James S. Sovey

Alphaport, Inc., Cleveland, Ohio

George C. Soulas and Daniel A. Herman

Glenn Research Center, Cleveland, Ohio

NEXT Propellant Management System Integration

With Multiple Ion Thrusters

NASA/TM—2011-217040

May 2011
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James S. Sovey

Alphaport, Inc., Cleveland, Ohio

George C. Soulas and Daniel A. Herman

Glenn Research Center, Cleveland, Ohio

NEXT Propellant Management System Integration

With Multiple Ion Thrusters

NASA/TM—2011-217040

May 2011

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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NASA/TM—2011-217040 1 

NEXT Propellant Management System Integration

With Multiple Ion Thrusters

James S. Sovey

Alphaport, Inc. 

Cleveland, Ohio 44135 

George C. Soulas and Daniel A. Herman

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Abstract

As a critical part of the NEXT test validation process, a multiple-string integration test was performed 

on the NEXT propellant management system and ion thrusters. The objectives of this test were to verify 

that the PMS is capable of providing stable flow control to multiple thrusters operating over the NEXT 

system throttling range and to demonstrate to potential users that the NEXT PMS is ready for transition to 

flight. A test plan was developed for the sub-system integration test for verification of PMS and thruster 

system performance and functionality requirements. Propellant management system calibrations were 

checked during the single and multi-thruster testing. The low pressure assembly total flow rates to the 

thruster(s) were within 1.4 percent of the calibrated support equipment flow rates. The inlet pressures to 

the main, cathode, and neutralizer ports of Thruster PM1R were measured as the PMS operated in

1-thruster, 2-thruster, and 3-thruster configurations. It was found that the inlet pressures to Thruster 

PM1R for 2-thruster and 3-thruster operation as well as single thruster operation with the PMS compare 

very favorably indicating that flow rates to Thruster PM1R were similar in all cases. Characterizations of 

discharge losses, accelerator grid current, and neutralizer performance were performed as more operating 

thrusters were added to the PMS. There were no variations in these parameters as thrusters were throttled 

and single and multiple thruster operations were conducted. The propellant management system power 

consumption was at a fixed voltage to the DCIU and a fixed thermal throttle temperature of 75 °C. The 

total power consumed by the PMS was 10.0, 17.9, and 25.2 W, respectively, for single, 2-thruster, and 

3-thruster operation with the PMS. These sub-system integration tests of the PMS, the DCIU Simulator, 

and multiple thrusters addressed, in part, the NEXT PMS and propulsion system performance and 

functionality requirements. 

Nomenclature

DCA discharge cathode assembly

DCIU digital control interface unit 

FCD flow control device

GRC Glenn Research Center

HPA high pressure assembly

LPA low pressure assembly

MFC mass flow controller

NCA neutralizer cathode assembly

NEXT NASA’s Evolutionary Xenon Thruster 

NSTAR NASA’s Solar Electric Propulsion Technology Application Readiness 

PFCV proportional flow control valve 
NASA/TM—2011-217040 2 

PM prototype model 

PMS propellant management system

RTD resistance temperature detector

SSIT single string integration test

VF vacuum facility

XFSE xenon feed support equipment

Introduction

The NASA Glenn Research Center (GRC) is responsible for the development of NASA’s Evolutionary 

Xenon Thruster (NEXT) ion propulsion system (Refs. 1 and 2). This system is a next generation ion 

propulsion system to follow the successful NASA’s Solar Electric Propulsion Technology Applications 

Readiness (NSTAR) ion propulsion system that propelled NASA’s Deep Space 1 spacecraft and is presently

propelling the Dawn spacecraft (Refs. 3 and 4) Propulsion system elements under development by the 

NEXT program include a high-performance, 7 kW ion thruster; a modular, high-efficiency 7 kW power 

processor unit; a highly flexible advanced xenon propellant management system; and a compact, lightweight thruster gimbal. This design approach was selected to provide future NASA science missions with 

the greatest value in mission performance benefit at a low total development cost (Ref. 1).

As a critical element of the NEXT test validation process, a multiple string integration test of the 

propellant management system (PMS) and three operational thrusters was performed. The objectives of 

this test were to verify that the PMS is capable of providing stable flow control to single or multiple 

thrusters operating over the system throttling range and to demonstrate to potential users that the NEXT 

PMS is ready for transition to flight. Propulsion system elements included in this integration test were:

• an advanced engineering model ion thruster, labeled PM1R, that was manufactured by the NEXT 

program’s industrial partner Aerojet and successfully completed environmental testing at 

qualification levels (Refs. 5 and 6); 

• two engineering model ion thrusters (EM1, EM5) that were previously employed in the NEXT 

Multi-Thruster Engineering Demonstration Test (Ref. 7); 

• a propellant management system that was designed and manufactured by the NEXT program’s 

industrial partner Aerojet and successfully completed environmental testing at qualification levels

(Refs. 8 and 9); 

• three power consoles that were previously employed in the NEXT Multi-Thruster Engineering 

Demonstration Test (Refs. 7 and 10); 

• a breadboard model digital control interface unit (DCIU) that acted as a test console for

performing ground-based testing of the propulsion system, and was designed and manufactured 

by the NEXT program’s industrial partner Aerojet (Ref. 9). 

This paper will present test results of the NEXT PMS multiple string integration test using the 

propulsion system elements described in this section. 

Propulsion System Element Descriptions

The following sections describe the propulsion system elements tested during the PMS integration test. 

They include the ion thruster, the propellant management system, and the DCIU simulator. Also included 

in a separate section is a description of the propulsion system interfaces.
NASA/TM—2011-217040 3 

Ion Thrusters

The advanced engineering model thruster (previously labeled prototype model, or PM) used in the 

integration test is labeled PM1R, and is shown in Figure 1 along with Thrusters EM1 and EM5. The thruster 

was developed by NASA GRC, and the technology was transferred to Aerojet, who designed and built this 

flight-like thruster. The PM thruster design is functionally identical to the thruster developed by NASA

(Ref. 11). The PM design improved upon the GRC thruster design with emphasis on surviving vibration and 

thermal environments and on reduced thruster mass. Manufacturability was also improved with this new 

design. The PM thruster design included innovative coatings to increase emissivity for enhanced thermal 

margin, more uniform ion optics aperture diameters with much shallower cusps, a 36 cm beam extraction 

diameter to reduce edge aperture erosion, and graphite discharge cathode keeper to mitigate keeper erosion. 

