NASA Technical Memorandum 102337 

I. 

Electric Propulsion Options for 10 kW Class 

Earth Space Missions 

M.J. Patterson and F.M. Curran 

Lewis Research Center 

Cleveland, Ohio 

Prepared for the 

1989 JANNAF Propulsion Meeting 

Cleveland, Ohio, May 23-25, 1989 

NASA 

- -- - _- 

(NA SA-TB- 10 23 3 7) E LE CT E1 C PBOPU LS IO El 

OPTIONS FOR 10 kL CLASS EARTH SPACE NISSIO8S 

N 89- 269 06 

(NASA. Lewis Research Center) 29 pCSCL 21C Unclas 

G3/20 0225966 
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ELECTRIC PROPULSION OPTIONS FOR 10 KW CLASS 

EARTH-SPACE MISSIONS 

M.J. Patterson and F.M. Curran 

NASA Lewis Research Center 

Cleveland, Ohio 44135 

Abstract 

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system perforniance and system component mass estimates. Overall mission performance of the electric propulsion 

systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment 

capability available when integrated onto a generic spacecraft under fixed mission model assumptions. 

Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA I1 

6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 

hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v’s range from 

1.2 x lo7 to 2.1 x lo’ N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the 

arcjet propelled spacecraft launched on the DELTA I1 6920 vehicle range from approximately 528 hours 

for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum 

total impulse values and maximum delta-v’s range from 2.2 x lo8 to 3.6 x 10‘ N-s, and approximately 662 

t,o 1072 m/s, respectively. 

INTRODUCTION 

A high level of interest has been developing for a near-term demonstration of primary solar electric 

propulsion’-*. The proposed ”strawman” electric propulsion (EP) space experiment would act to assist 

transition of primary electric propulsion to operational status through development and flight qualification 

of realistic electric propulsion system(s). A successful demonstration of the electric propulsion system(s) 

would verify thruster system performance and lifetime, and establish the capabilities of this technology to 

accomplish high-energy missions. Other advantages of such a space experiment may include the demonstration of autonomous flight control of a low thrust spacecraft, investigations of spacecraft/plasma/EP system 

int,eractions, and concurrent demonstration of advanced solar array technologies. 

Ion and arcjet propulsion are options for primary electric propulsion missions because of their specific impulse, efficiency, and development maturity. This paper examines the ion and arcjet thruster and propulsion 

syst,ein options available for a near-term space demonstration mission. First-order calculations of propulsion 

system performance and component masses are provided. Overall mission performance of the proposed EP 

syst,eiiis is quantified in terms of total impulse, delta-v, and thrusting time capability based on a fixed mission 

inodel. 

MODEL DESCRIPTION 

Assessing a propulsion system requires selecting the thruster technology (performance and operating condition) and deriving the system parameters (component masses and architecture) based on that technology. 

This section describes the thruster and system mass models used for the ion and arcjet propulsion systems, 

and the mission model used in characterizing their performance. 

“Approved for public release; distribution is unlimited.” 

1 
ION PROPULSION SYSTE MS 

Thruster - The ion thruster, shown in Figure 1, generates thrust by electrostatic acceleration of ions extracted 

from a plasma discharge. Predicting the performance (thrust, specific impulse, power level, etc.) of an ion 

thruster with various propellants is fairly straightforward because the device operation is determined by well 

known physical operating principles. In general, the performance of a given ion thruster can be specified to 

within 5 percent accuracy for most parameters prior to it's operation. The assumptions used in determining 

the thruster performance for this study are listed in Appendix A. Only 30 cm diameter (28.7 cm effective 

beam diameter) ion thrusters were considered in these analyses due to the relative maturity of this technology. 

A ring-cusp discharge chamber was assumed as it has demonstrated improved efficiencies compared to the 

SEPS divergent-field ion thr~ster~-~. Further, high power operation of the divergent-field thruster with 

xenon has shown unacceptable levels of discharge component erosion over a wide range of input power levels 

of interest5. This indicates that the divergent-field thruster has insufficient life for high total impulse nussions 

using inert gas propellants. This high erosion has not been observed in high power ring-cusp ion thruster 

testing to date. Xenon was selected as the propellant because of its high atomic mass which gives a high 

thrust-to-power ratio, and because it is the propellant for which the majority of high power testing has 

been conducted. Figures 2 and 3 show thruster efficiencies, and thrust levels, respectively, versus specific 

impulse achievable with present 30 cm ion thruster technology on xenon propellant. The assumptions used 

in obtaining these values for both figures are listed in Appendix A. These predicted performance levels for 30 

cm ion thruster technology correspond favorably to results obtained at NASA-LeRC and else~here~-~. As 

indicated in Figure 2, 30 cm xenon ion thrusters can run from approximately 2450 seconds to 4520 seconds 

specific impulse, which corresponds to 62 to 79% overall thruster efficiency. The efficiency of electrostatic 

thrusters increases with specific impulse, as an increasing fraction of the input power goes directly into the 

beam. For completeness, data were included over a range of discharge chamber propellant efficiencies (from 

90 to 95%), which accounts for the vertical band in the efficiency curve. Thrust levels with xenon propellant 

range from approximately 0.04 to 0.42 Newtons, which corresponds to thruster input power levels from 

approxiniately 1 to 12 kW (Figure 3). 

