts its coast phase to

reach the target planet.

A limited trade study was conducted on the launch vehicle and the BOL array power to see if the Hall

performance trends varied with either of these parameters relative to the performance capability of NEXT. When the

target planet was Saturn, the same solar array power was chosen for both launch vehicles. This way, the effect of a

larger launch vehicle could be assessed when comparing performance results for different electric propulsion

technologies. When Neptune was targeted, both power and launch vehicles were varied. For the mission to Jupiter,

neither power nor launch vehicles were varied; Hall performance relative to NEXT on a smaller launch vehicle can

be inferred from the Saturn and Neptune performance results.

Optimizing the combination of power and

number of thrusters was not done in this study; an

artificial limit was placed on the maximum

number of operational thrusters. Understandably,

more 3-kW Hall thrusters are required for these

deep space missions as compared to a 7 or 8-kW

thruster, especially for a trip to Neptune. The

maximum and minimum number of operational

thrusters considered for the 8-kW Hall and the 7-

kW NEXT were four and two, respectively. The

maximum number of operational thrusters selected

for the 3-kW Hall was six; the minimum was

Table 6 Jupiter trade study parameters

Atlas V 551

BOL array 

power (1 AU), 

kW

Throttle profile 

(high thrust or 

high Isp)

# operating 

thrusters 

3-kW Hall 15 n/a 3 - 6

8-kW Hall 15 both 2 - 4

7-kW NEXT 15 both 2 - 3

Thruster

Table 5 SEPTOP launch vehicle coefficients and

reference altitude

Launch Vehicle A, kg B, kg C, km/s alt, km

Delta IV 4450 19815.035 5491.56 4.7702 185

Atlas V 551 21753.49 2085.33 3.3997 185

0

1000

2000

3000

4000

5000

6000

7000

8000

-15 -5 5 15 25 35 45 55 65 75 85 95 105 115 125

C3, km2/s2

Injected Mass, kg

Delta IV 4450

Atlas V 551

Figure 7. Launch vehicle performance data and curve fits.

NASA/TM—2006-214044 9
three. All cases assumed a spare thruster and PPU for redundancy. Tables 6-8 summarize the trade-study parameters

for each target planet.

E. Results

All net delivered masses reported in this section include the mass of all elements of the spacecraft (including its

own power source), science payload, any necessary support structure and/or adapter(s). The propellant and SEP

module masses have been subtracted out. Due to convergence issues sometimes associated calculus-of-variations

based optimization codes, like SEPTOP, some of the time spans for a given combination of trade study parameters

are significantly shorter than others; however, there was no difficulty in ascertaining the overall relative

performance.

1. Jupiter

NDM to Jupiter as a function of trip time, number of operating Hall thrusters, and throttle profile (for the 8-kW

Hall thruster) is shown in Fig. 8 and Fig. 9 for the 3-kW and 8-kW Hall thrusters, respectively.

Table 7 Saturn trade study parameters

Delta IV 4450 Atlas V 551

BOL array 

power (1 AU), 

kW

Throttle profile 

(high thrust or 

high Isp)

# operating 

thrusters 

BOL array 

power (1 AU), 

kW

Throttle profile 

(high thrust or 

high Isp)

# operating 

thrusters 

3-kW Hall 18 n/a 3 - 6 18 n/a 3 - 6

8-kW Hall 18 both 2 - 4 18 both 2 - 4

7-kW NEXT 18 both 2 - 4 18 both 2 - 4

Thruster

Table 8 Neptune trade study parameters

Delta IV 4450 Atlas V 551

BOL array 

power (1 AU), 

kW

Throttle profile 

(high thrust or 

high Isp)

# operating 

thrusters 

BOL array 

power (1 AU), 

kW

Throttle profile 

(high thrust or 

high Isp)

# operating 

thrusters 

3-kW Hall 30 n/a 6 22 n/a 6

8-kW Hall 30 both 2 - 4 22 both 2 - 4

7-kW NEXT 30 both 3 - 4 22 both 3 - 4

Thruster

1500

1550

1600

1650

1700

1750

1800

1850

1900

1950

2000

2050

3.5 4 4.5 5 5.5 6

Trip Time, Years

Net Delivered Mass to Jupiter, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 15 kW

5+1

4+1

3+1

6+1

3-kW Hall

Figure 8. NDM to Jupiter as a function of trip

time and number of operating thrusters for the

spacecraft using 3-kW Hall thrusters.

1500

1550

1600

1650

1700

1750

1800

1850

1900

1950

2000

2050

3.5 4 4.5 5 5.5 6

Trip Time, Years

Net Delivered Mass to Jupiter, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 15 kW

8-kW Hall High Thrust, 2+1

High Thrust, 3+1

High Thrust, 4+1

High Isp, 2+1

Figure 9. NDM to Jupiter as a function of trip

time and number of operating thrusters for the

spacecraft using 8-kW Hall thrusters.

NASA/TM—2006-214044 10
As Fig. 8 shows, the 5+1 thruster configuration is the top performer and is selected for comparison with the

capability of NEXT. Figure 9 shows that the 8-kW Hall high thrust throttle profile outperforms the high Isp throttle

profile; the 2+1 thruster configuration yields the largest NDM and is selected for comparison with NEXT. Figure 10

shows the results for the NEXT propelled spacecraft. The largest masses delivered by the NEXT propelled

spacecraft use the high thrust throttle profile. So, by looking at Fig. 8-10, one can conclude that the optimal specific

impulses for this mission scenario are below 3,000 seconds. This fact is clearly indicated in Fig. 11, which shows the

Hall and NEXT delivered mass and arrival V∞ comparison.

Both versions of the Hall thruster deliver at least 175 kg more dry mass than NEXT for the trip times shown. Fewer

8-kW Hall thrusters are needed than 3-kW Hall thrusters to deliver nearly the same mass. The spacecraft using Hall

propulsion arrives with slightly less V∞ than the NEXT propelled spacecraft for transfer times less than four years.