A more detailed discussion of the PM thruster design can be found in Reference 5.

The PM1R thruster is a reworked version of the PM1 thruster that was initially performance tested in 

the summer of 2006 (Ref. 12).The PM1R thruster successfully completed environmental testing and 

single-string testing. Detailed results can be found in References 6, 13, and 14. 

Also shown in Figure 1 are NEXT engineering model thrusters, EM1 and EM5. These thrusters were 

used in the multi-thruster array demonstration test (Ref. 7) and were also employed in this PMS 

integration test. General thruster configurations are shown in Table 1. Thrusters EM1 and EM5 are 

similar to the EM thruster tested for 2000 hr in 2003 (Ref. 15). The EM thrusters utilize a 40 cm beam 

extraction diameter. These thrusters use a hollow cathode electron emitter and a semi-conic chamber with 

a ring-cusp magnetic circuit employing high strength rare earth magnets. Compact high voltage propellant 

isolators with voltage isolation capability greater than 1800 V are also used on the EM thrusters. The 

neutralizer cathode assembly is mechanically similar to the Hollow Cathode Assembly used on the 

International Space Station Plasma Contactor (Ref. 16). 

An abbreviated version of the NEXT thruster throttle table is shown in Table 2. Predicted

performance for several throttle points used in this investigation is shown for power levels ranging from 

0.55 to 6.86 kW. 

TABLE 1.—CONFIGURATION OF THE THRUSTERS USED WITH THE NEXT PMS

Component/Assembly PM1R EM1 EM5

DCA Pre-operated in single 

string integration test 

Pre-operated in multithruster array testPre-operated in multithruster array test

NCA Pre-operated in single 

string integration test 

Pre-operated in multithruster array testPre-operated in multithruster array test

Ion optics (Beam dia.) Pre-operated in single 

string integration test 

(36 cm)

Pre-operated in multithruster array test

(40 cm)

Pre-operated in multithruster array test

(40 cm)

Discharge chamber Pre-operated in single 

string integration test 

Pre-operated in multithruster array testPre-operated in multithruster array test
NASA/TM—2011-217040 4 

TABLE 2.—THRUSTER THROTTLE TABLE WITH PREDICTED THRUST AND

SPECIFIC IMPULSE FOR RELEVANT THROTTLE POINTS

[All neutralizer keeper currents are 3.0 A.]

Nominal 

thruster 

power, 

kW

Beam 

current,

A 

Beam 

supply 

voltage,

V

Accel. 

voltage,

V 

Main

flow rate,

sccm

Cathode 

flow rate, 

sccm

Neut. 

flow rate, 

sccm

Thrust, 

mN

Specific 

impulse,

s 

6.86 3.52 1800 –210 49.6 4.87 4.01 236 4190

4.74 3.52 1179 –200 49.6 4.87 4.01 192 3390

3.66 2.70 1179 –200 37.6 4.26 3.50 147 3360

2.78 2.00 1179 –200 25.8 3.87 2.50 108 3490

1.12 1.20 679 –115 14.2 3.57 3.00 49.2 2450

0.55 1.00 275 –350 12.3 3.52 3.00 25.6 1410

Propellant Management System 

The propellant management system used for the system integration test was an engineering model 

system. The overall design approach was developed by a NEXT integrated product team led by the NEXT 

program’s industrial partner Aerojet, who designed and manufactured the engineering model hardware. A 

schematic of the PMS single string is shown in Figure 2. The PMS single string is composed of High 

Pressure Assembly (HPA) and Low Pressure Assembly (LPA). The HPA reduces xenon tank pressure up 

to a maximum expected operating inlet pressure of 18,600 kPa (2700 psia) to a regulated outlet pressure 

of 240 kPa (35 psia). The outlet pressure is regulated with a proportional flow control valve (PFCV) using 
NASA/TM—2011-217040 5 

an outlet pressure transducer for feedback. The HPA includes a redundant PFCV and outlet pressure 

transducer for fault tolerance. A single HPA can provide flow to multiple LPAs for systems utilizing 

multiple thrusters.

The LPA provides independent flow control to each of the three thruster propellant inputs, labeled 

neutralizer, cathode, and main in Figure 2. During normal operation, each LPA branch flow rate is set by 

regulating the pressure to a heated porous plug, or thermal throttle, with a separate PFCV and pressure 

transducer. Thermal throttle temperature is controlled using a sheathed heater with a temperature sensor 

for feedback. As with the HPA, thermal throttle inlet pressure is regulated with a PFCV using a pressure 

transducer for feedback. The thermal throttle temperature is typically set to 75 °C and thermal throttle 

inlet pressures range from 77.9 to 189 kPa (11.3 to 27.4 psia) to achieve the commanded flow rates. To 

support fault tolerance in the design, the LPA design includes latch valves between the three branches of 

the LPA and the thermal throttle was designed for operation up to 400 °C. If a branch’s PFCV were to fail 

closed or pressure transducer not operate, for example, the failed branch’s latch valve would be opened. 

The working PFCV of the coupled branches would then regulate pressure to both branches to a constant 

level, and xenon flow rate would be varied by changing thermal throttle temperatures. The thermal 

throttle design also incorporated a redundant heater and temperature sensor for fault tolerance.

Photographs of the NEXT engineering model HPA and LPA are shown in Figures 3 and 4, 

respectively. The tested HPA’s redundant PFCV and outlet pressure transducer for this work were mass 

model mock-ups to reduce assembly cost. The LPA main flow branch can output a xenon flow rate of up 

to 50 sccm, while the cathode and neutralizer branches can output up to 6 sccm each. The HPA and LPA 

weigh 1.9 and 3.1 kg, respectively. These masses include each assembly’s mounting plate, connectors, 

and component heaters used for the system integration test. 

The engineering model HPA and LPA have completed environmental testing at qualification levels. 