As indicated in Figure 3, the plot of thrust versus specific impulse defines an operating envelope for 30 

cm ion thrusters. The boundaries for the envelope include: 

1. a lower-limit range of specific impulse, defined by a net-to-total accelerating voltage ratio (R ratio) 

of 0.55. Reducing the R ratio below this value with two grid optics results in defocussing and directimpingement of the ion beamlets onto the accelerator electrode. The lower limit to the specific impulse 

could be reduced further by the use of three grid optics which permit operation at lower values of R, 

but this would result in a reduction in the absolute value of the thrust. 

2. an upper-limit range of specific impulse, defined by an R ratio of 0.85. Increasing the R ratio beyond this 

value results in electrons in the neutralized ion beam backstreaming to the screen (positive) electrode. 

3. a maximum specific impulse, thrust, and thruster input power, defined by the electric field strength 

linit (or equivalently, ion extraction linlit) of the ion optics. 

4 an upper-limit to the thruster input power, and thrust, defined by the heat rejection capability of 

the thruster discharge chamber. As noted in Appendix A, ring-cusp ion thrusters incorporate rareEarth cobalt magnets in the discharge, which arc susceptible to irreversible loem when their surface 

temperature exceeds 300 "C. This corresponds to approximately 800 Watts discharge power for present 

30 cm ring-cusp thruster designs. Figure 3 shows for xenon propellant, the discharge thermal limit and 

the ion optics extraction limit are reached at approximately the same condition. The thruster thermal 

limit may be increased by an improved thermal design (e.g.- larger diameter discharge chamber) or by 

an improved plasma containment geometry which reduces discharge losses. 

Thirt,y centimeter diameter ion thrusters have operated for short duration tests at performance levels 

subst.antially higher than those shown in Figure 34. However, the authors believe that the performance 

parameters indicated in this figure encompass the range of moderate risk operating conditions presently 

available, which lead to adequate thruster lifetimes. The xenon operating envelope spans the range of 

performance obtained from the 1.4 kW, 25 cm diameter thruster of Beattie, et al.', to the 10 kW, 30 cin ion 

thrusters presently under testing at NASA-LeRC. 

2 
It is noted that all the operating conditions within the performance envelopes of Figure 3 are not of equal 

risk, in terms of thruster endurance. Qualitatively, the minimum risk operating conditions are at the lowest 

power density and near the center of the specific impulse band. Reduced thruster life occurs by increasing 

input power (thrust), or by changing the specific impulse to a value near the prescribed limits. These occur 

when the total accelerating voltage across the optics is increased, when the beam current is increased (either 

by increasing discharge power or total accelerating voltage), or when the R ratio is decreased/increased 

to beyond 0.55/0.85. Jon thrusters are typically operated near 0.80 R ratio, below the point of electron 

backstreaming but high enough to maximize the engine thrust. Substantial reductions in the R ratio lead 

to defocussing and direct impingement (as previously discussed) and also lead to increased charge-exchange 

erosion of the accelerator grid electrode due to the higher operating voltages. 

Although ion thruster lifetime at high power levels has not been verified, it is believed that projections 

based on ion erosion of the screen electrode to 50% of its initial thickness provides a conservative measure of 

thruster life, Based on the 10 kW wear test of a xenon divergent-field thruster', and short-term erosion rate 

measurements performed on a 1.4 kW xenon ring-cusp thruster", a 35,000 A-h screen grid (beam current 

times operating time) lifetime limit at 28 V discharge voltage is proposed as the best estimate for xenon 

thruster life presently available. Based on this criterion, the xenon ion thruster lifetimes associated with the 

operating conditions of Figure 3 would range from approximately 5,800 to 35,000 hours, corresponding to 

the range of input power levels from approximately 12 kW down to 1 kW. 

System Mass __ Model ~ - The propulsion system mass model used in this study was derived from the architecture/methodology proposed in reference 11. It's implementation is consistent with that of reference 12. 

This mass model defines the propulsion system into a thrust module and an interface module (see Table 

I). The thrust, module includes the thrusters (self radiating), gimbals, power processor units (ppu) and ppu 

therinal control, thruster structure, and propellant distribution. The interface module consists of converters, 

controllers, power distribution cabling, structural mass (comprising the 'dry' interface module mass), and 

the propellant and propellant storage and assembly (comprising the propellant module mass). The sum of 

the thrust module mass and the 'dry' interface module mass are defined here as the propulsion system dry 

mass. Reference 12 lists the equations used to calculate the component masses of the thrust and interface 

modules. 

No redundancy was included in the propulsion system, except for the power distribution cabling, converters, and controllers. Also, the mass of the guidance/navigation system was assumed to be part of the 

spacecraft (payload) and hence was not included in the evaluation of the propulsion systems. The question 

of what is the correct number of active thrusters to incorporate into a system design is a function of the 

propulsion system power, mass, and volume allotment, as well as the tradeoffs in risk (thruster endurance) 

versus propulsion system cost (including qualification testing) and complexity. Based on mission model 

assumptions, systems ranging from 4 active thrusters processing a total of 5 kW to 1 thruster processing 10 

kW were examined over the range of operating conditions identified in Figure 3. 