They have roughly the same arrival V∞ at four years, after which NEXT system provides slightly lower V∞.

Table 9 provides the breakdown of Hall and NEXT system masses for a 3.8-year transfer time. The Atlas launch

vehicle injects an average of 23% more mass for the spacecraft using Hall thrusters than NEXT—more than enough

to make up for the difference in wet mass. For the same transfer time, a trajectory and power and thruster variation

profile is shown in Fig. 12 and Fig. 13, respectively, for the 8-kW

Hall thruster.

Starting with two thrusters, the optimal thruster variation is 2-1-

2-0 (short coast period)-2-1 thrusters. At the highest A.U. portion

of the trajectory, one thruster provides the necessary thrust for

lowering perihelion. For most of the trajectory the spacecraft flies

with only a single thruster operating. Two thrusters operate at full

power only once during the transfer—during the Venus flyby.

Approximately 120 days after the gravity assist, the thrusters turn

off and the SEP module is jettisoned, thus the spacecraft begins its

coast phase to Jupiter.

The fact the spacecraft operates only one thruster at its high

radial distances from the sun is not a given. A brief discussion on

why a spacecraft would operate more than one thruster during this

region may be worthwhile. Total PPU power for most of this

region is less than 5-kW. Let’s suppose that the spacecraft operates

two thrusters for a portion of this region. This means it would have

to split this power between those two thrusters. The spacecraft

could operate both thrusters at 2-kW. The thrust produced by these

electric thrusters is given by Eq. (5).

21 At this lower power setting,

the ratio of overall efficiency to specific impulse may or may not

be larger than the same ratio at 4-kW. This means that it is very

possible that two thrusters operating at 2-kW each produce more

1500

1550

1600

1650

1700

1750

1800

1850

1900

1950

2000

2050

3.5 4 4.5 5 5.5 6

Trip Time, Years

Net Delivered Mass to Jupiter, kg

6

6.2

6.4

6.6

6.8

7

7.2

7.4

7.6

7.8

8

Jupiter Arrival Vinf, km/s

3-kW Hall Arrival Vinf

8-kW Hall Arrival Vinf

7-kW NEXT Arrival Vinf Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 15 kW

3-kW Hall, 5+1

8-kW Hall, 2+1, High Thrust

7-kW NEXT, 2+1, High Thrust

Arrival Vinf

Figure 11. NDM to Jupiter as a function of trip time

for the top performing Hall and NEXT propelled

spacecraft.

1500

1550

1600

1650

1700

1750

1800

1850

1900

1950

2000

2050

3.5 4 4.5 5 5.5 6

Trip Time, Years

Net Delivered Mass to Jupiter, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 15 kW

7-kW NEXT

High Thrust, 2+1

High Isp, 2+1

High Isp, 3+1

High Thrust, 3+1

Figure 10. NDM to Jupiter as a function of trip

time and number of operating thrusters for the

spacecraft using 7-kW NEXT.

Table 9 Electric propulsion system

masses for a 3.8-year transfer to

Jupiter when launching on an Atlas

V 551

System/parameter 3-kW Hall 8-kW Hall 7-kW NEXT

# of thrusters/PPUs 6 3 3

NDM/Payload, kg 2030 2022 1834

Attitude Control, kg 3.0 3.0 3.0

Power, kg 127.6 127.6 101.7

Propulsion, kg 132.5 167.4 167.0

Structure, kg 324.8 319.9 286.9

Cabling, kg 54.0 49.4 45.0

Thermal, kg 54.6 76.9 62.8

SEP dry, kg 696.5 744.2 666.4

Contingency 30% 30% 30%

SEP dry w/contingencies 951 1018 881

Xe (includes 5% 

contingency for nav 

errors -assessed by 

SEPTOP), kg

909 1018 503

Xe contingency (other) 5% 5% 3%

Xe contingency (other), 

kg 45 51 15

Total launch mass, kg 3890 4057 3219

LV contingency 10% 10% 10%

C3, km2/s2 21.3 19.0 31.5

mp/m0 0.25 0.26 0.16

NASA/TM—2006-214044 11
(or less) total thrust than one thruster operating a 4-kW. The spacecraft propelled by 8-kW Hall thrusters operates a

single thruster during this very power-limited region because the thrust generated is sufficient. It should not be

assumed that any electric thruster would exhibit this characteristic; operating more than one thruster could generate

more thrust and may be required.

In terms of the propellant mass fraction, NEXT is more efficient. However, this did not result in a higher final

mass than the Hall propelled spacecraft. As a consequence of the relatively lower thrust produced by the NEXT

propelled spacecraft, more energy is required from the launch vehicle relative to the Hall propelled spacecraft.

!

T =

2"sysP

gIsp

, N (5)

2. Jupiter Summary

For the selected array power, launch vehicle and trajectory, the advanced Hall thrusters’ specific impulses are

more optimal than NEXT. Using Hall thrusters for the spacecraft’s primary propulsion system enables at least an

additional 175 kg to be delivered to Jupiter relative to the NEXT propelled spacecraft. The spacecraft using the 8-

kW Hall thrusters delivers slightly less mass than the 3-kW Hall propelled spacecraft; however, it does so with three

less operating thrusters. The average throughput per thruster for this configuration is over 500 kg.

3. Saturn

The trade-offs between the number of operating thrusters, throttling profile (when applicable), trip time, and

NDM for the 3-kW Hall, 8-kW Hall and 7-kW NEXT when launching on a Delta IV 4450 are shown in Fig. 14-16.

Figures 17-19 show the same trade studies when the spacecraft is launched on an Atlas V 551. BOL array power

was set to 18-kW.