Vibration testing was conducted in April of 2005 and included three random vibration tests to 14.1 grms 

in each axis for 2 min each. Sinusoidal vibrations sweeps were conducted before and after each test to 

assess PMS health. Thermal vacuum testing was conducted in May of 2007 and included three full 

thermal cycles from 12 to 70 °C with 2 hr dwell times at each temperature, followed by two separate

24 hr dwell times at each temperature. Flow calibration checks were conducted prior to and following the 

vibration test and throughout the thermal vacuum test. Proof pressure tests, leakage checks, and electrical 

checks were conducted on the HPA and LPA prior to and following the vibration and thermal vacuum 

tests. PMS, HPAs, and LPAs successfully completed environmental testing, and the results of these tests 

can be found in References 17 and 18. 

For this test with multiple ion thrusters, the PMS is comprised of one HPA and three LPAs that 

deliver regulated xenon flow to each of the three thrusters. Each thruster has three propellant lines: one 

for the discharge chamber plenum, one for the discharge cathode, and one for the neutralizer cathode. The 

LPA that is integrated with Thruster PM1R is constructed using flight-rated parts. To decrease development costs, the two LPAs integrated with Thrusters EM1 and EM5 were assembled using several nonflight components that are representative of flight hardware. The non-flight components included the latch 

valves, service valves, and pressure transducers. The PMS integration tests were conducted with the

 
NASA/TM—2011-217040 6 

HPA operated over an inlet pressure range from 329 kPa (47.7 psia) to 349 kPa (50.6 psia). In all cases the 

HPA outlet pressure was 242 kPa (35.1 psia). A PMS test with Thruster PM1R was conducted with HPA 

inlet pressures as high as 6,200 kPa (900 psia) during the NEXT Single String Integration Test (Ref. 14).

All tests were performed with the LPAs operating in the standard mode. That is to say the LPAs did 

not simulate failed mode operation where the LPAs could operate with either a failed open PFCV or a 

failed closed PFCV. Such fault handling was demonstrated during the NEXT Single String Integration 

Test (Ref. 14).

DCIU Simulator

The DCIU used for the integration test was a simulator that was built by NEXT’s industrial partner 

Aerojet. The DCIU Simulator was a data acquisition and control system used to demonstrate the 

performance objectives of the propulsion system elements (Ref. 9). In contrast to the other propulsion 

system elements, the objective of the DCIU Simulator was to act as a test console for performing groundbased testing of the propulsion system (Ref. 9). As a result, the DCIU Simulator was made of 

commercially available industrial control hardware. The only exception was the PFCV driver circuit 

which regulates pressure with a PFCV using pressure transducer for feedback. The PFCV driver circuit 

was designed to a brassboard level for this test, while the remainder of the DCIU Simulator was designed 

to a breadboard level (Ref. 19). 

The DCIU Simulator controlled and read telemetry from the PMS. The DCIU Simulator also operated 

the Xenon Feed System Equipment (XFSE) built to support single string integration testing of the PMS 

(Ref. 18).

The DCIU Simulator utilized two computers to improve response times. One computer handled 

hardware communications and stored test data while the other computer provided a graphical user interface. 

Communications between the computers and the DCIU Simulator hardware were handled by an Ethernet 

hub. The DCIU Simulator software was written in Visual C#. The software was not written for autonomous 

thruster operation. So thruster functions that included cathode conditioning, ignition, throttling, beam 

current regulation, and shutdown had to be conducted manually by the user with the software’s graphical 

interface. A schematic of the DCIU Simulator and the PMS interfaces is shown in Figure 5.
NASA/TM—2011-217040 7 

For the PMS interface, the DCIU Simulator had to power the HPA and LPAs, as well as provide 

analog commands and digitize the incoming analog telemetry. As a result, the DCIU Simulator included 

separate HPA and LPA interfaces that:

• provided an adjustable 24 to 32 Vdc housekeeping power to operate the pressure transducers and 

latch valves; 

• utilized industrial hardware for digital-to-analog conversion for analog commands and analog-todigital conversion for processing analog telemetry;

• operated the PFCV driver circuits for pressure regulation using analog commands; and

• operated the thermal throttles using commercial DC power supplies and software-controlled 

proportional-integral-derivative loops for temperature control.

The DCIU Simulator also included an XFSE interface box that: 

• operated separate pressure transducers to monitor PMS operation;

• operated separate mass flow controllers to independently monitor total PMS flow rates and to 

provide a separate flow path for thruster operation by actuating separate solenoid vales; and

• interfaced to a facility interlock to terminate propulsion system operation from a high facility 

pressure.

A brief description of the DCIU Simulator hardware can also be found in Reference 19. 

Summary of Propulsion System Interfaces 

A 22 to 34 Vdc low power bus is used by the DCIU for DCIU and PMS operation. Laboratory-class 

power consoles were used to power Thrusters PM1R, EM1, and EM5. 

The DCIU Simulator provides power to the PMS to operate the pressure transducers, latch valves, 

PFCVs, and thermal throttle heaters. For flow control, the DCIU regulates LPA and HPA pressures with 

the PFCVs and regulates thermal throttle temperature with the throttle heaters to predefined values stored 

by the DCIU software. The DCIU Simulator receives analog telemetry from the PMS pressure 

transducers and thermal throttle temperature sensors, which are used to regulate pressures and thermal 

throttle temperatures.

The PMS delivers regulated xenon flow to each of the three thrusters. Each thruster has three 

propellant lines and the size of this tubing can affect the xenon flow rate via backpressure (i.e., the 

pressure just downstream of the LPA thermal throttles). This pressure is a function of thruster inlet 

pressure, flow rate, and tubing dimensions. 

Test Setup

The following sections describe the test setup for the PMS multiple-string integration test. The overall 

test setup and vacuum facilities are described in the next section. This is followed by descriptions of the 

ion thruster, the PMS and XFSE test setups. 

