NASA/TM--2002-211892 AIAA-2002-3969

Mission Advantages of NEXT: NASA's

Evolutionary Xenon Thruster

Steven Oleson, Leon Gefert, Scott Benson, and Michael Patterson

Glenn Research Center, Cleveland, Ohio

Muriel Noca and Jon Sims

Jet Propulsion Laboratory, Pasadena, California

September 2002
The NASA STI Program Office... in Profile

Since its founding, NASA has been dedicated to

the advancement of aeronautics and space

science. The NASA Scientific and Technical

Information (STI) Program Office plays a key part

in helping NASA maintain this important role.

The NASA STI Program Office is operated by

Langley Research Center, the Lead Center for

NASA's scientific and technical information. The

NASA STI Program Office provides access to the

NASA STI Database, the largest collection of

aeronautical and space science STI in the world.

The Program Office is also NASA's institutional

mechanism for disseminating the results of its

research and development activities. These results

are published by NASA in the NASA STI Report

Series, which includes the following report types:

TECHNICAL PUBLICATION. Reports of

completed research or a major significant

phase of research that present the results of

NASA programs and include extensive data

or theoretical analysis. Includes compilations

of significant scientific and technical data and

information deemed to be of continuing

reference value. NASA's counterpart of peerreviewed formal professional papers but

has less stringent limitations on manuscript

length and extent of graphic presentations.

TECHNICAL MEMORANDUM. Scientific

and technical findings that are preliminary or

of specialized interest, e.g., quick release

reports, working papers, and bibliographies

that contain minimal annotation. Does not

contain extensive analysis.

CONTRACTOR REPORT. Scientific and

technical findings by NASA-sponsored

contractors and grantees.

CONFERENCE PUBLICATION. Collected

papers from scientific and technical

conferences, symposia, seminars, or other

meetings sponsored or cosponsored by

NASA.

SPECIAL PUBLICATION. Scientific,

technical, or historical information from

NASA programs, projects, and missions,

often concerned with subjects having

substantial public interest.

TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific

and technical material pertinent to NASA's

mission.

Specialized services that complement the STI

Program Office's diverse offerings include

creating custom thesauri, building customized

databases, organizing and publishing research

results.., even providing videos.

For more information about the NASA STI

Program Office, see the following:

• Access the NASA STI Program Home Page

at http://www.sti.nasa.gov

• E-mail your question via the Internet to

help@sti.nasa.gov

• Fax your question to the NASA Access

Help Desk at 301-621-0134

• Telephone the NASA Access Help Desk at

301-621-0390

Write to:

NASA Access Help Desk

NASA Center for AeroSpace Information

7121 Standard Drive

Hanover, MD 21076
NASA/TM--2002-211892 AIAA-2002-3969

Mission Advantages of NEXT: NASA's

Evolutionary Xenon Thruster

Steven Oleson, Leon Gefert, Scott Benson, and Michael Patterson

Glenn Research Center, Cleveland, Ohio

Muriel Noca and Jon Sims

Jet Propulsion Laboratory, Pasadena, California

Prepared for the

38th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Indianapolis, Indiana, July 7-10, 2002

National Aeronautics and

Space Administration

Glenn Research Center

September 2002
NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Available from

National Technical Information Sel_ice

5285 Port Royal Road

Springfield, VA 22100

Available electronically at http //gltrs.grc.nasa.gov
MISSION ADVANTAGES OF NEXT: NASA'S EVOLUTIONARY XENON

THRUSTER

Steven Oleson, Leon Gefert, Scott Benson, and Michael Patterson

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio

Contact: _._:._)._.f_!_ _L_.!___._._!_:__!:2.!__'_

Muriel Noca and Jon Sims

Jet Propulsion Lab

California Institute of Technology

Pasadena, Califol7lia

With the demonstration of the NSTAR propulsion system on the Deep Space One mission, the range of the

Discovery class of NASA missions can now be expanded. NSTAR lacks, however, sufficient performance

for many of the more challenging Office of Space Science (OSS) missions. Recent studies have shown that

NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system is best choice for many exciting

potential OSS missions including outer planet exploration and inner solar system sample returns. The

NEXT system provides the higher power, higher specific impulse and higher throughput required by these

science missions.