ARCJET P_RLUPUSECHNBTEMS 

3krustc.r - Arcjet thrusters are electrothermal devices in which an arc is used to heat the propdlant. A 

cutaway scheinatic of a typical laboratory model arcjet thruster is shown in Figure 4. A high-voltage pulse 

is used to start the arcjet. In operation the arc cathode attachment point is a small, molten spot at the 

cathode tip. The arc attaches in a diffuse manner in the divergent section of the nozzle which also acts as 

the anode. Propellant is injected into the chamber tangentially to provide swirl stabilization of the arc. 

T11e current NASA-sponsored research program on low-powered arcjet thrusters was st trted in 1983. 

Under this program, operation in the 1 to 2 kW range has been demonstrated using hydrazine, or mixtures 

simulating it '5 decomposition products, as the propellant. Hydrazine was chosen because the near-term goal 

of this work is the application of the arcjet to North-South Station Keeping (NSSK) on geosynchronous 

communication satellites which currently use this propellant. Ammonia has also been tested in the range of 

interest. 

A 1000 hour/500 cycle lifetest was successfully completed on a lab model thruster to demonstrate longterm reliability". Typical performance measurements taken using this thruster sre shown in Figure 5. These 

data show that there was no significant change in the thruster operating characteristics over the course of 

the test. While this test was run on mixtures simulating hydrazine decompostion products, other tests have 

shown the operating characteristics of the arcjet are virtually identical when operated on hydrazi~~e'~. In 

addition to the above, both the impacts of the and electromagnetic interference (EMI) are ultder 

investigation. 

3 
An engineering model flight-type arcjet system is currently nearing completion and lifetats of this ryrtem 

are planned for the near future. A preliminary investigation of arcjet operating characteristics at increased 

power levels has also been started. 

Additionally, the Jet Propulsion Laboratory (JPL) has been establishing a large performance and lifetime 

database for 30 kW-class ammonia arc jet^"-^^. This program supports the development of arcjet thruster 

technology for the SP-100 Flight Experiment for which arcjet EP has been baselined as the active load'. 

System Mass Model - The arcjet propulsion system mass model used in these analyses was based on the 

methodology developed in reference 11, and was consistent with that shown in Table I and published by 

Hardy et al.lg, except for the modifications identified in Appendix B. Three different thrusters and thruster 

power levels (approximately 1.5, 4.6, and 9.4 kW) were selected based on the database established with the 

1 kW NASA-LeRC arcjet13. This criterion resulted in system designs which ranged from 1 to a maximuin 

of 6 active thrusters available to process 5 to 10 kWe from the power source. The issue of redundancy was 

treated as described previously in the ion system mass model description. 

MISSION MODEL 

~ - - _____ 

A mission model was used to calculate the capabilities of each of the low power ion and arcjet systems, 

once the dry masses of these propulsion systems were determined. The mission model is described in the 

following assumptions: 

A low power ion or arcjet system would fly on a dedicated electric propulsion space experiment. The 

total payload mass (not including the low power ion or arcjet propulsion unit or propellant) would 

be 2950 kg. This would include the spacecraft bus, an energy storage module, solar arrays, ai1 EP 

diagnostic package, and an autonomous EP control unit. 

Either 5 or 10 kWe would be available from a solar array for the low power ion or arcjet propulsion 

module. 

The EP spacecraft would be launched from the Eastern Space and Missile Center (ESMC) either on 

an ATLAS I, ATLAS 11, DELTA I1 6920, or DELTA I1 7920 expendable launch vehicle. The ATLAS 

vehicles would do a direct planar ascent insertion of the EP spacecraft into a circular orbit of 550 kiii 

altitude. The DELTA I1 vehicles would insert the EP spacecraft into a circular orbit at 550 km altitude 

using a similar two-stage mission profile, but at a 28.7 degree inclination. The initial circular orbit 

altihde of 550 km was selected to ensure at least an order of magnitude in the thrust-to-drag ratio for 

all the proposed spacecraft. This altitude was determined assuming a maximum acceptable drag force 

of 5.0 x lo-' N/m2, using nominal atmospheric density projections for peak conditions during solar 

cycle 22 from reference 20. The payload capability of each vehicle to this orbit is 3600 kg, 4500 kg, 

4600 kg, and 5100 kg, for the DELTA I1 6920, ATLAS I, DELTA I1 7920, and ATLAS 11, respectively. 

These launch vehicle performance numbers were obtained from references 21 and 22. 

The mission would require a total thrusting time demonstration sufficient to verify the performance 

and lifetime of the thruster technology. The thrusting time would be specified based on either the 

launch vehicle payload margin (termed "Mission Mode I") or by an estimate of future EP system 

inissioii requirements (termed "Mission Mode 11"). In these analyses, both approaches were addressed. 

The performance of the low power ion and arcjet systems in Mission Mode I were characterized by the 

maximum total impulse, velocity increment, and thrusting time capability of each system, using the 

launch vehicle payload margin for the addition of propellant. Eased on the evaluation of arcjet and 

ion system mass and thrust, and the lift mass of the selected launch vehicles, a thrusting time of 5000 

hours for the ion systems, and 1000 hours for the arcjet systems was used for Mission Mode 11. The 

t,wo EP system types provide comparable values of total impulse for these thrusting times. 