Overall Test Setup and Vacuum Facility

The integration test was conducted in two vacuum facilities. The thrusters and PMS, along with part of 

the XFSE, were installed in the NASA GRC’s Vacuum Facility 6 (VF6), shown in Figure 6. This vacuum 

facility has a diameter of 7.6 m and an overall length of 22.9 m. It is cryogenically pumped with twelve 

internal cryogenic pumps for a total measured pumping speed of 290,000 L/s with xenon and a base 

pressure of about 1.3×10–5 Pa (1×10–7 torr). The facility also has a turbomolecular pump for the removal of 

lighter gases and a residual gas analyzer for monitoring residual gas partial pressures. Background pressures 

during full power thruster operation were within 3.7×10–4 Pa (2.8×10–6 torr), as measured by an internal ion 

gage located 41 cm below the centerline of and 25 cm behind the PM1R thruster. 
NASA/TM—2011-217040 8 

Ion Thruster Setup

Thrusters PM1R, EM1, and EM5 were mounted in VF6 on the multi-thruster array test mount, as 

shown in Figure 7 (Ref. 7). The setup was nearly identical to that used for the NEXT Single String 

Integration Test (SSIT) (Ref. 14). The PMS and XFSE were mounted behind the thrusters, as shown in 

Figure 7. Thrusters were operated with the PMS in one-string, two-string, and three-string configurations. 

Thrusters and specific throttle points selected are shown in Table 3. Thrusters were operated in four 

sequences, namely: (1) Thruster PM1R at full power, (2) Thruster PM1R at full power with Thruster EM5 

at low throttle conditions, (3) Thruster PM1R at full power with Thruster EM1 operated over the entire 

throttle range, and (4) Thrusters PM1R and EM1 at full power with Thruster EM5 operated over the entire 

throttle range. 
NASA/TM—2011-217040 9 

TABLE 3.—THRUSTERS AND THROTTLE POINTS USED

TO DEMONSTRATE PMS PERFORMANCE

Throttle point for 

thruster PM1R

Throttle point for 

thruster EM1

Throttle point for 

thruster EM5

Data set Beam 

current, 

A

Beam 

voltage, 

V

Beam 

current, 

A

Beam 

voltage, 

V

Beam 

current, 

A

Beam 

voltage, 

V

1 3.54 1800 ----- ----- ----- ----- 

2 3.54 1800 ----- ----- 1.01 275

3 3.54 1800 ----- ----- 1.20 679

4 3.54 1800 ----- ----- 2.00 1180

5 3.54 1800 1.00 276 ----- -----

6 3.54 1800 1.21 679 ----- -----

7 3.54 1800 2.01 1179 ----- -----

8 3.54 1800 2.71 1179 ----- -----

9 3.54 1800 3.54 1800 ----- -----

10 3.54 1800 3.54 1800 1.01 275

11 3.54 1800 3.54 1800 1.21 679

12 3.54 1800 3.54 1800 2.02 1179

13 3.54 1800 3.54 1800 2.72 1179

14 3.54 1800 3.54 1800 3.54 1179

15 3.54 1800 3.54 1800 3.55 1800

16 3.54 1800 3.54 1800 3.56 1800

17 3.54 1800 3.54 1800 3.55 1800

Power Consoles

Thrusters PM1R, EM1, and EM5 were operated with power consoles similar to that described in 

Reference 10. These power consoles utilized six commercially available power supplies to operate the main 

discharge, discharge cathode heater, neutralizer keeper discharge, and neutralizer heater. The consoles also 

provided constant voltages for the accelerator grid and for beam ion acceleration. Although the NEXT peak 

input power is about 7 kW with a maximum power supply voltage of 1800 V, these consoles can provide 

ion thruster input powers in excess of 10 kW with beam power supply voltages up to 2000 V.

Propellant Management System and XFSE Setup

The PMS was integrated into a XFSE assembly, as shown on the schematic in Figure 8. The XFSE is 

ground test-support equipment and, therefore, not a part of the flight system but was still controlled by the 

DCIU Simulator. The XFSE was used to monitor the operation of the PMS during the integration test and 

to provide a secondary thruster gas feed system. The XFSE was divided into two systems. The external 

XFSE was located on atmospheric side of VF6. It included mass flow controllers for checking LPA flow 

rate calibrations and could independently provide flow to the thruster using mass flow controllers, thus 

allowing for thruster operation using a standard flow regulation configuration. The internal XFSE was 

located within VF6. It included manual and solenoid valves that allowed for configuration changes, such 

as thruster operation with and without the PMS. The internal XFSE included one HPA and three LPAs for 

multi-thruster PMS operation. The internal XFSE and PMS were covered with a polyimide foil for 

protection against back-sputtered material deposition. 

Tubing and components of the PMS and XFSE were wrapped with heater tape for a bake-out to 

remove air and adsorbed moisture on surfaces exposed to atmosphere. About twenty thermocouples were 

mounted on key PMS and XFSE components and were monitored and recorded by the DCIU Simulator 

throughout testing. To simulate pressure drops due to viscosity, the tubing leading from the LPA outlet to 

the thruster inlet included 3.5 m of 0.32 cm diameter × 0.71 mm wall tubing, which is similar to the 

longest tubing length used with the Dawn (Ref. 4) ion thruster propellant management system. 
NASA/TM—2011-217040 10

For this integration test, the DCIU Simulator monitored and recorded all pressures, temperatures, and 

flows measured by the PMS and XFSE at a rate of 1 Hz. The DCIU Simulator also monitored and 

recorded PMS component input currents and voltages other than those of the pressure transducers, which 

were manually recorded. For the PMS, the DCIU Simulator regulated HPA outlet pressure, LPA outlet 

flow rates, and thermal throttle temperatures, and controlled latch valve position and the selection of 

primary or redundant thermal throttle heaters. For the XFSE, the DCIU Simulator controlled and recorded 

telemetry from the mass flow controllers.

Test Results and Discussion

This section will present the test results of the PMS integration test with multiple thrusters. PMS flow 

rate calibration checks and integration of the PMS were performed with NEXT Thrusters PM1R, EM1, 

and EM5. The PMS-multi-thruster integration test demonstrated that the PMS could successfully interface 

with and operate NEXT ion thrusters with no anomalous thruster or PMS behavior. Thrusters and throttle 

points used to demonstrate PMS performance are shown in Table 3. 