Introduction

With the demonstration of the NSTAR

propulsion system on the Deep Space One

mission, the range of the NASA Discovery class

of missions can now be expanded. NSTAR

lacks, however, sufficient performance for many

of the more challenging Office of Space Science

(OSS) missions. Recent studies have shown that

NASA's Evolutionary, Xenon Thruster (NEXT)

ion propulsion system is best choice for many

exciting potential OSS missions including outer

planet exploration and inner solar system sample

returns.

The NEXT system provides quicker missions,

delivers heavier payloads, and reduces

complexity compared to NSTAR and other

options while exhibiting redundancy, de-rating,

and optimal Iw. The key improvements over

NSTAR, embodied in the NEXT system, are

higher power and lower specific mass, and to a

lesser extent higher I_p. Specific potential OSS

missions where the NEXT system is applicable

are outer planet missions: Neptune Orbiter, Titan

Explorer. Europa lander, Europa Subsurface,

Saturn Ring Observer: and inner solar system

missions: Venus Sample Return, Comet Nucleus

Sample Return, and Mars Sample Return.

The NEXT ion propulsion system was recently

chosen to be developed by the Next Generation

Ion (NGI) NASA Research Announcement

(NRA) competition. The NEXT propulsion

system will be developed over the next few years

to a technology level ready to be injected into a

flight development program. Mission analyses

similar to those shown in this paper are

undergoing to support the development of the

final design to be developed. Due to these

further mission analyses the final NEXT ion

propulsion system design may be different than

the proposed design discussed in this paper.

An outline of this work is as follows. First the

proposed NEXT ion propulsion is described and

modeled, followed by the modeling of the solar

electric propulsion (SEP) spacecraft. The

mission analysis methodology and assumptions

are then explained. After a discussion of past

specific mission analyses, the design reference

missions included in the NGI NRA and their

assumptions are described. Finally. mission

performance for some other SEP missions using

the NEXT ion propulsion system are briefly

reviewed.

NASA/TM--2002-211892 1
System Analyses

Propulsion Systems

A team of analysts, technologists and engineers

from Glenn Research Center (GRC), the Jet

Propulsion Lab (JPL), Boeing Electron Dynamic

Devices (BEDD), and General Dynamics (GD -

formerly Primex Aerospace) developed the

propulsion system analysis models of the NEXT

propulsion system for the mission trades.

!

Figure 1. NASA's Evolutionary Xenon

Thruster

The NEXT ion propulsion system is assumed to

consist of several main components including the

thruster (figure 1), the gimbal, the power

processing unit (PPU), the digital control

interface unit (DCIU), and the feed system (split

into a main component and a per engine

component), t The estimated system masses are

shown in table 1. The propellant tank is assessed

with the spacecraft. Structure and thermal

support of the propulsion system subsystems is

handled by the spacecraft and modeled with the

rest of the subsystems.

Table 1. NEXT Propulsion System Masses

NEXT Propulsion System Masses for Mass

DSDRM .b (kg)

each

Ion thrqster 12

Thruster gimbai +actuators 2.6

Power Processing Unit 27 4

PPU to Thruster Cabling _ 26

DCIU 2,5

Feed system fixed 0,23

Feed System per thruster ° I. 13

a Structure. Thermal Contmgerlcy

ir_cludet with spacecraft

b Tanks not included

c 2 m. cross-strapped thermal handled

by spacecraft

d IncMdes tubing and fittings

Since the NEXT system is being developed for

solar powered interplanetary spacecraft where

power level varies with distance from the Sun

some amount of power throttling will be required

of the system. The NEXT system will be

designed for operation with input powers from 1

kW up to perhaps as high as 10 kW. The

throttle table used in the analysis is shown in

Table 2.