-6

-4

-2

0

2

4

6

-6 -4 -2 0 2 4 6

Launch: 2 July 2009

Venus flyby: 17 March 2011

Arrival: 20 April 2013

!V: 6.0 km/s

Figure 12. 3.8-year spacecraft trajectory

to Jupiter using 8-kW Hall propulsion

(bold curve depicts thrusting).

0

5

10

15

20

25

30

0 150 300 450 600 750 900

Trip Time, Days

Power, kW

0

1

2

Number of Operating Thrusters

# of thrusters

1/R2 array power

Total PPU power short coast period

Figure 13. Spacecraft power and 8-kW Hall

thruster variation history for a 3.8-year transfer to

Jupiter.

NASA/TM—2006-214044 12
800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

Delta IV 4450, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

3-kW Hall

3+1

5+1

4+1

6+1

Figure 14. NDM to Saturn as a function of trip time

and number of operating thrusters for the spacecraft

using 3-kW Hall thrusters when launching on a Delta

IV 4450.

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

Delta IV 4450, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

8-kW Hall

2+1, High Isp

2+1, High Thrust

3+1, High Isp

3+1, High Thrust

4+1, High Isp

4+1, High Thrust Figure 15. NDM to Saturn as a function of trip time,

throttle profile, and number of operating thrusters

for the spacecraft using 8-kW Hall thrusters when

launching on a Delta IV 4450.

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

Delta IV 4450, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

7-kW NEXT

3+1, High Thrust

2+1, High Thrust

4+1, High Thrust

3+1, High Isp (circles)

2+1, High Isp

4+1, High Isp

Figure 16. NDM to Saturn as a function of trip time,

throttle profile, and number of operating thrusters for

the spacecraft using 7-kW NEXT when launching on a

Delta IV 4450.

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

3-kW Hall

6+1

3+1

4+1

5+1

Figure 17. NDM to Saturn as a function of trip

time and number of operating thrusters for the

spacecraft using 3-kW Hall thrusters when

launching on an Atlas V 551.

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

8-kW Hall 2+1, High Thrust

2+1, High Isp

3+1, High Isp 4+1, High Thrust

3+1, High Thrust

4+1, High Isp

Figure 18. NDM to Saturn as a function of trip

time, throttle profile, and number of operating

thrusters for the spacecraft using 8-kW Hall

thrusters when launching on an Atlas V 551.

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

7-kW NEXT

3+1, High Thrust

2+1, High Isp

3+1, High Isp

2+1, High Thrust

4+1, High Thrust

4+1, High Isp

Figure 19. NDM to Saturn as a function of trip

time, throttle profile, and number of operating

thrusters for the spacecraft using 7-kW NEXT

when launching on an Atlas V 551.

NASA/TM—2006-214044 13
By comparing the performance results of the spacecraft propelled by the 3-kW Hall thrusters when launched on the

Delta IV 4450 or Atlas V 551 (Fig. 14 and Fig. 17), one can see greater sensitivity to the number of operating

thrusters when launched on an Atlas V 551. Switching launch vehicles increases the NDM of the 5+1 thruster

configuration by roughly 475 kg.

The same comparison for the spacecraft using 8-kW Hall thrusters (Fig. 15 and Fig. 18) is interesting for two

reasons. First, notice that when the spacecraft is launched on the Delta IV 4450, performance is primarily dependant

on the number thrusters. This indicates a strong sensitivity to the dry mass. In fact, reducing the total number of

thrusters from four to three results in an additional 200 kg of dry mass at Saturn. Secondly, notice that when the

spacecraft is launched on the Atlas V 551, performance is primarily dependant on the throttle profile. The high

thrust throttle profile yields the best results regardless of the total number of thrusters. This indicates a strong

sensitivity to the thrusters’ specific impulses. Roughly an additional 550 kg is delivered to Saturn when the

spacecraft is launched on the Atlas V 551.

In regards to these same types of comparisons for the NEXT propelled spacecraft when launched on the Delta IV

4450 (Fig. 16), the NDM to Saturn is primarily dependant on the throttle profile for most of the transfer times. A

NDM gain of roughly 60 kg is achieved just by changing the throttle profile from high Isp to high thrust. Figure 19

shows the same throttle profile sensitivity for the NEXT propelled spacecraft launched on an Atlas V 551. Switching

launch vehicles results in roughly a 350 kg gain in the NDM to Saturn. Due to the relatively lower thrust of the

NEXT propelled spacecraft, more of the capability of the Atlas launch vehicle is spent on a higher injection velocity.

This results in a performance increase of roughly 125 kg and 200 kg less than what the 3-kW Hall and 8-kW Hall

propelled spacecraft achieved when using a larger launch vehicle.

The top performance curves from Fig. 14-19 were selected for comparison and are shown in Fig. 20 and Fig. 21.

When the spacecraft is launched on a Delta IV 4450, two operational 8-kW Hall thrusters and five operational 3-kW

Hall thrusters deliver essentially the same mass to Saturn as two or three operational 7-kW NEXTs for transfer times

greater than 7-years. For transfer times less than 7-years, the 8-kW Hall thruster offers the benefit of delivering

nearly the same mass as NEXT with one fewer thruster than NEXT and three fewer thrusters than the 3-kW Hall.

Looking at Fig. 21, one can see that the Hall thrusters exhibit greater sensitivity to the launch vehicle

performance capability. Just by switching launch vehicles, the differences in final masses between the two electric

propulsion systems are substantial. The 8-kW Hall thrusters enable the spacecraft to deliver roughly 200 kg more,

while using one less thruster, to Saturn than the NEXT propulsion system. For the transfer times shown, average

throughput per thruster for the 8-kW Hall varies by 8% and never exceeds 635 kg.

Tables 10 and 11 provide each electric propulsion system masses for a 5.7-year transfer to Saturn. For this same

transfer time, a trajectory and thruster variation history profile is shown in Fig. 22 and Fig. 23 when the spacecraft

utilizes 8-kW Hall thrusters and launches on an Atlas V 551.