Flow Rate Calibration Checks With Single and Multiple Thrusters

Calibration checks were conducted with the PMS under vacuum and while connected to various 

combinations of operational Thrusters PM1R, EM1, and EM5. Figure 8 shows the flow path and valve 

positions. Calibration checks were performed on LPA branches for single or 2-thruster operation. Total 

LPA flow rates were compared with the calibrated flow rate from the XFSE mass flow controllers. A 

maximum of two thrusters were used in this checkout procedure because the XFSE mass flow controllers 

had a maximum flow rate of only 75, 10, and 10 sccm for the main, cathode, and neutralizer branches, 

respectively. The XFSE flow controller valves were fully opened so it would measure and not regulate the 

flow rate. To reduce mass flow controller measurement error, flow rate calibration checks were made in

5 s intervals for 5 min and averaged to determine the actual flow rate. Further, sufficient time was allowed 

between mass flow rate changes for the pressure downstream of the XFSE metering to reach equilibrium.

The XFSE mass flow controller accuracy had to be considered when assessing LPA flow rate error. The 

main mass flow controller accuracy was ±1 percent of the indicated reading for flow rates from 26 to

75 sccm. Both XFSE cathode and neutralizer flow controller accuracies were ±1 percent for flow rates from 

3.5 to 10 sccm. For all the flow rates presented herein, the accuracy of all XFSE mass flow controllers was 

±1 percent. LPA indicated flow rates were simply compared to the calibrated XFSE mass flow rates. The 

results of the calibration checks are shown in Table 4. First, Thruster PM1R was operated at full power with 

the PMS, and the total LPA flow rate was within 0.4 percent of the calibrated total flow rates of the XFSE 

mass flow controllers. Next, Thruster PM1R was operated at full power of 6.86 kW, and Thruster EM5 was 

operated at three power levels from 0.55 to 2.78 kW. In these three cases the total LPA flow rate differed 

from the XFSE calibrated total flow rate by 1.0 to 1.3 percent. Finally, Thruster PM1R was run at full power 

while Thruster EM1 was throttled from 0.55 to 2.78 kW. Again the LPA total flow rate differed from the 

XFSE total flow rate by 1.1 to 1.3 percent. Overall, the data of Table 4 indicate that the total LPA flow rates 

compare very favorably, within 1.4 percent, to the calibrated XFSE flow rates. 
NASA/TM—2011-217040 11

TABLE 4.—PMS FLOW RATE CALIBRATION CHECKS DURING SINGLE AND MULTIPLE THRUSTER OPERATION

First thruster—LPA indicated

flow rate

Second thruster—LPA indicated 

flow rate

Data set Main 

flow 

rate, 

sccm

Cathode 

flow rate, 

sccm

Neut. 

flow 

rate, 

sccm

Main 

flow 

rate, 

sccm

Cathode 

flow rate, 

sccm

Neut. 

flow 

rate, 

sccm

Total LPA 

indicated 

flow rate, 

sccm

Total 

MFC 

flow 

rate, 

sccm

Overall 

LPA error, 

percent

Thruster PM1R

1 49.7 4.87 4.01 ---- ---- ---- 58.6 58.8 0.4

Thruster PM1R Thruster EM5

2 49.7 4.87 4.01 12.4 3.53 3.00 77.5 76.5 –1.3

3 49.7 4.87 4.02 14.3 3.58 3.01 79.5 78.6 –1.0

4 49.7 4.87 4.02 25.9 3.88 2.51 90.9 90.0 –1.0

Thruster PM1R Thruster EM1

5 49.6 4.87 4.01 12.4 3.52 3.01 77.4 76.5 –1.2

6 49.7 4.87 4.01 14.3 3.57 3.01 79.5 78.5 –1.1

7 49.7 4.87 4.02 25.8 3.87 2.51 90.8 89.6 –1.3
NASA/TM—2011-217040 12

Thruster PM1R Inlet Pressure

The repeatability of PMS performance can be examined by measuring the thruster inlet pressures to the 

main plenum, the discharge cathode, and the neutralizer. These pressure measurements were made at the 

inlet to Thruster PM1R as it performed 1-thruster, 2-thruster, and 3-thruster operation with the PMS.

Pressure transducer error bands were less than ±0.3 percent of the full-scale pressure. The inlet pressure data 

are shown in Figure 9. For 2-thruster and 3-thruster operation, the main plenum and cathode inlet pressures 

are within 1.7 and 1.1 percent of each other, respectively. The main plenum pressure at full power is about 

13.4 kPA or 100 torr. The neutralizer inlet pressure did vary by as much as 13 percent. This result may 

indicate that the neutralizer cathode assembly had not yet reached thermal equilibrium. Data Set 1 provides 

inlet pressures for Thruster PM1R for single thruster operation. The main inlet pressure in this case is lower 

than all other cases, and this effect can be attributed to lack of thermal equilibrium since data were taken 

more rapidly in this situation. The main plenum inlet pressure data from Data Sets 4, 9, and 16 are nearly the 

same as Thruster PM1R data taken during the Single String Integration Test (Ref. 14). The fact that the inlet 

pressure data for 2-thruster and 3-thruster operation as well as SSIT 1-thruster operation compare very 

favorably indicates that the flow rates to Thruster PM1R were similar in all cases.

Figure 10 shows the temporal variation of the three inlet pressures to Thruster PM1R for operation of 

three thrusters at full power, 6.8 kW. LPA inlet pressures, as well as the HPA outlet pressures, were very 

steady indicating stable flow delivery with three thrusters operating at full power. 

The pressures of all LPA branches were examined every 5 s. Indicated changes in the LPA branch 

pressures were less than 0.1 percent when thrusters were added to the PMS, when cathodes were ignited, 

and when ion extraction voltages were applied. The change in HPA outlet pressure was also less than

0.1 percent when cathodes were ignited or high voltages were applied to the grids of the thrusters that 

were operated in single-, two-, or three-string configurations. However, changes in the HPA outlet 

pressure were noted when a thruster was added to the PMS and flow through the LPAs was initiated. 