With proper PPU design the NEXT thruster

should be capable of a limited range of Isr

throttling at each power point. The limits of this

throttling are noted as High Isp and High Thrust

in figure 2. Mission analyses traded these

limiting throttle curves to determine which was

most appropriate for each mission. Current

analysis techniques are unable to allow throttling

throughout the range at each power level.

However, new techniques have recently been

developed and are currently being used to

support the development of the thruster.

Table 2. NEXT Throttle Table

NEXT Nominal THROTTLE TABLE

Power into Thrust (raN)

PPU (W)

Total Specific

Propulsion Impulse (s)

System

Efficiency

(thruster &

PPU)

10.5 364 0.67 3940

8.60 299 0.67 3900

6 38 221 0.66 3870

4,91 170 0 66 3860

2 28 76 0.61 3760

7.75 282 066 3680

73 209 065 3660

.42 160 0 65 3630

207 7I 0.60 3550

4.97 192 064 3360

382 147 0.63 3340

1.81 65 058 3250

3.37 137 0,62 3110

1.61 61 056 3020

1.17 49 0 51 2460

NASA/TM--2002-211892
Due to the high specificimpulseof ion

propulsionsystems,longburntimes(years)are

neededto provide the required mission energy

(velocity change or V). This challenge to

create a long-lived propulsion system was a

major part of the NSTAR development

(including a ground life test) and will be integral

to the development of the NEXT thruster.

Analyses have shown that throughput lifetimes

for the NEXT thruster should be at least 300 kg,

if not more. _ The mission analyses adjusted the

number of thrusters to handle the required

throughput.

To demonstrate the need and advantages of the

NEXT system the missions were also evaluated

using the state-of-art 2.3 kW ion thruster

NSTAR (flown on Deep Space 1), as well as a

high I_v 30 cm ion propulsion system analyzed in

the Integrated In-Space Transportation Planning

(IISTP) studies (hereafter denoted as the IISTP5kW-30 cm). Mass. throttle and life information

on the NSTAR and the IISTP-5kW-30cm was

obtained from the NRA and the IISTP analysis,

respectively. 2.3

4000 _sp vs Input Power for 40 cm NEXT,_,

/ iligh ................... _!o!_

/ ! / _"

2800 1

2400

0 3000 6000 9000 ] zooc

Pin(Into PPU)(180

Figure 2. High lsp and High Thrust Throttle

Fits

Solar Electric Propulsion Spacecraft

The rest of the SEP spacecraft, such as the solar

array, thermal system, and structure, were

modeled using a spacecraft model developed

during the JPL integrated project design center

(Team X) analysis of the Titan Explorer mission

for the IISTP. 3.4.5.6 Glenn Research Center and

Marshall Space Flight Center supported this

exercise with technology analysis and mission

support. The propulsion subsystem is integrated

into the spacecraft using this model.

To be consistent with the JPL conceptual design

guidelines used in the IISTP analysis, 30% mass

and 30% power contingencies were applied to all

spacecraft subsystems, and a 10% launch vehicle

margin was assumed. The structures/cabling

masses are not based on a specific design but are

a percentage of the subsystems to which the

structures apply (tTpically 26% of the propulsion

system and 16% of the power system for

structures). These percentages are based on

historical data and are consistent with the design

guidelines of Team X. However. the structural

mass includes additional mass to carry the

spacecraft above the SEP stage to overcome

launch loads, mass to hold the tanks, mass for

the SEP stage separation mechanism and

additional mass for the system assembly

hardware (bolts. epoxies, tie downs...). The

payload spacecraft side of the launch vehicle

adapter is also included in the SEP stage

structure.

One spare ion engine for every four operating

ion engines and one spare PPU and digital

control interface unit (DCIU) are included for

single-fault tolerance. Each thruster was

gimbaled separately.