As low thrust trajectory analyses (including occultation, eclipse and drag penalties, array degradation, 

etc.) was beyond the scope of this investigation, the velocity increments were determined solely from the 

'rocket equation' without consideration for thruster duty cycle/restart requirements, or throttling strategies. 

4 
RESULTS AND DISCUSSION 

This section describes the ion thruster performance parameters, ion propulsion system options, and 

arcjet propulsion system options available for an for an electric propulsion space experiment. The mission 

performance of each system, in terms of total impulse, delta-v and thrusting time capability is also quantified 

and presented. 

ION THRUSTER PERFORMANCE AND PROPULSION SYSTEM OPTIONS 

As suggested in Figures 2 and 3, the operating conditions available to a an ion thruster with a single 

propellant are quite varied. Specifying the power delivered to the thruster or thruster ppu is not sufficient 

to define the thruster operating point, as several operating conditions (specific impulse, efficiency, thrust, 

etc.) are available at any one power level. The mission model assumes both 5 and 10 kW are available 

from a solar array to the ion propulsion module. Since the ppu efficiency and line losses were assumed to 

be 92% and 0.5%, respectively, the total power delivered to the ion thrusters was approximately 4.6 and 

9.2 kW. As discussed earlier, the number of thrusters per power level would range from 1 to a maximum 

of 4. Consequently, a selection of thruster operating condition must be made so that an integral number 

of thrusters will process the power available, and yet be consistent with the operating envelope and criteria 

previously discussed. 

One way to select the ion thruster operating conditions would be to select the rpecific impulse (I,p) 

from within the available range determined by the engine parameters. Selecting the Isp specifies the beam 

voltage. To ensure that an integral number of thrusters would process all the available power requires that 

the thrusters be throttled by reducing the beam current. However, this could reduce the total propulsion 

system thrust significantly. 

By judiciously selecting the total accelerating voltage and beam voltage, and thereby specifying the I,, 

and power level, an integral number of thrusters can be obtained without throttling the beam current This 

method ensures that the minimum total voltage required is selected, the R ratio is kept near 0.80 (minimizing 

the accelerator voltage), and the thrust is maximized. Table I1 lists the 30 cm ion thruster performance 

parameters for 6 operating conditions with xenon propellant which arose from the 5 and 10 kW single and 

niultiple engine system designs. The thruster input power levels range from approximately 1 to 9 kW, for 

an Isp range of approximately 3200 to 4400 seconds. 

Table 111 outlines the top-level design parameters of the 5 and 10 kW xenon ion propulsion system options, 

including total propulsion system thrust, efficiency, and dry mass. Total thrust for the xenon propulsion 

systems range from 0.180 to 0.208 Newtons for the 5 kW options, to 0.329 to 0.392 Newtons for the 10 kW 

options. These thrust levels correspond to overall propulsion system efficiencies of 65 to 72%, respectively. 

Table IV provides more system detail as it breaks down the ‘dry’ component masses of these ion propulsion 

systems. As indicated, the propulsion system dry masses are strongly dependent on total input power and 

number of thrusters. The propulsion system dry masses range from a maximum of 234 kg for a 4-engine 10 

kW system, to a minimum of 79 kg for a 1-engine 5 kW system. Figure 6 presents the propulsion system 

dry mass as a function of specific impulse for these various systems. The number of thrusters per system is 

also identified in this figure. 

Figures 7(a) and 7(b) show the dry mass distribution of the 5 and 10 kW xenon ion propulsion systems, 

respectively. The 10 kW ion thruster systems are approximately 12 to 28 percent heavier, for the same 

number of the thrusters, than the 5 kW systems. The mass distribution as a function of number of thrusters 

is generally the same for both power levels. The majority of the mass is in the power processing, constituting 

over 40% of the total dry mass for all the system options. The thruster/gimbal mass is the second most 

massive component, except for the single-engine 10 kW case where the thermal control mass exceeds the 

thruster and gimbal masses. 

ARCJETTHRUSTER PERFORMANCE ANQ PROPULSION SYSTEM OPTIONS 

Performance estimates for the 1.5, 4.6, and 9.4 kW ammonia and hydrazine arcjets are listed in Table 

V. Listed values include thrust, I,p, efficiency, and propellant mass flowrate as a function of thruster input 

power. The performance estimates used are based on conservative projections and are consistent with results 

obtained at NASA-LeRCI3 and at JPL”. 

Table VI outlines the top-level system parameters for the 5 and 10 kW ammonia and hydrazine arcjet 

propulsion system options. The parameters include total system thrust, efficiency, and total dry mass. Total 

5 
arcjet system thrust with both propellants is typically a factor 2 to 3 higher than that projected for the 

xelloll ion propulsion systems at the same input power. Hydrazine arcjet system thrust levels rarlge from 

0.423 to 0.593 Newtons for the 5 kW options, and 0.810 to 1.187 Newtons for the 10 kW options. These 

thrust levels correspond to system efficiencies of 30-32%. 