HPA outlet pressure decreased by 12.6 percent when thruster EM1 cold flow was added to the PMS that 

was providing xenon to Thruster PM1R. When Thruster EM5 cold flow was added to the two-string 

system of Thrusters PM1R and EM1, the HPA inlet pressure decreased by 8.9 percent. In all cases the 
NASA/TM—2011-217040 13

HPA outlet pressure recovered within 30 s. The change in HPA outlet pressures and pressure recovery 

times are dependent on the volume between the HPA and LPA branches. Although this effect did not 

change LPA branch pressures, it may have to be examined in subsequent flight-like feed systems that will 

have much smaller volumes between the HPA and LPA branches.

Thruster Performance Data

The PMS demonstration test with single and multiple thrusters included examining any changes in 

thruster performance as more operating thrusters were added to the PMS. Performance characterizations 

included discharge losses, accelerator grid current, and neutralizer performance.

Discharge Chamber Performance

Figure 11 shows the thruster discharge voltage for 1-thruster, 2-thruster, and 3-thruster operation with 

the PMS. All data were taken at full power of 6.8 kW. All three thrusters were operated with identical 

main plenum, discharge cathode, and neutralizer flow rates. Given the fixed flow rates, the discharge 

voltage for Thrusters PM1R, EM1, and EM5 varied by less than 0.40, 0.43, and 0.50 V, respectively, as 

single or multiple thrusters were tested with the PMS. These data are consistent with the SSIT data and 

the results of two performance acceptance tests of Thruster PM1R (Ref. 14). The very small changes in 

discharge voltage for single or multiple thruster testing is a good signature of steady and repeatable flow 

rates provided by the PMS.

Figure 12 displays the variation in discharge losses at full power as one, two, or three thrusters are 

operated with the PMS. During the course of single or multiple thruster testing with the PMS, the 

discharge losses, for a given thruster, vary by less than 4 W per beam ampere. This shows that xenon flow 

delivery and thruster discharge chamber performance are insensitive to the number of thrusters operated 

with the PMS.
NASA/TM—2011-217040 14
NASA/TM—2011-217040 15 

The discharge losses of Thruster EM5 are shown in Figure 13 at throttle levels of 0.55, 1.12, and

2.78 kW. Thrusters PM1R and EM1 were run at highest throttle level of 6.86 kW. These data were taken 

for single and multiple thruster operation with the PMS. For Thruster EM5 throttle levels of 0.55, 1.12, 

and 2.78 kW, the variation in discharge losses was <5.3 percent for 1-thruster, 2-thruster, or 3-thruster 

operation with the PMS. This result is not much different from dispersions in discharge chamber 

performance when thrusters are operated individually. For example, the dispersion in discharge losses for 

Thrusters EM1, EM2, EM4, and PM1 at full power, was found to be about 7 percent (Ref. 20). 

At all three throttle levels the discharge losses of Thruster EM5 were lower when three thrusters were 

operating. This result is likely independent of PMS operation and may be due to more efficient operation 

due to xenon ingestion at the higher vacuum facility pressures. 

Accelerator Grid Current

Figure 14 shows displays the thruster accelerator grid current as one, two, or three thrusters were used 

with the NEXT PMS. All data were at the full power throttle level: 3.52 A beam current and 1800 V beam 

power supply voltage. Accelerator currents for Thrusters PM1R and EM1 were nearly the same for single 

thruster operation. During single thruster operation, the accelerator current for Thruster EM5 was

8.8 percent higher than Thruster EM1 at full power. In dispersion analyses of the EM thrusters (Ref. 20),

Thruster EM5 had a much larger peak ion current density, and this might account for the difference in 

accelerator grid currents. Additionally, the local vacuum facility pressure in the vicinity of Thruster EM5 

may be slightly higher because it is immediately adjacent to both Thrusters EM1 and PM1R. 

For 2-thruster or 3-thruster operation with the PMS, the accelerator grid currents of Thrusters PM1R 

and EM1 were the same. Higher accelerator grid currents for multiple thruster operation are expected due 

to increased charge exchange ion impingement caused by higher vacuum facility pressures. The 

accelerator current for Thruster EM5 again is about 9 percent higher than that of Thrusters PM1R and 

EM1 during 3-thruster operation. For the most part the accelerator grid current values were as expected, 

and no anomalies were apparent due to single or multiple thruster operation with the PMS.
NASA/TM—2011-217040 16

Neutralizer Performance

The neutralizer keeper and coupling voltages for full power operation are shown in Figure 15. 

Neutralizer keeper and coupling voltages were within 1.0 and 1.6 V, respectively, for all three thrusters 

operating as single or multiple thrusters with the PMS. Dispersion data taken at full power with Thruster 

PM1R and Thruster EM1, operating in 1-thruster, 2-thruster, and 3-thruster configurations, indicated the 

neutralizer keeper and coupling voltages varied by less than 0.5 and 1.1 V, respectively. Thruster EM5, 

operating in 1-thruster and 3-thruster configurations, had neutralizer keeper and coupling voltage 

variations of only 0.37 and 0.22 V, respectively at full power. The small changes in neutralizer voltages 

support the fact that xenon flow rates are constant for single or multiple thruster testing with the PMS. 

PMS Power Consumption

The propellant management system power consumption is tabulated in Table 5. These powers were 

determined from currents and voltages that were measured by the DCIU Simulator for the thermal 

throttles and the proportional flow control valves, and by multi-meters for the pressure transducers. In all 

cases the thermal throttle temperatures were set at 75 °C, and the input voltage to the DCIU Simulator 

was 28 V. The total power consumed by the HPA and LPAs was 10.0, 17.9, and 25.2 W, respectively, for 

single, 2-thruster, and 3-thruster operation. In all cases, the HPA consumed 1.6 to 1.7 W. The thermal 

throttle heaters accounted for about 52 percent of the total PMS power for single or multiple thruster 

operations. The PMS power consumption for single thruster operation was nearly identical to results 

obtained during the single-string integration test of the PMS and Thruster PM1R (Ref. 14). The PMS 

demonstrated successful operation over a wide throttling range for single and multiple thrusters. 
NASA/TM—2011-217040 17

TABLE 5.—NOMINAL PMS POWER CONSUMPTION FOR ONE-, TWO-, AND THREE-STRING OPERATION

[Low power bus voltage: 28 V. Thermal throttle temperatures: 75 °C.]