The tankage fraction was calculated assuming

cylindrical composite tanks. Those tanks have a

propellant storage efficiency (Tank Fraction TF)

of about 2.5% for Xenon when stored as a

supercritical gas (-2000 psia). Furthermore, a

10% propellant contingency was added to the

deterministic propellant mass to account for

residuals, attitude control and margin. Other

assumptions are as follows:

SEP Stage Assumptions

Minimum two thrusters operating (actual number

optimized [but 2 or greater] by SEPTOP codeJ

Spare Thruster and PPU Added

Delta IV M+ (4,2) Launcher

Common, 100VDC Solar Array

5% Fuel added and expended during flight for errors

and misc.

5% Fuel reserve added

10% Trajectory Management Coast Time

174 W/kg Ultraflex Solar Arrays (150 W/kg with

structure )

Power System Specific Mass: 9.2 kgJ'kW

Power and Thermal Structure: 16%

Thermal control per PPU : 24 kg/kWthermal

Stage Cabling (Not PPU) : 6% of Power Mass

Tankage Mass Fraction (includes Structure[26%

dry,4%fuel] and 5% residual fuel): 15.6%

Structure to Support Payload Spacecraft on SEP Stage:

8 5% of Payload mass

Spacecraft Adapter: 42 kg

Launch Vehicle Adapter: 65 kg

NASA/TM--2002-211892 3
Alternate Ion propulsion systems

c, NSTAR SOA: Masses and Performance

from DSDRM (reference 2) except

advanced tanks used with NEXT

c_ IISTP5 kW 30 cm: Mass and Performance

Data from Reference 3. (I1STP Analyses)

and advanced Tanks used with NEXT

Mission Analyses

Once the propulsion system has been modeled in

terms of mass and performance and incorporated

into a spacecraft model its delivered payload

mass and required trip time can be determined

using a trajectory optimization code. The code

of choice for this analysis was the Solar Electric

Propulsion Trajectory Optimization Program

(SEPTOP) developed by Carl Sauer of JPL.

SEPTOP is a calculus-of-variations code which

optimizes two body interplanetary trajectories.

SEPTOP models discrete numbers of operating

xenon thrusters throughout a trajectory. The

number of operating thrusters is switched by

SEPTOP in an optimal fashion. Additionally.

SEPTOP throttles the thrusters in power as

required by that available from the Sun.

SEPTOP was used to evaluate trades of

delivered payload and trip time.

In order to examine trades of payload mass

payload versus trip time. the SEPTOP code

needs a simplified system mass model. A top

level mass equation was developed which was

derived from the Team X modeling for input into

SEPSPOT:

M_,,nch = MAdapter + (7.PowerP + MsEP_Fixe d +

Mp_ylo_d (1 + ksn,ct,,reMpayZo_d ) +

MP, om,_,,, (1 + kT,,,_g e Mpayload )

where.

Mt .... h = Total Launch Mass (kg)

MAdapter = LV Adapter Mass (kg)

Mvaylo_d= Science Payload Mass (kg)

Mvropdl_t = Xenon Propellant (kg)

¢tpo_._ = Specific Power Mass (kg/kW)

[determined by component mass summation]

kstructure = Payload Structure Fraction (%)

k._ = Propellant Storage Fraction (%)

MSEP-F,_d= Fixed Mass (kg)

P = Array 1 A.U. Power (kW)

The values of kstructure, kt_k_g_. P are set to 8.5 %.

15.6 % and 25 kW (from DSDRM assumptions)

respectively. The linear scaling factors k_t_c_r_.

kt_k_s_ and Ctpo,,ermodel the rate in which their

systems functionally grow and MsEP-Fixed includes

the fixed masses of systems that do not grow as a

function of other parameters.