As indicated in Table VI, the propulsion system dry mass is independent of propellant and depends only 

on total power and the number of thrusters. The propulsion system dry masses range from a minimum of 

38 kg for a single-engine 5 kW system, to a maximum of 87 kg for a 6-engine 10 kW system. These dry 

mass values are a factor of 2 to 3 lower than those projected for the ion propulsion systems for the same 

power and numbers of thrusters (see Table 111). Figure 8 shows the arcjet propulsion system dry masses as 

a function of specific impulse. 

Table VI1 breaks the mass of the arcjet propulsion system options down into the subsystem level. For all 

of the arcjet systems, the power processor is the most massive system component, constituting anywlierr from 

28 to 34% of the total dry mass. The second most massive component is the dry interface module, except 

for the 6- and 2-engine 10 kW systems where the thermal control is the second heaviest component. Unlike 

the ion propulsion systems, the thruster/gimbal masses in the arcjet systems are fairly small, constituting 

only 6-8% of the total mass. Figure 8 shows the arcjet propulsion system dry masses as a function of specific 

impulse. Figure 9 shows the dry mass distribution of each of the arcjet propulsion systems. 

MISSION PERFORMANCE 

Values for total initial dry spacecraft mass for each combination of electric propulsion (EP) system 

option/payload were determined by individually adding each EP system dry mass to the 2950 kg payload. 

Subtracting these dry mass values from the mass lift capability of the launch vehicles defines a positive 

payload margin. The payload margin, in part or in whole, is then consumed by the addition of propellalit 

and tankage (the propellant module mass) for the EP system. 

Mission Mode I - In this mode the mission parameters for the EP options are determined by the launch 

vehicle payload margin. This is accomplished by filling the payload margin with propellant for the EP 

system to increase the thrusting time and delta-v capability of the ion and arcjet systems. Tables VI11 

(a) and (b) list the spacecraft system and imssion parameters calculated for this extended-duration mission 

using the various ion and arcjet propulsion system options. The ammonia arcjet numbers are not shown, as 

the system and mission parameters with this propellant are comparable to those obtained with hydrazine 

propellant which are listed. 

The xenon 4-engine 10 kW ion propelled spacecraft has the lowest positive payload margin of any system 

option when launched on the DELTA I1 6920 vehicle. Yet, as indicated in Table VIII, there is still sufficient 

payload margin to on-load enough propellant to run the propulsion system for more than 8,600 hours, giving 

a delta-v of 3600 m/s. Enough propellant could be on-loaded on most of the ion propelled spacecraft to 

permit thruster operation greatly in excess of the projected thruster lifetimes. That is, the capacity to 

fully demonstrate and qualify the ion propulsion systems is not limited by the lift capacity of the proposed 

launch vehicles, but is restricted by the engine life. Maximum available thrusting times for the ion-propelled 

spacecraft on the DELTA I1 6920 vehicle range from approximately 8,600 hours for the 4-engine 10 kW 

system to more than 29,600 hours for the single-engine 5 kW system. Maximum total impulse values and 

maximum delta-v’s on the DELTA I1 6920 range from 1.2 x IO’ to 2.1 x lo’ N-s, awl 3550 to 6200 m/s, 

respectively. Ion propulsion system and spacecraft mission parameters obtained with launches 011 the ATLAS 

I, DELTA I1 7920, and ATLAS I1 vehicles are also listed in Table VIII. 

Since the specific impulse of the arcjet propulsion options is substantially lower than that of the ion 

systeiiis, the launch vehicle payload ni rrgins are also much lower. Maximi ni available thrusting times 

for the arcjet propelled spacecraft on the DELTA I1 6920 vehicle range from approximately 528 hours for 

the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maxiiiiuiii total 

impulse values and maximum delta-v’s on the DELTA I1 6920 range froin 2.2 x 10‘ to 3.6 x lo6 N-s, and 

approximately 662 to 1072 m/s, respectively. These values of total impulse and delta-v are approxiniately a 

factor of 5 to 6 lower than that obtained with the ion systems for the same payload and launch vehicle. It 

is noted that only three arcjet options provide total thrusting times in excess of 5000 hours. These are the 

single-engine 5 kW system on the ATLAS I (5736 hours), DELTA I1 7920 (6096 hours), and the ATLAS I1 

(7992 hours) launch vehicles. 

The highest performance in terms of maximum delta-v for both the ion and arcjet systems are obtained 

6 
with the 10 kW single-engine options. Figures 10 and 11 show the maximum thrusting time versus launch 

vehicle option, and maximum delta-v versus launch vehicle option for these systems. As indicated in Figure 

10, the maximum thrusting time for the single-engine 10 kW ion-propelled spacecraft ranges from approximate 17,400 hours on the DELTA I1 6920 to 64,200 hours on the ATLAS 11. Corresponding values for 

t.lie single-engine 10 kW arcjet spacecraft range from approximately 1248 hours to 4368 hours. Maximum 

delta-v's for this ion-propelled spacecraft range from 6.2 km/s on the DELTA I1 to 18.6 km/s on the ATLAS 

11. Corresponding values for the single-engine 10 kW arcjet spacecraft range from 1.1 km/s to 3.0 km/s. 

Mission Mode I1 - In defining a total thrusting time for the EP systems (5000 hours for ion/1000 hours for 

arcjet), a delta-v, total impulse, and initial spacecraft mass can be determined. Tables IX (a) and (b) list 

the propulsion system and mission parameters calculated for this mission mode using the various ion and 

arcjet propulsion system options. 