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

4

5

6

7

8

9

10

11

12

13

14

Arrival Vinf, km/s

Delta IV 4450, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

Arrival Vinf

7-kW NEXT, 2+1, High 

Thrust

7-kW NEXT, 3+1, High 

Thrust

8-kW Hall, 2+1, High Isp

3-kW Hall, 5+1

Figure 20. NDM to Saturn as a function of trip time for

the top performing Hall and NEXT propelled spacecraft

when launching on a Delta IV 4450.

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

4 5 6 7 8 9 10 11 12

Trip Time, Years

Net Delivered Mass to Saturn, kg

4

5

6

7

8

9

10

11

12

13

14

Saturn Arrival Vinf, km/s

Arrival Vinf

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 18 kW

8-kW Hall, 2+1, High Thrust

3-kW Hall, 6+1

7-kW NEXT, 3+1, High Thrust

Figure 21. NDM to Saturn as a function of trip time

for the top performing Hall and NEXT propelled

spacecraft when launching on an Atlas V 551.

NASA/TM—2006-214044 14
The spacecraft using NEXT has lower propellant mass fractions than the Hall propelled spacecraft when

launched on either launch vehicle. However, this not does translate to better performance due to the difference in

required launch energies (C3) that are direct result of the relatively low thrust produced by the ion thrusters. A single

8-kW Hall thruster generates the necessary thrust to lower perihelion for a flyby of Venus as shown in Fig. 23.

Table 10 Electric propulsion system masses

for a 5.7-year transfer to Saturn when

launching on a Delta IV 4450

System/parameter 3-kW Hall 8-kW Hall 7-kW NEXT

# of thrusters/PPUs 6 3 4

NDM/Payload, kg 1364 1373 1362

Attitude Control, kg 3.0 3.0 3.0

Power, kg 146.2 146.2 119.8

Propulsion, kg 131.8 158.9 218.6

Structure, kg 286.0 275.0 285.2

Cabling, kg 52.8 47.3 52.6

Thermal, kg 54.6 76.9 78.8

SEP dry, kg 674.3 707.3 758.0

Contingency 30% 30% 30%

SEP dry w/contingencies 921 961 1005

Xe (includes 5% 

contingency for nav 

errors -assessed by 

SEPTOP), kg

894 837 637

Xe contingency (other) 5% 5% 3%

Xe contingency (other), 

kg 45 42 19

Total launch mass, kg 3180 3171 3003

LV contingency 10% 10% 10%

C3, km2/s2 11.8 11.9 14.3

mp/m0 0.30 0.28 0.22

Table 11 Electric propulsion system masses

for a 5.7-year transfer to Saturn when

launching on an Atlas V 551

System/parameter 3-kW Hall 8-kW Hall 7-kW NEXT

# of thrusters/PPUs 7 3 4

NDM/Payload, kg 1817 1897 1615

Attitude Control, kg 3.0 3.0 3.0

Power, kg 146.2 146.2 119.8

Propulsion, kg 157.1 179.1 220.1

Structure, kg 334.3 326.0 302.9

Cabling, kg 60.2 51.6 53.8

Thermal, kg 62.0 76.9 78.9

SEP dry, kg 762.9 782.8 778.5

Contingency 30% 30% 30%

SEP dry w/contingencies 1048 1081 1033

Xe (includes 5% 

contingency for nav 

errors -assessed by 

SEPTOP), kg

1132 1264 683

Xe contingency (other) 5% 5% 3%

Xe contingency (other), 

kg 57 63 20

Total launch mass, kg 3997 4242 3331

LV contingency 10% 10% 10%

C3, km2/s2 19.8 16.5 29.7

mp/m0 0.30 0.31 0.21 -12

-10

-8

-6

-4

-2

0

2

4

6

8

10

-10 -5 0 5 10

Launch: 8 Jan 2011

Venus flyby: 20 Sept 2012

Arrival: 20 Sept 2016

!V: 7.4 km/s

Figure 22. 5.7-year spacecraft trajectory

to Saturn using 8-kW Hall propulsion

(bold curve depicts thrusting).

0

5

10

15

20

25

30

35

40

0 200 400 600 800 1000

Trip Time, Days

Power, kW

0

1

2

Number of Operating Thrusters

# of thrusters

1/R2 array power

Total PPU power

Figure 23. Spacecraft power and 8-kW Hall thruster

variation history for a 5.7-year transfer to Saturn.

NASA/TM—2006-214044 15
4. Saturn Summary

The NEXT and Hall propelled spacecraft deliver roughly the same mass to Saturn when launched on a Delta IV

4450. For transfer times less 7-years, one less 8-kW Hall thruster is required relative to NEXT.

Switching to a launch vehicle with an increased performance capability (and holding BOL solar array power

constant) shows that the Hall thrusters’ performance is more sensitive to the larger launch vehicle than the NEXT

propulsion system for this mission scenario. A 200 kg delivered mass advantage for the Hall propelled spacecraft

relative to the NEXT propelled spacecraft is achieved with the Atlas V 551. The 8-kW Hall thruster performed well

for this mission scenario; it offers a substantial mass advantage with one fewer thruster than NEXT. The trade-off

for this increased performance is a relatively high average throughput per thruster (> 600 kg).

5. Neptune

The same kind of trade studies that were performed for a Saturn mission was performed for a Neptune mission

with one exception: BOL solar array power was reduced roughly 25% when launched on the Atlas V 551. This way,

a trade-off between two different SEP mission architectures could be assessed. One mission architecture would

require a relatively large BOL solar array combined with the performance of a relatively smaller launch vehicle, the

other visa versa. The spacecraft performance results when launched on a Delta 4450 and an Atlas V 551 are shown

in Fig. 24 & 25, and Fig. 26 & 27, respectively. Because of the higher power requirement for a larger ΔV mission to

Neptune, the trade-off between the number of operational 3-kW Hall thrusters was not performed—only the

performance of the 6+1 thruster configuration was assessed.