Mission Trades

IISTP Studies

Previous to the release of the NGI NRA. the

Integrated In-Space Transportation Planning

(IISTP) studies compared many potential

technologies for various NASA. government and

commercial missions/TM These studies

indicated that a high power ion propulsion

system (of which NEXT was one of the samples)

is the most important technology for

development due to its outstanding performance

versus perceived development and recurring

costs for interplanetary solar electric propulsion

missions. One of the best applications of a high

power electric propulsion system was as an

integral part of a Solar Electric Propulsion (SEP)

stage to send a payload to outer planet targets

which would use aeroshells to capture into the

outer planetary orbits. 4 The IISTP studies

showed that either trip time or launch vehicle

class can be significantly reduced when

compared to state-of-the-art systems. 4__

Deep Space Design Reference Mission Studies

The SEP stage mission developed for the IISTP

studies is very similar to the Deep Space Design

Reference Missions (DSDRMs) defined in the

NRA. The main difference was that the

DSDRMs had greater desired payloads. The

role of the DSDRMs was to allow proposers to

demonstrate their concept's ability to perform

missions of interest. 2 Thus to show the

advantages of the NEXT ion thruster for the

DSDRM missions. JPL and Glenn Research

Center analysts used the SEPTOP numerical

optimization code (previously used for the IISTP

analyses) to evaluate the NEXT's performance

for two outer planet missions: Titan Explorer and

Neptune Orbiter. These two missions represent

targets at Saturn or closer, and beyond Saturn.

respectively, as desired in the DSDRM. In order

to quantify the mission benefits of the NEXT ion

propulsion system analyses were run to

determine the mission performance using SOA

NSTAR systems. An additional approach was

also analyzed (similar to that in the IISTP

analyses), that of a 30 cm thruster at 5 kW. 3.4.5._

NASA/TM--2002-211892 4
Table 3. Summary of DSDRM Inputs for

SEPTOP

DSDRM

Spacecraft

Assumptio

ns

[SEPTOP

Inputs]

Target

# Thrusters

/# PPUs

Fixed mass

[kg] (Dry.

SEP Stage

less power

& Tanks I

[alfa(2)]

Payload

[alfa(7)]

Payload

Structure

Fraction

[alfa(8)]

Tankage

(include

structure &

5%

Residual )

[kt]

Power

System

Specific

Mass

(k_W)

[alfa( 1)]

× << _

Z

Neptune

4/4 9/9

601 696

85O

0085

6/6

570

.156 .28 .15

9.2

Saturn

3/3 9/9

501 696

1400

0.085

5 ,/5

510

.15 28 .15

9,2

The DSDRM clearly defined the power and

payload spacecraft to be assumed for each

mission. A total solar array power of 30 kW is

provided at Earth's orbital distance from the Sun

but only 25 kW maximum power was to be

available to the propulsion system, allowing 5

kW for housekeeping and reserve. As the

spacecraft journeys closer to the Sun to flyby

Venus the arrays are assumed to be manipulated

to cap the power to the propulsion system at 25

kW. The payload orbiter spacecraft with its

aerocapture system was assumed to be 850 kg

and 1400 kg for the Neptune (past Saturn

example) and Titan Orbiter (Saturn and closer

example), respectively. After the specific SEP

stage parameters were calculated by the design

model they were interpreted into those required

by the SEPTOP model as described earlier.

These parameters are shown in table 3. Each of

the DSDRM missions and the analysis results

will now be described.

Neptune Orbiter

Placing an orbiter around Neptune is of great

interest after the tantalizing views sent back by

Neptlllle

Figure 3. Neptune Orbiter Trajectory

Voyager. The SEP system's role for the mission

is that of a stage to build up energy in the inner

solar system, redirect the energy with a Venus

flyby and separate the radioisotope powered

Neptune orbiter on a fast trajecto_ _ to Neptune.

Once at Neptune an advanced aeroshell system is

used to capture in orbit.