Under the constraint of a fixed thrusting time, the maximum spacecraft delta-v and total impulse are 

obtained with the 4-engine 10 kW xenon ion propulsion system, and (for arcjet) the 6-engine 10 kW hydrazine 

arcjet propulsion system. The xenon ion system provides a 2121 m/s delta-v capability, at a total impulse 

value of 7.1 x 10' N-s and total mass of 466 kg for a 5000 hour total burn-time. The hydrazine arcjet system 

provides a 1161 ni/s delta-v capability, at a total impulse value of 4.3 x 10' N-s and total mass of 1161 kg 

for a 1000 hour total burn-time. 

The propulsion system dry masses for the ion options are 2 to 3 times more massive than the arcjet 

options as previously indicated. However, the total EP system mass (including the propellant and tankage) 

of the arcjet systems is approximately a factor of 2 higher than that of the ion systems at the same power 

level and number of thrusters because of the lower values of specific impulse. The 10 kW total ion propulsion 

system masses range from 466 kg for the 4-engine option, to 257 kg for the single-engine option. The 10 kW 

hydrazine arcjet system masses range from 1161 kg for the 6-engine option to 536 kg for the single-engine 

option. 

Table IX also lists the total spacecraft mass for all the EP system options, and the positive payload 

margin available on the four proposed launch vehicles with these spacecraft. The total spacecraft masses are 

most massive with the arcjet propulsion system options; consequently the payload margins are lower than 

with the ion system options. The payload margin with the 6-engine 10 kW arcjet propelled spacecraft on the 

DELTA I1 6290 is negative, which indicates that this launch vehicle option is not available. To accomplish 

the baseline nlission scenario of a 1000 hour thrusting-time would require that this system be launched on a 

heavier lift vehicle such as the ATLAS I. Positive payload margins for the arcjet propelled spacecraft range 

from a low of 57 kg on the DELTA I1 6920 to a maximum of 1848 kg on the ATLAS 11. All the payload 

iiiargins for the ion propelled spacecraft are positive, ranging from a low of 184 kg on the DELTA I1 6920 to 

a high of 1975 kg on the ATLAS 11. That is, the baseline mission scenario (5000 hour total thrusting time) 

could be accomplished with the DELTA I1 6920 for any of the ion system options. 

Figures 12 and 13 show the spacecraft mass distribution for the 2-engine 10 kW ion and 6-engine hydrazine 

arcjet system options, respectively. As indicated in figure 12, the total ion propulsion system mass (including 

propellant and tankage) represents only 10% of the 3288 kg spacecraft mass. The propellant and tankage 

for the ion system is more than 55% of the total EP mass. The arcjet propulsion system mass (see figure 13) 

represents more than 28% of the 4111 kg total spacecraft mass, with the propellant and tankage constituting 

nearly 93% of the total EP mass. 

CONCLUDING REMARKS 

Ion and arcjet propulsion system options for a near-term space demonstrztion experiment were identified, 

and analyses were conducted to determine first-order propulsion system performance, and system component 

mass estimates. Overall mission performance of the EP systems was quantified in terms of total impulse, 

delta-v, and thrusting time capability based on a mission model which assumed a 2950 kg payload mass 

(including power) and either 5 or 10 kW available from a solar array to the EP system. 

The thirty centimeter diameter ring-cusp engine operating on xenon propellant was baselined as the ion 

thruster option. This thruster can run from approximately 2450 seconds to 4520 seconds specific impulse, 

which Corresponds to 62 to 79% overall thruster efficiency. Thrust levels range from approximately 0.04 to 

0.42 Newtons, which corresponds to thruster input power levels from approximately 1 to 12 kW. Although 

ion thruster lifetime at high power levels has not been verified, it is believed a 35,000 A-11 screen grid 

(beam current times operating time) lifetime limit at 28 V discharge voltage with xenon propellant is a 

7 
conservative measure of thruster life. Ion thruster Lifetimes ranging from approximately 5,800 to 35,000 

hours, corresponding to the range of input power levels from 12 kW down to 1 kW, would be anticipated. 

Both ammonia and hydrazine propellants were assumed for the arcjet. Thruster performance estimates on 

hydrazine range from 523 seconds to 800 seconds specific impulse, corresponding to 34% thruster efficiency. 

Thrust levels with hydrazine were estimated to be in the range of 0.20 to 0.81 Newtons for input power levels 

from approximately 1.5 to 9.4 kW. 

The ion and arcjet propulsion systems were derived using comparable propulsion system mass and mission 

models. The system inass model used defined the propulsion system into a thrust module and an interface 

module. The mission model assumed that a low power ion or arcjet propulsion system would fly on a 

dedicated electric propulsion space experiment. Either 5 or 10 kW was assumed to be available from a 

solar array on the sapceuaft to the ion or arcjet propulsion module. Two mission modes were assumed; an 

extended-duration thrusting time dependent upon the launch vehicle payload margin, and a thrusting time 

denionstration of 5000 hours for the ion systems and 1000 hours for the arcjet systems. A DELTA I1 6920, 

ATLAS I, DELTA I1 7920, or ALTAS I1 launch vehicle deployment of the spacecraft to 550 km circular orbit 

was assumed. 