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

8.5 9 9.5 10 10.5 11 11.5 12 12.5 13 13.5 14

Trip Time, Years

Net Delivered Mass to Neptune, kg

Delta IV 4450, LV Contingency: 10%

BOL Array Power (1 AU): 30 kW

8-kW Hall

3+1, High Isp

4+1, High Thrust

3+1, High Thrust

2+1, High Isp

4+1, High Isp

Figure 24. NDM to Neptune as a function of trip

time, throttle profile and number of operating

thrusters for the spacecraft using 8-kW Hall

thrusters when launching on a Delta IV 4450.

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

8.5 9 9.5 10 10.5 11 11.5 12 12.5 13 13.5 14

Trip Time, Years

Net Delivered Mass to Neptune, kg

Delta IV 4450, LV Contingency: 10%

BOL Array Power (1 AU): 30 kW

7-kW NEXT

4+1, High Thrust

4+1, High Isp

3+1, High Thrust

3+1, High Isp

Figure 25. NDM to Neptune as a function of trip

time, throttle profile and number of operating

thrusters for the spacecraft using 7-kW NEXT

when launching on a Delta IV 4450.

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

8.5 9 9.5 10 10.5 11 11.5 12 12.5 13 13.5 14

Trip Time, Years

Net Delivered Mass to Neptune, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 22 kW

8-kW Hall 3+1, High Thrust

4+1, High Thrust

2+1, High Thrust

3+1, High Isp

4+1, High Isp

2+1, High Isp

Figure 26. NDM to Neptune as a function of trip

time, throttle profile and number of operating

thrusters for the spacecraft using 8-kW Hall

thrusters when launching on an Atlas V 551.

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

8.5 9 9.5 10 10.5 11 11.5 12 12.5 13 13.5 14

Trip Time, Years

Net Delivered Mass to Neptune, kg

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 22 kW

7-kW NEXT

4+1, High Thrust

3+1, High Isp

4+1, High Isp 3+1, High Thrust

Figure 27. NDM to Neptune as a function of trip

time, throttle profile and number of operating

thrusters for the spacecraft using 7-kW NEXT

when launching on an Atlas V 551.

NASA/TM—2006-214044 16
Although the optimal transfers time to Neptune are more than 13 years, as shown in Figures 24-27, past mission

analysis work has focused on delivering an average payload of nearly 1,000 kg to Neptune in 10-11 years.

11,22

Figure 24 shows that the NDM to Neptune (Delta IV 4450 launcher) is primarily dependant on the 8-kW Hall

throttle profile. The high Isp throttle profile yields a larger NDM than the high thrust profile. When the BOL array

power is reduced from 30 kW to 22 kW and the launch vehicle changed from a Delta IV 4450 to an Atlas V 551

(Fig. 26), the Hall propelled spacecraft can reach Neptune several months sooner or deliver roughly an additional

300 kg in final mass for transfer times between 10-11 years by using the high thrust throttle profile. So, using a

launch vehicle with a greater performance capability and reducing the array power resulted in a modest decrease of

the mission’s optimal specific impulses (the specific impulse differences between the high Isp and high thrust throttle

profile).

The NEXT propelled spacecraft does not exhibit the same results. When the spacecraft is launched on the Delta

IV 4450 or an Atlas V 551, performance is primarily dependant on the total number of thrusters; four operational

thrusters enable a larger delivered mass than three. Better performance is achieved for the high thrust throttle

profiles. The high Isp throttle profile may deliver a higher mass than the high thrust throttle profile for transfer times

on the order of 13 years or longer. Although the performance capability of the Atlas V 551 is significantly more than

the Delta IV 4450, the spacecraft’s performance actually is reduced because of its sensitivity to power. This results

in additional couple months of trip time or a rough average of 50 kg less mass at Neptune.

Once again, the best performance curves from Fig. 24-27 are selected for comparison and are shown in Fig. 28

and Fig. 29

From viewing these two figures, one can see that the spacecraft propelled by Hall thrusters is more sensitive to

the additional performance capability of the launch vehicle, and less sensitive to the BOL array power. In contrast,

the NEXT propelled spacecraft shows greater sensitivity to BOL array power. When launched on a Delta IV 4450,

NEXT outperforms the Hall system; the opposite is true when the launch vehicle is the Atlas V 551 and power is

reduced. When the spacecraft is launched on the Atlas V 551, 8-kW Hall thrusters enable a trip time reduction on

the order of a year or more or about a 300 kg gain in final mass at Neptune compared to the NEXT propelled

spacecraft. Notice at this reduced power level, the performance of the 3-kW Hall thrusters and NEXT are nearly

identical.

Because the maximum power of the 8-kW Hall and 7-kW NEXT differ by ~ 1-kW, four operational 8-kW Hall

can process an additional 4-kW of power relative to NEXT inside 1 A.U. To assess what kind of effect this

additional available power would have on performance, the maximum power for the 8-kW Hall thrusters was

constrained to 7.3-kW for the 4+1, Atlas V 551 case. The mass delivered to Neptune for this case was reduced by

3%. This small performance decrease is consistent with the performance shown by the 3-kW Hall thrusters; its mass

delivery capability is nearly the same as NEXT while the maximum power processed is 11-kW less than NEXT

(inside 1 A.U.).