Earth

Figure 4. SEP Portion of Neptune Orbiter

Trajectory

A sample trajectory for the SEP stage is shown

in figure 3. The SEP stage provides a veD' fast

path to Neptune after the Venus flyby. A closer

view of the inner solar system trajectory showing

the Venus flyby is shown in figure 4. Since the

available power varies as the spacecraft moves

closer to and further from the sun before and

after the Venus flyby, the number of thrusters

which may be operated varies. This variation of

power and operating thrusters is shown in figure

5. The number of operating thrusters in this

case varies 4-3-2-3-4-3-2 thrusters as power

varies with solar distance.

NASA/TM--2002-211892 5
60 -

50-

_ 40 ira30

20 _

I0

0

!

# Operating ] ! Thrusters

0 200 400 600 800

Trip Time (da.vs)

i:

2 .=

,I

o

Figure 5. Power and Operating Thruster

Variation with Thrust Time

Maximum power into the propulsion system was

set by the DSDRM at 25 kW. Note however,

when the spacecraft approaches Venus the

potential power level exceeds this power. A

small study was performed to see if this

additional power (up to 40 kW) would be

beneficial to the mission, but the additional

power management mass to deliver the power to

the thrusters offset any, benefits. This result was

reasonable due to the relatively short time the

power is available. When the power falls to

around 2 kW at the end of the inner solar system

portion of the trajectory the propulsion system is

shut down since the minimum throttled thruster

power is 1 kW.

lloo

1050

looO

Z a_

BOO

7_

7OO

Neptune Performance

NEXT,

/f

j op_. I! _xT.

| f_ _ _T4f_usters High

5 10 15 20

Trip Time (years)

Figure 6. Performance Trade of High Isp and

High Thrust NEXT Options

Before comparing the NEXT propulsion system

performance to that of the other ion systems a

comparison was made of the performance of the

High Isp and High Thrust throttle options shown

in figure 1. Results are shown in figure 6.

While both systems delivered the required

payload in roughly, the same time. the high thrust

option was slower for higher payload masses.

Since most science payloads tend to grow in

mass the high Iw option was chosen as superior

and used for the rest of the analyses.

Results shown in figure 7 for the Neptune

Orbiter mission clearly show the trip time

advantages of the NEXT system versus the other

systems. Such savings in trip time can save >$5

million a year in operations costs for missions in

addition to allowing scientific investigators more

timely' research. The use of fewer thrusters

should also result in a cost savings as well as a

reduced complexity. The SOA NSTAR system is

hampered by its low power capability,

necessitating an exorbitant propulsion system

dr), mass. The IISTP-5kW-30 cm system has too

high an I_r and must rely upon the launch vehicle

to perform more of the mission (higher excess

velocib_ [C3]. less mass at escape) than for the

NEXT ion propulsion system case.

1100

1000

9oo

"_ 8OO

7OO

6OO

Neptune Performance

NEXT,

4 thrusters //IISTP- 30cm-SkW,

6 thrusters

t......//: ill

l

i r

10 15 20

Trip Time (years)

Figure 7. Neptune Performance Trade of Ion

Propulsion Options

The NEXT system performance results shown

are for a 4 thruster / 4 PPU case where the power

is capped at 25 kW as directed in the DSDRM.

For the desired 850 kg payload the 4 thruster

configuration's maximum power level and fuel

throughput (for each thruster/PPU) is decidedly

derated at 6.25 kW and just under 200 kg.

respectively. If a thruster or PPU fails the

mission can be easily completed with only three

NASA/TM--2002-211892
thrusters at a still derated performance of 8.3 kW

and 260 kg. This clearly demonstrates the

NEXT thrusters are not oversized for the mission

and consequently will have plenty of growth

potential for even higher power missions.

,Saturn Performance

NEXT,

1400 _k 1 / w.

]1200| 1 / 4 thrusters

1lot) NSTAR,

1000 / 9 thrusters

4 5 6 7 8

Trip Time (years)

Figure 8. Saturn (Titan) Performance Trade

Ion Propulsion Options

Titan Explorer

The other DSDRM mission analyzed used a

similar SEP Venus flyby trajectory to place an

aerocapture orbiter about Saturn's moon Titan.