Several Candidate 5 and 10 kW ion and arcjet propulsion system options were identified. The majority 

of dry mass of the ion systems was in the power processor, which constituted greater than 40% of the total 

dry mass for all the system options. The arcjet propulsion system dry masses ranged from a maximum of 87 

kg for a 6-engine 10 kW system, down to a minimum of 38 kg for a single-engine 5 kW system. For all of the 

arcjet systems, the power processor was the most massive system component, constituting anywhere from 

28 to 34% of the total dry mass. The second most massive component was the dry interface module, except 

for the 6- and 2-engine 10 kW arcjet systems where the thermal control was the second heaviest coniponent 

system component. 

Values for total initial dry spacecraft mass for each combination of electric propulsion system option 

and payload were determined. This was done by individually adding each EP system dry mass to the 2950 

kg payload and subtracting this mass value from the mass lift capability of the launch vehicle to define a 

payload margin. The payload margin, in part or in whole, was then consumed by the addition of propellant 

and tankage for the EP system. 

The thrusting times and delta-v’s could be determined by on-loading propellant to the spacecraft to fill 

the launch vehicle payload margin. Maximum available thrusting times for the ion-propelled spacecraft on 

the DELTA I1 6920 vehicle range from approximately 8,600 hours for the 4-engine 10 kW system to more 

than 29,600 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v’s 

on the DELTA I1 6920 range from 1.2 x lo7 to 2.1 x 10’ N-s, and 3550 to 6200 m/s, respectively. Enough 

propellant could be on-loaded to most of the ion-propelled spacecraft to permit thruster operation greatly 

in excess of the projected thruster lifetimes. 

Since the arcjet total propulsion system masses were substantially higher than that of the ion systems, 

the launch vehicle payload margins were much lower. Maximum available thrusting times for the arcjet 

propelled Spacecraft on the DELTA I1 6920 vehicle range from approximately 528 hours for the 6-engine 10 

kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and 

maximum delta-v’s on the DELTA I1 6920 range from 2.2 x 10‘ to 3.6 x lo6 N-s, and approximately 662 to 

1072 m/s, respectively. These values of total impulse and delta-v are approximately a factor of 5 to 6 lower 

than that obtained with the ion systems for the same payload and launch vehicle. Only three arcjet options 

provided total thrusting times in excess of 5000 hours. These were the single-engine 5 kW system on the 

ATLAS I (5736 hours), DELTA I1 7920 (6096 hours), and the ATLAS I1 (7992 hours) launch vehicles. 

The highest performance in terms of maximum delta-v for both the ion and arcjet systems were obtained 

with the 10 kW single-engine options. Maximum delta-v’s for the single-engine ion-propelled spacecraft range 

from 6.2 km/s on the DELTA I1 to 18.6 km/s on the ATLAS 11. Corresponding values for the single-engine 

10 kW arcjet spacecraft range from 1.1 km/s to 3.0 km/s. 

Under the constraint of a fixed thrusting time, the maximum spacecraft delta-v and total impulse were 

obtained with the 4-engine 10 kW xenon ion propulsion system, and (for arcjet) the 6-engine 10 kW hydrazine 

propulsion system. The xenon ion system provides a 2121 m/s delta-v capability, at a total impulse value 

of 7.1 x 10‘ N-s and total mass of 466 kg for a 5000 hour total burn-time. The hydrazine arcjet system 

provides a 1161 m/s delta-v capability, at a total impulse value of 4.3 x 10‘ N-s and total mass of 1161 kg 

for a 1000 hour total burn-time. The baseline mission scenario (5000 hour total thrusting time) could be 

8 
accomplished with the DELTA I1 6920 for all of the ion system options. However, to accomplish a 1000 hour 

thrusting time would require that the 6-engine 10 kW arcjet systems be launched on a heavier lift vehicle 

such as the ATLAS I. 

REFERENCES 

1. Deininger, W.D.; and Vondra, R.J.: Electric Propulsion for Constellation Deployment and Spacecraft 

Maneuvering. AIAA Paper 88-2833, July 1988. 

2. Sponable, J.M.; and Penn, J.P.: An Electric Orbital Transfer Vehicle for Delivery of NAVSTAR 

Satellites. AIAA Paper 87-0985, May 1987. 

3. Sovey, J.S.: Iiiiproved Ion Containment Using a Ring-Cusp Ion Thruster. Journal of Spacecraft and 

Rockets, 21, No. 5, pp. 488-495, September-October 1984. 

4. Patterson, M.J.;and Rawlin, V.K.: Performance of 10-kW Class Xenon Ion Thrusters. AIAA Paper 

88-2914, July 1988. 

5. Rawlin, V.K.: Internal Erosion Rates of a 10 kW Xenon Ion Thruster. AIAA Paper 88-2912, July 

1988. 

6. Rawlin, V.K.: Operation of the J-Series thruster Using Inert Gas. AIAA Paper 82-1929, November 

1982. 