A SEP mass breakdown for a 10.7-year transfer to Neptune when launched on a Delta IV 4450 is shown in Table

12. Table 13 provides the same mass breakdown when the spacecraft is launched on an Atlas V 551. When the

NEXT propelled spacecraft is launched on the Atlas V 551, the increment in required launch energy is large enough

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

8.5 9.5 10.5 11.5 12.5 13.5

Trip Time, Years

Net Delivered Mass to Neptune, kg

10

11

12

13

14

15

16

17

18

19

20

Neptune Arrival Vinf, km/s

Delta IV 4450, LV Contingency: 10%

BOL Array Power (1 AU): 30 kW

Arrival Vinf

7-kW NEXT, 4+1, High Thrust

8-kW Hall, 3+1, High Isp

3-kW Hall, 6+1

Figure 28. NDM to Neptune as a function of trip

time for the top performing Hall and NEXT

propelled spacecraft when launching on a Delta

IV 4450.

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

8.5 9 9.5 10 10.5 11 11.5 12 12.5 13 13.5 14

Trip Time, Years

Net Delivered Mass to Neptune, kg

10

11

12

13

14

15

16

17

18

19

20

Neptune Arrival Vinf, km/s

Atlas 551, LV Contingency: 10%

BOL Array Power (1 AU): 22 kW

8-kW Hall, 3+1, High Thrust 8-kW Hall, 4+1, High Thrust

3-kW Hall, 6+1

7-kW NEXT, 4+1, High Thrust

Arrival Vinf

Figure 29. NDM to Neptune as a function of trip

time for the top performing Hall and NEXT

propelled spacecraft when launching on an Atlas

V 551.

NASA/TM—2006-214044 17
that the launch vehicle actually injects less mass than the Delta IV 4450. This is because 22-kW is far from its

optimal power level. In contrast, the total thrust produced by the Hall propelled spacecraft is sufficient to handle

this reduction in power.

Figure 30 shows the trajectory profile for a 10.7-year transfer to Neptune and Fig. 31 provides the corresponding

power and thruster variation history. All three thrusters are turned on only for two relatively short periods—at the

beginning of the trajectory and during the Venus flyby. Dual thruster operations are used primarily for throttling

from three to two thrusters and visa versa.

-6

-4

-2

0

2

-2 2 6 10 14 18 22 26 30

Launch: March 21, 2010

Venus flyby: May 14, 2012

Arrival: Dec. 1, 2020

!V: 10.5 km/s

Figure 30. 10.7-year spacecraft trajectory to Neptune using 8-kW Hall

propulsion (bold curve depicts thrusting).

Table 12 Electric propulsion system masses

for a 10.7-year transfer to Neptune when

launching on a Delta IV 4450

System/parameter 8-kW Hall 7-kW NEXT

# of thrusters/PPUs 4 5

NDM/Payload, kg 946 1004

Attitude Control, kg 3.0 3.0

Power, kg 220.3 192.2

Propulsion, kg 215.4 275.1

Structure, kg 295.4 305.5

Cabling, kg 60.2 65.1

Thermal, kg 97.1 94.8

SEP dry, kg 891.4 935.7

Contingency 30% 30%

SEP dry w/contingencies 1221 1244

Xe (includes 5% 

contingency for nav 

errors -assessed by 

SEPTOP), kg

1240 909

Xe contingency (other) 5% 3%

Xe contingency (other), 

kg 62 27

Total launch mass, kg 3407 3156

LV contingency 10% 10%

C3, km2/s2 8.8 12.1

mp/m0 0.38 0.30

Table 13 Electric propulsion system masses

for a 10.7-year transfer to Neptune when

launching on an Atlas V 551

System/parameter 3-kW Hall 8-kW Hall 7-kW NEXT

# of thrusters/PPUs 7 4 5

NDM/Payload, kg 965 1252 956

Attitude Control, kg 3.0 3.0 3.0

Power, kg 170.9 170.9 143.9

Propulsion, kg 162.1 232.9 269.3

Structure, kg 290.3 322.2 287.5

Cabling, kg 59.3 59.9 60.8

Thermal, kg 62.0 97.1 94.8

SEP dry, kg 747.6 886.0 859.3

Contingency 30% 30% 30%

SEP dry w/contingencies 1034 1232 1139

Xe (includes 5% 

contingency for nav 

errors -assessed by 

SEPTOP), kg

1238 1612 745

Xe contingency (other) 5% 5% 3%

Xe contingency (other), 

kg 62 81 22

Total launch mass, kg 3236 4096 2840

LV contingency 10% 10% 10%

C3, km2/s2 31.2 18.5 38.1

mp/m0 0.40 0.41 0.27

NASA/TM—2006-214044 18
6. Neptune Summary

When launched a Delta IV 4450, the NEXT propelled spacecraft delivers ~30 kg more than the Hall propelled

spacecraft for 10-11 year transfer times. For longer trip times, the mass advantage increases.

Selecting a launch vehicle with more performance capability and decreasing solar array power by roughly 25%

reveals something important: an electric thruster that produces relatively higher thrust can enable a reduction in solar

array power and the number of operational thrusters. The thrust produced by NEXT is simply not high enough to

take advantage of this; therefore, the Hall propelled spacecraft outperforms the NEXT propelled spacecraft by

roughly 300 kg for 10-11 year transfer times.

This result is important because if the delivered mass requirement grows significantly enough to necessitate the

use of a launch vehicle with more performance capability, an electric thruster that produces relatively more thrust

could enable a reduction in power and required thrusters that a higher specific impulse thruster could not. In fact,

since the time this analysis was performed, a detailed Neptune mission trade study has established a SEP baseline

trajectory that must deliver ~2,230 kg to Neptune’s SOI—over a 100% mass growth from previous analysis.

23 To

meet this high mass requirement, that study used two gravity assists, a Delta IV 4050H launch vehicle, and five

operational NEXT thrusters that use the high Isp throttle profile.