The trajectories and throttling requirements for

Titan look similar to those for the Neptune

orbiter.

Applying the NEXT ion propulsion technology

to the Titan Explorer mission showed slightly

different results (figure 8.). The larger 1400 kg

desired payload began to stress the assumed

launch vehicle with the required performance

moving past the knee of the performance

(payload versus trip time) curve. Regardless. the

NEXT system is again a year quicker than the

IISTP-5KW-30 CM system and the SOA

NSTAR is unable to deliver the required

payload.

Comet Sample Return

Returning samples from a comet are also of great

interest. Use of SOA NSTAR for a Comet

Return mission would require too many thrusters

(9) due to the limited power capability, making

the delivered mass performance poor. A study

using the NEXT ion propulsion system was

made using the same models developed for the

DSDRM missions. Results showed that almost

25% greater payload could be carried. Use of

the NEXT would also greatly reduce the

complexity" of the 30 kW SEP spacecraft by

reducing the number of thrusters from 9 NSTAR

to 4 NEXT (both concepts include spares).

Mars Sample Return

The NEXT technoiog)' can also be applied to

other NASA missions to great effect. Returning

samples from solar system targets is of particular

interest to scientists. A recent stud)', conducted

by JSC, Spectrum Astro. and SA1C. explored the

use of SEP for a Mars sample return mission

(these trades used assumptions different than

those of the DSDRM missions, but are

representative). 7 The SEP system was traded for

various phases of the mission from just returning

the sample from low Mars orbit to performing all

of the interplanetary and near planet propulsion.

Using a Delta II class launch vehicle a l0 kW

SEP spacecraft equipped with two NEXT

thrusters returned the Mars sample in a total

transit time of 4 ?'ears (2 years Mars to Earth):

the lander and Mars ascent vehicle (to place the

sample into Mars orbit) would be launched by

another launch vehicle. Use of a 40 kW SEP

spacecraft along with six NEXT thrusters could

perform the entire mission, deliver'3' of orbiter

and lander and realm of the samples in 4.9 years

round trip using a Delta 4 Heax3' launch vehicle.

To perform this mission using chemical SOA

and aerocapture would require two Atlas V 551

medium class launch vehicles. In these cases

the NEXT proved useful in both the l0 kW and

40 kW SEP spacecraft cases, demonstrating the

wide range of application of the technology.

Commercial Geostationary

The NEXT is also advantageous to governmental

and commercial geosynchronous missions.

Currently, Boeing is using four 4.5 kW ion

thrusters for both orbit insertion and station

keeping of their 17 kW 702 spacecraft. As

power levels of the spacecraft grow the larger

ion propulsion systems such as NEXT will be

able to provide payload increases of at least 50%

for a few months trip times. 8 With sufficient

power (40 kW) it should be possible to provide a

launch vehicle step-down for trip times of less

than six months. ')

Conclusions

These trade studies, while not encompassing all

potential missions, clearly demonstrate the

advantages of the NEXT system for NASA outer

planet, sample return, and interplanetary

missions, as well as commercial earth orbit

missions. Compared to the SOA NSTAR the

NEXT provides a higher power, lighter, longer

NASA/TM--2002-211892 7
life system which can be used for a wide range

of SEP power levels. The NEXT can provide

quicker trip times for better science and cheaper

missions. Mission trades were made to evaluate

the best way to throttle the thruster and the use of

extra power at Venus was also considered.

References

1. M. Patterson. J. Foster, T. Haag, V. Rawlin

and G. Soulas, NEXT: NASA's Evolutionary

Xenon Thruster, AIAA-2002-3832, 2002 Joint

Propulsion Conference, Indianapolis, Indiana,

July 2002.