Thruster(s) HPA Power, W LPA Power, W Power, W

Pressure 

transducers

PFCV Total Pressure 

transducersa

PFCV Thermal 

throttle

Total

OnestringPM1R 0.96 0.64 1.60 1.42 1.78 5.20 8.40 10.0

TwostringPM1R 1.42 1.87 4.58 7.87

EM1 1.42 2.19 4.77 8.38

PM1R + EM1 0.96 0.68 1.64 2.84 4.06 9.35 16.25 17.9

ThreestringPM1R 1.42 1.93 4.10 7.45

EM1 1.42 2.25 4.30 7.97

EM5 1.42 2.16 4.50 8.08

 PM1R + EM1 

+ EM5

0.96 0.73 1.69 4.26 6.34 12.90 23.50 25.2

a

Pressure transducer power levels assumed EM transducer power values since the two commercial pressure transducer’s power consumption was not 

representative of flight-rated pressure transducers.
NASA/TM—2011-217040 18

Propellant Management System and Propulsion System Satisfied Requirements

This sub-system integration test of the PMS, the DCIU Simulator, and multiple thrusters addressed, in 

part, some of the NEXT PMS and ion propulsion system performance and functionality requirements. 

Some of these requirements are identified here. The PMS demonstrated propellant flow control 

commensurate with single or multiple thruster operations over the system throttling range. The PMS 

design had the capability to simultaneously deliver xenon to 3 thrusters operating at full power. With all 

three thrusters operating, the PMS, DCIU Simulator, and thruster data demonstrated the xenon flow rates 

and thruster performance were consistent with project requirements. The DCIU Simulator did provide 

sub-system throttling capabilities with single or multiple thrusters over the throttling range. The subsystem engineering data stream, rate of data collected for storage, and collection intervals were adjustable 

through the DCIU Simulator. Finally, the PMS did accommodate the back-pressure at the three thruster 

interfaces over the throttling range. 

Concluding Remarks

As a critical part of the NEXT test validation process, a multiple-string integration test was performed 

on the NEXT propellant management system and ion thrusters. The objectives of this test were to verify 

that the PMS is capable of providing stable flow control to single or multiple thrusters operating over the 

system throttling range and to demonstrate to potential users that the NEXT PMS is ready for transition to 

flight. Propulsion sub-system elements included in this integration test were: 1) an engineering model ion 

thruster, labeled PM1R, that has successfully completed environmental testing at qualification levels;

2) two development-type engineering model thrusters; 3) an engineering model propellant management 

system that has successfully completed environmental testing at qualification levels; 4) three laboratoryclass power consoles; and 5) a breadboard DCIU Simulator that acted as a test console. 

A test plan was developed for the sub-system integration test for verification of PMS and IPS 

performance and functionality requirements. Propellant management system calibrations were checked 

during the single and multi-thruster testing. Calibration checks were performed on total LPA branch flow 

rates by comparing them to XFSE mass flow controller flow rates that were previously calibrated against 

a standard. A maximum of two thrusters were used in this procedure because the XFSE mass flow 

controllers had a maximum xenon flow rate of only 75, 10, and 10 sccm for the main plenum, cathode, 

and neutralizer branches, respectively. Overall, the calibration check of LPA total flow rates to the 

thruster(s) compare very favorably, within 1.4 percent, of the calibrated XFSE flow rates. Also, the inlet 

pressures to the main, cathode, and neutralizer ports of Thruster PM1R were measured as the PMS 

operated in 1-thruster, 2-thruster, and 3-thruster configurations. Variations of the inlet pressures with time 

were found to be insignificant, indicating steady flow to Thruster PM1R. It was found that the inlet 

pressures to Thruster PM1R for 2-thruster and 3-thruster operation as well as single thruster operation 

with the PMS during the SSIT compare very favorably indicating that flow rates to Thruster PM1R were 

similar in all cases. 

Characterizations of discharge losses, accelerator grid current, and neutralizer performance were 

performed as more operating thrusters were added to the PMS. There were very small changes in 

discharge voltage for single or multiple thruster testing which is a good signature of steady repeatable 

flow rates provided by the PMS. With thrusters operating a full power, the discharge losses, for a given 

thruster, varied by less than 4 W per beam ampere during the course of single or multiple thruster 

operations with the PMS. There were also no unexpected variations in discharge losses as thrusters were 

throttled and single and multiple thruster operations were conducted. The accelerator grid current values 

were as expected, and no anomalies were apparent due to single or multiple thruster operations with the 

PMS. Neutralizer keeper and coupling voltages were within 1.0 and 1.6 V, respectively, for all three 

thrusters operating a single or multiple thrusters with the PMS. This result confirms that flow rates to the 

thrusters are constant for single or multiple thruster testing.
NASA/TM—2011-217040 19

The propellant management system power consumption was at a fixed voltage to the DCIU and a 

fixed thermal throttle temperature of 75 °C. The total power consumed by the HPA and the LPAs was 

10.0, 17.9, and 25.2 W, respectively, for single, 2-thruster, and 3-thruster operation with the PMS. 

This sub-system integration test of the PMS, the DCIU Simulator, and multiple thrusters addressed, in 

part, some of the PMS and IPS performance and functionality requirements. Some of the key 

requirements addressed were: 1) The PMS demonstrated providing propellant flow control commensurate 

with single or multiple thruster operations over the system throttling range; 2) With all three thrusters 

operating, the PMS, DCIU Simulator, and thruster data demonstrated the xenon flow rates and thruster 

performance were consistent with project requirements. 3) The DCIU Simulator did provide sub-system 

throttling capabilities with single or multiple thrusters over the throttling range; 4) The PMS did 

accommodate the back-pressure at the three thruster interfaces over the throttling range. 

References

1. Patterson, M.J. and Benson, S.W., “NEXT Ion Propulsion System Development Status and 

Performance,” AIAA Paper 2007–5199, July 2007. 

2. Benson, S.W. and Patterson, M.J., “NEXT Ion Propulsion System Progress Towards Technology 

Readiness,” AIAA Paper 2008–5285, July 2008.