IV.  Conclusion

This study is believed to be the first of its kind in that it assessed the performance of deep space SEP missions

using next generation Hall thrusters for primary propulsion applications—mission scenarios previously reserved for

thrusters that generate higher specific impulses. In order to access the performance of a spacecraft using advanced

Hall thrusters as its primary propulsion system, trade-studies were conducted on mission parameters such as thruster

count, throttle profile, and launch vehicle to identify the spacecraft configuration that delivered the largest mass for a

range of trip times to each of three target planets: Jupiter, Saturn, and Neptune. These same trade-studies were

conducted for a spacecraft that utilized the ASOA NEXT ion system for its primary propulsion system and the top

performing spacecraft configuration was also identified. A relative comparison was made between the existing

NEXT system and the projected next generation Hall primary propulsion systems.

Results in this study show that for four of out the five mission scenarios assessed, a spacecraft propelled by an

advanced Hall thruster (of the type under development) offered at least one of the following benefits relative to the

existing higher Isp NEXT:

(1) Better performance—larger payload or shorter flight time;

(2) Fewer thrusters;

(3) Less BOL array power.

0

10

20

30

40

50

60

0 200 400 600 800 1000 1200

Trip Time, Days

Power, kW

0

1

2

3

Number of Operating Thrusters

# of thrusters

1/R2 array power

Total PPU power

Figure 31. Spacecraft power and 8-kW Hall thruster variation history for a

10.7-year transfer to Neptune.

NASA/TM—2006-214044 19
These promising advantages of advanced Hall thrusters should be realized in the context that every mission is

unique, and that this study only provides a snapshot of the relative performance. Because each mission is unique,

it will have specific science objectives and cost-caps that will strongly affect important parameters such as the

launch vehicle and array power. And as this study shows, the combination of launch vehicle and array power has

dramatic consequences when comparing the performance of the advanced Hall thrusters and the ASOA NEXT

ion system. One should not assume a priori that advanced Hall thrusters would offer one or more of these

benefits for every deep space mission; more analyses needs to be performed.

Having hopefully placed these results in the proper context, several important conclusions can be drawn:

(1) The optimal specific impulses for SEP deep space missions of the kind that this study considers are

determined by various mission design parameters including destination, required payload, power, and

launch vehicle. However, this study finds that advanced Hall thrusters offer the more optimal Isps for

most of the mission scenarios considered.

(2) Increased launch vehicle performance capability tends to significantly lower these missions’ optimal

specific impulse. As a result, a mission designer could choose to take on the added propellant load

that is inherent with a higher thrust system, but have enough of a performance margin to reduce

power and therefore cost.

(3) Higher specific impulses do not necessarily translate to better performance. Often, better

performance is achieved with the higher thrust throttle profile of NEXT rather than the high Isp

throttle profile.

(4) The relatively higher thrust generated by the Hall thrusters enables a reduction in solar array power

requirements that a higher Isp thruster cannot.

(5) For missions to Jupiter and beyond, the 3-kW Hall thruster is not appropriate because it requires

several more operational thrusters than the 7-kW NEXT and 8-kW Hall. Additionally, operation at

power levels below 3 kW is not driven by the available power from the sun.

(6) Increased lifetime capabilities are essential for optimal performance of the 8-kW Hall thrusters.

Average throughput per thruster requirements is on the order of 400 – 600 kg.

(7) The mission benefits provided by advanced Hall thrusters may warrant the future development of

propulsion systems based on this emerging technology if thruster throughput can be adequately

addressed.

These findings are important because if the advanced Hall thruster throughput can be adequately addressed, the

application of advanced Hall thrusters to deep space SEP missions potentially enables a reduction in mission class.

For example, if the science/payload objectives of a Flag Ship class mission can be met with fewer thrusters and less

array power, the resulting cost-savings may be substantial enough that the mission becomes a New Frontiers class

mission. Launch vehicle costs would also have to be considered. Of course, cost analyses would need to be

performed, and more mission analyses would need to be done, but the potential exists.

Other possible advanced Hall thruster applications include using the Hall thrusters to spiral out from the Earth’s

gravity well and for the heliocentric transfer. This scenario would reduce the size of the launch vehicle required and

the complexity associated with carrying a different propulsion system for the heliocentric transfer. However, spare

thrusters may be needed if throughput requirements become an issue.

In light of this analysis, one must wonder what the results would be of a similar trade-study for SEP missions in

the inner solar system: missions such as a comet surface sample return (CSSR) or another type of small body

rendezvous. These missions may or may not require a gravity assist. The BOL array power requirements may be

substantially less than missions to the outer solar system. The launch vehicle performance capability that is required

to enable these types of missions to small bodies may be less than missions to the outer solar system. Despite these

variants, recent mission analyses with the objective of assessing the applicability of electric thrusters such as NEXT

and Hall for Discovery class missions have shown similar results: advanced Hall thrusters can offer a performance

advantage relative to NEXT (and NSTAR).

24 Additionally, that study found that advanced Hall thrusters could offer

a cost benefit.

As a final closing remark, although this study showed substantial performance benefits for advanced Hall

thrusters relative to NEXT, the technologies are at different maturity levels. The advanced Hall thrusters are only at

TRL 4, and have not yet been fully validated. Substantiation of these performance benefits requires further technical

development.

NASA/TM—2006-214044 20
References 1

Gulczinski, F. S., Analysis of Hall-Effect Thrusters and Ion Engines for Orbit Transfer Missions,” 32nd

AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA 96-2973, July 1-3, 1996. 2

Leifer, S., and Noca, M.,“Negative C3 Launch Options for Solar System Exploration,” AIAA 98-3925, 1998. 3

Gulczinski, F. S., Spores, R. A., and Stuhlberger, J., “In-Space Propulsion,” AIAA/ICAS International Air and Space 

Symposium and Exposition: The Next 100 Years,” AIAA 2003-2588, July 14-17, 2003. 4

Oleson, S. R., and Sankovic, J. M., “Electric Propulsion for Low Earth Orbit Constellations,” NASA TM-208821, 1998. 5

Gefert, L. P., Hack, K. J., and Kerslake, T. W., Options for the Human Exploration of Mars using Solar Electric Propulsion, 

AIP Conference Proceedings, No. 458, pp 1275-1280, STAIF-99, Nov. 1999. 6

Brophy, J.R., and Noca, M., “Electric Propulsion for Solar System Exploration,” Journal of Propulsion and Power, Vol. 14, 

No. 5, 1998, pp. 700-707. 7

Oleson, S. R., and Sankovic, J. M., “Advanced Hall Electric Propulsion for Future In-Space Transportation,” NASA TM210676, 2001.