2. NASA Research Announcement for the Next

Generation Ion Engine Technology, NRA-01-

OSS-01, section A.9.2.

3. Noca, M.. Frisbee, R.. Johnson. L., Kos, L.,

Gefert. L.. Dudzinski, L., Evaluating Advanced

Propulsion Systems for the Titan Explorer

Mission, IEPC-01-175, 27 th IEPC, Oct. 2001,

Pasadena. CA.

4. Integrated In-Space Transportation Planning

(IlSTP), Summary of Process and Technology

Prioritization, May 30. 2001. (contact Larry Kos,

MSFC. 256 544-1522)

5. Neptune Orbiter Mission Package, In-Space

Integrated Space Transportation Activity, by the

I1STP Systems Analysis Team, Apr. 4th, 2001,

(contact Larry Kos. MSFC. 256 544-1522)

6. Titan Explorer Mission, In-Space Integrated

Space Transportation Activity, Team X/A

Results, Apr. 19th, 2001. (Muriel Noca)

7. SEP for Mars Sample Return. MPSET Status

Briefing. Dec. 5. 2001. (Contact Jeff George.

NASA/JSC)

8. Oleson, S.R., Myers, R.M., Launch Vehicle

and Power Level Impacts on Electric GEO

Insertion, 32nd Joint Propulsion Conference,

July, 1996, AIAA-96-2978, NASA Contractor

Report 198520.

9. Oleson, S.R.. Advanced Electric Propulsion

for RL V Launched Geosynchronous Spacecraft, ,

26th International Electric Propulsion

Conference, Kitak3'ushu, Japan, October. 1999,

IEPC 99-185.

NASA/TM--2002-211892

REPORT DOCUMENTATION PAGE FormApproved

OMB No. 0704-0188

Public reporting burden for this collectionof information is estimated to average 1 hour per response,includingthe time for reviewing instructions,searchingexistingdata sources,

gatheringand maintaining the data needed, and completing and reviewing the collection of information. Send commentsregarding this burden estimate or any other aspect of this

collectionof information,including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington,VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

September 2002 Technical Memorandum

4. TITLE AND SUBTITLE

Mission Advantages of NEXT: NASA's Evolutionary Xenon Thruster

6. AUTHOR(S)

Steven Oleson, Leon Gefert, Scott Benson. Michael Patterson, Muriel Noca, and

Jon Sims

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135 - 3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

5. FUNDING NUMBERS

WU-344-96-8D--O0

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-13574

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--2002-211892

AIAA-2002-3969

11. SUPPLEMENTARY NOTES

Prepared for the 38th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE, and ASEE,

Indianapolis, Indiana, July 7-10, 2002. Steven Oleson, Leon Gefert, Scott Benson, and Michael Patterson, NASA Glenn

Research Center; Muriel Noca and Jon Sims, Jet Propulsion Laboratory, California Institute of Technology, Pasadena,

California. Responsible person, Steven Oleson, organization code 6910, 216-977-7426.

12a. DISTRibUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Categories: 13, 20, and 91 Distribution: Nonstandard

Available electronically at http://gltr_. _rc.nasa. gOv

This publication is available from the NASA Center for AeroSpace Information, 301-621_)390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

With the demonstration of the NSTAR propulsion system on the Deep Space One mission, the range of the Discovery class

of NASA missions can now be expanded. NSTAR lacks, however, sufficient performance for many of the more challenging

Office of Space Science (OSS) missions. Recent studies have shown that NASA's Evolutionary Xenon Thruster (NEXT)

ion propulsion system is the best choice for many exciting potential OSS missions including outer planet exploration and

inner solar system sample returns. The NEXT system provides the higher power, higher specific impulse, and higher

throughput required by these science missions.

14. SUBJECT TERMS

Ion propulsion; Electric propulsion; Planetary spacecraft

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

]15. NUMBER OF PAGES

14

16. PRICE CODE

20. LIMITATION OF ABSTRACT

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-1B

298-102