3. Polk, J.E., Brinza, D.E., Kakuda, R.Y., Brophy, J.R., Katz, I., Anderson, J.R., Rawlin, V.K., 

Patterson, M.J., Sovey, J.S, and Hamley, J.A., “Demonstration of the NSTAR Ion Propulsion System 

on the Deep Space One Mission,” IEPC Paper 01-075, October 2001. 

4. Brophy, J.R., Garner, C.E., and Mikes, S.C., “Development Dawn Ion Propulsion System – Initial 

Checkout after Launch,” AIAA Paper 2008-4917, July 2006. 

5. Hoskins, W.A., Wilson, F.C., Polaha, J., Talerico, L.A., Patterson, M.J., Soulas, G.C. and Sovey, 

J.S., “Development of a Prototype Model Ion Thruster for the NEXT System,” AIAA Paper 2004-

4111, July 2004.

6. Snyder, J.S., Anderson, J.R., Van Noord, J.L. and Soulas, G.C., “Environmental Testing of the NEXT 

PM1R Ion Engine,” IEPC Paper 2007-276, September 2007. 

7. Patterson, M.J., Foster, J., McEwen, H., Pencil, E., Van Noord, J.L. and Herman, D.A., “NEXT 

Multi-Thruster Array Test – Engineering Demonstration,” AIAA Paper 2006-5180, July 2006.

8. Aadland, R.S., Frederick, J., Benson, S.W. and Malone, S.P., “Development Results of the NEXT 

Propellant Management System,” JANNAF Proceedings, December 2005.

9. Hoskins, W.A., Wilson, F.C., Aadland, R.S., Monheiser, J., Talerico, L.A. and Polaha, J., “Overview 

of the NEXT Ion Propulsion System Program at Aerojet,” AIAA Paper 2005-3885, July 2005.

10. Pinero, L.R., Patterson, M.J., and Satterwhite, V.E., “Power Console Development for NASA’s 

Electric Propulsion Outreach Program,” IEPC Paper 93-250, September 1993. 

11. Soulas, G.C., Domonkos, M.T., and Patterson, M.J., “Performance Evaluation of the NEXT Ion 

Engine,” AIAA Paper 5278, July 2003. 

12. Herman, D.A., Soulas, G.C., and Patterson, M.J., “Performance Evaluation of the Prototype-model 

NEXT Ion Thruster,” AIAA Paper 2007-5212, July 2007. 

13. Snyder, J.S., Anderson, J.R., Van Noord, J.L. and Soulas, G.C., “Environmental Testing of the NEXT 

PM1 Ion Engine,” AIAA Paper 2007-5275, July 2007.

14. Soulas, G.C., Patterson, M.J., Pinero, L.R., Herman, D.A., and Snyder, J.S., “NEXT Single String 

Integration Test Results,” AIAA Paper 2009-4816, July 2009.

15. Soulas, G.C., Kamhawi, H., Patterson, M.J., Britton, M.A., and Frandina, M.M., “NEXT Ion Engine 

2000 Hour Wear Test Results,” AIAA Paper 2004-3791, July 2004.

16. Patterson, M.J., Verhey, T.R., Soulas, G.C., and Zakany, J.S., “Space Station Cathode Design, 

Performance, and Operating Specifications,” IEPC Paper 97-170, August 1997. 

17. Aadland, R.S., Monheiser, J.M., Driscoll, E.A., Wilson, F.C. and Benson, S.W., “Development Status 

of the NEXT Propellant Management System,” AIAA Paper 2004-3974, July 2004.
NASA/TM—2011-217040 20

18. Aadland, R.S. and Talerico, L.A., “NEXT Propellant Management System (PMS) Thermal Vacuum 

Cycle Test Report,” Aerojet Internal Report 2007-R-2948, Aug. 14, 2007. 

19. Hoskins, W.A., Aadland, R.S., Meckel, N.J., Talerico, L.A., and Monheiser, J.M., “NEXT Ion 

Propulsion System Production Readiness,” AIAA Paper 2007-5856, July 2007. 

20. Soulas, G.C. and Patterson, M.J., “NEXT Ion Thruster Performance Dispersion Analyses,”

AIAA Paper 2007-5213, July 2007.
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NEXT Propellant Management System Integration With Multiple Ion Thrusters 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

As a critical part of the NEXT test validation process, a multiple-string integration test was performed on the NEXT propellant management system and ion thrusters. The 

objectives of this test were to verify that the PMS is capable of providing stable flow control to multiple thrusters operating over the NEXT system throttling range and to 

demonstrate to potential users that the NEXT PMS is ready for transition to flight. A test plan was developed for the sub-system integration test for verification of PMS and 

thruster system performance and functionality requirements. Propellant management system calibrations were checked during the single and multi-thruster testing. The low 

pressure assembly total flow rates to the thruster(s) were within 1.4 percent of the calibrated support equipment flow rates. The inlet pressures to the main, cathode, and 

neutralizer ports of Thruster PM1R were measured as the PMS operated in 1-thruster, 2-thruster, and 3-thruster configurations. It was found that the inlet pressures to Thruster 

PM1R for 2-thruster and 3-thruster operation as well as single thruster operation with the PMS compare very favorably indicating that flow rates to Thruster PM1R were 

similar in all cases. Characterizations of discharge losses, accelerator grid current, and neutralizer performance were performed as more operating thrusters were added to the 

PMS. There were no variations in these parameters as thrusters were throttled and single and multiple thruster operations were conducted. The propellant management system 

power consumption was at a fixed voltage to the DCIU and a fixed thermal throttle temperature of 75 °C. The total power consumed by the PMS was 10.0, 17.9, and 25.2 W, 

respectively, for single, 2-thruster, and 3-thruster operation with the PMS. These sub-system integration tests of the PMS, the DCIU Simulator, and multiple thrusters 

addressed, in part, the NEXT PMS and propulsion system performance and functionality requirements.

15. SUBJECT TERMS

Ion thruster; Ion engine; Xenon propellant; Xenon feed system; Xenon; Ion propulsion system; Propellant management system 

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