8

Oleson, S. R., “Mission Advantages of Constant Power, Variable Isp Electrostatic Thrusters,”, 36th

AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA 2000-3413, July 16-19, 2000. 9

Fiehler, D., and Oleson, S., “A Comparison of Electric Propulsion Systems for Mars Exploration,” 39th

AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA 2003-4574, July 20-23, 2003. 10Oh, D., Benson, S., Witzberger, K., and Cupples, M., “Deep Space Mission Applications for NEXT: NASA’s Evolutionary 

Xenon Thruster,” 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA 2004-3806, July 11-14, 2004. 11Oleson, S., Gefert, L., Benson, S., Patterson, M., Noca, M., Sims, J., “Mission Advantages of NEXT: NASA’s Evolutionary 

Xenon Thruster,” 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA 2002-3969, July 7-10, 2002. 12Integrated Technology Plan for the Civil Space Program, JCM-7410, March 17, 1991. 13Oleson, S., and Myers, R., "Launch Vehicle and Power Level Impacts on Electric GEO Insertion," 32nd

AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA 1996-2978, July, 1996. 14Manzella, D., Jacobson, D., and Jankovsky, R., "High Voltage SPT Performance,", 37th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference, AIAA 2001-3774, July 2001. 15Hofer, R., and Gallimore, A., "Efficiency Analysis of a High-Specific Impulse Hall Thruster," 40th

AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA 2004-3602, July, 2004. 16Sengupta, A., Brophy, J., Anderson, J., Garner, C., Banks, B., and de Groh, K., "An Overview of the Results from the 

30,000 Hr Life Test of Deep Space 1 Flight Spare Ion Engine," 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 

AIAA 2004-3608, July 2004. 17de Grys, K., Fisher, J., Wilson, F., Beal, B., Dimicco, J., and Khayms, V., "4.5 KW Hall Thruster System Qualification 

Status," AIAA 2004-3603, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July, 2004. 18Polk, J., Kakuda, R, Anderson, J., Brophy, J., Rawlin, V., Patterson, M., Sovey, J., and Hamley, J., "Performance of the 

NSTAR Ion Propulsion System on the Deep Space One Mission," AIAA 2001-96, Jan., 2001. 19Advanced Projects Design Team (Team X), Titan Orbiter 2003-10, Report ID #658, October 7,9,10, 2003. 20http://elvperf.ksc.nasa.gov/elvMap/index.html 21Wertz, J. R., and Larson, W. J., Space Mission Analysis and Design, 3rd ed., Microcosm Press, El Segundo, CA., and 

Kluwer Academic Publishers, Dordrecht, The Netherlands, 1999, pg. 702. 22Neptune Orbiter Mission Analysis Assumptions, In-Space Integrated Space Transportation Activity by the 2004 IISTP 

Systems Analysis Team, March 2004, (contact Larry Kos, MSFC, 256-544-1522) 23Noca, M. A., and Bailey, R. W., “Mission Trades for Aerocapture at Neptune,” 40th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference, AIAA 2004-3843, July 11-14, 2004. 24Oh, David Y., “Evaluation of Solar Electric Propulsion for Discovery Class Missions,” 41th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference, AIAA 2005-4270, July 11-14, 2005.

NASA/TM—2006-214044 21
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AIAA–2005–4268

E–15401

WBS–22–800–90–09

27

Performance of Solar Electric Powered Deep Space Missions Using Hall

Thruster Propulsion

Kevin E. Witzberger and David Manzella

Electric propulsion; Hall thrusters; Ion thrusters

Unclassified -Unlimited

Subject Categories: 12, 13, and 20

Prepared for the 41st Joint Propulsion Conference and Exhibit cosponsored by AIAA, ASME, SAE, and ASEE,

Tucson, Arizona, July 10–13, 2005. Responsible person, Kevin E. Witzberger, organization code PBM, 216–433–3463.

Power limited, low-thrust trajectories were assessed for missions to Jupiter, Saturn, and Neptune utilizing a single Venus

Gravity Assist (VGA) and a primary propulsion system based on either a 3-kW high voltage Hall thruster, of the type

being developed by the NASA In-Space Propulsion Technology Program, or an 8-kW variant of this thruster. These Hall

thrusters operate with specific impulses below 3,000 seconds. A trade study was conducted to examine mission parameters that include: net delivered mass (NDM), beginning-of-life (BOL) solar array power, heliocentric transfer time,

required launch vehicle, number of operating thrusters, and throttle profile. The top performing spacecraft configuration

was defined to be the one that delivered the highest mass for a range of transfer times. In order to evaluate the potential

future benefit of using next generation Hall thrusters as the primary propulsion system, comparisons were made with the

advanced state-of-the-art (ASOA), 7-kW, 4,100 second NASA's Evolutionary Xenon Thruster (NEXT) for the same

mission scenarios. For the BOL array powers considered in this study (less than 30 kW), the results show that the

performance of the Hall thrusters, relative to NEXT, is largely dependant on the performance capability of the launch

vehicle, and that at least a 10 percent performance gain, equating to at least an additional 200 kg dry mass at each target

planet, is achieved over the higher specific impulse NEXT when launched on an Atlas 551.