7. Patterson, M.J.: Performance Characteristics of Ring-Cusp Thrusters with Xenon Propellant. AIAA 

Paper 86-1392, June 1986. 

8. Rawlin, V.K.; and Patterson, M.J.: High Power Ion Thruster Performance. Presented at the Fourth 

Space Nuclear Power Conference, January 1987. 

9. Beattie, J.R.; Matossian, J.N.; and Robson, R.R.: Status of Xenon Ion Propulsion Technology. AIAA 

Paper 87-1003, May 1987. 

10. Beattie, J.R.; Matossian, J.N.; Poeschel, R.L.; Rogers, W.P.; and Martinelli, R: Xenon Ion Propulsion 

Subsystem. AIAA Paper 85-2012, September 1985. 

11. Byers, D.C.; Terdan, F.F.; and Myers, I.T.: Primary Electric Propulsion for Future Space Missions. 

AIAA Paper 79-0881, May 1979. 

12. Galecki, D.L.; Patterson, M.J.: Nuclear Powered Mars Cargo Transport Mission Utilizing Advanced 

Ion Propulsion. AIAA Paper 87-1903, July 1987. 

13. Curran, F.M.; and Haag, T.M.: An Extended Life and Performance Test of a Low Power Arcjet. AIAA 

Paper 88-3106, July 1988 (NASA TM-100942). 

14. Knowles, S.K.: Arcjet Research and Technology, Phase 11, Final Report. NASA CR-182276, Rocket 

Research Co., Redniond, WA (to be published). 

15. Zana, L.M.: Langmuir Probe Surveys of an Arcjet Exhaust. AIAA Paper 87-1950, July 1987. (NASA 

TM-89924). 

16. Carney, L.M.: Evaluativm of the Communications Impact of a Low Power Arcjet Thruster. AIAA 

Paper 88-3105, July 1988. (NASA TM-100926). 

17. Pivirotto, T.J.; King, D.Q.; and Deininger, W.D.: Long Duration Test of a 30-kW Class Tlieriiial 

Arcjet Engine. AIAA Paper 87-1947, July 1987. 

18. Deininger, W.D.; Chopra, A.; King, D.Q.; and Pivirotto, T.J.: Thermal Design Improvements for 30 

kWe Arcjet Engines. IEPC Paper 88-073, October 1988. 

19. Hardy, T.L.; Rawlin, V.K.; and Patterson, M.J.: Electric Propulsion Options for the SP-100 Reference 

Mission. NASA TM-88918, January 1987. 

20. Penko, P.F.; Staiger, P.J.; and Burr, M.J.: An Analysis of Low-Thrust, Resistojet Reboost for the 

Space Station. AIAA Paper 85-2042, October 1985. 

9 
21. Mission Planner's Guide for the Atlas Launch Vehicle Family. Revision 1, March 1989. General 

Dynamics commercial Launch Services, Inc. 

22. Delta I1 Commercial Spacecraft Users Manual. July 1987. McDonnell Douglas Corporation. 

APPENDIX A 

The assumptions used in generating predicted 30 cm ion thruster performance on xenon propellant were 

as follows: 

1. The beam current (Jb) was calculated from an empirical equation which predicts the performance of 30 

cm small-hole-accelerator-grid (SHAG) 2-grid ion optics on a laboratory model ring-cusp ion thruster 

with xenon propellant. This expression is given by 

where A is t,he effective beam area, M is the propellant atomic mass unit, and Vt is the total accelerating 

voltage. This equation predicts 85%-of-maximum beam current attained experimentally for the optics 

described, at a cold-gap spacing of approximately 0.66 mm. 

2. The total accelerating voltage across the optics ranged from 1000 to 2000 volts, inclusive. Two-grid 

SHAG optics were assumed, with an effective range of net-to-total voltage (R) of 0.55 to 0.85, inclusive. 

The maximum total accelerating voltage attainable was assumed to be 2500 V at 0.66 mm gap, or 

equivalently, 3800 V/mm maximum electric field strength. 

3. A total thrust loss due to beam divergence (0.98, neglecting dependence on R-ratio) plus niultiplycharged ions was estimated. The value of multiply-charged ions (J++/J+) was estimated based on a 

simple curve-fit of data obtained from 30 cm ion thruster beams, given by 

(2) 

J++ 

~ = -7.1078 * N + 15.3317 * N2 - 14.7856 * N3 + 5.4228. N' + 1.2439 J+ 

where N is the discharge chamber propellant efficiency. 

4. The maximum discharge chamber propellant efficiency for xenon was assumed to be 0.95. The beam 

ion production cost was assumed to be 125 W/A. The discharge voltage was assumed to be 28 volts. 

5. A fixed power loss of 0.050 kW was assumed. 

6. A neutralizer mass flow rate of 3.2% of the beam current was assumed. 

7. A thruster thermal limit of 800 watts maximum discharge power was specified to preclude the rareearth cobalt magnets in the discharge chamber from exceeding 300 "C where irreversible losses in field 

st.re1igt.h may occur. 

APPENDIX B 

The modifications, from reference 19, employed in the low power arcjt t propulsion system modeling 

include the following: 

1. The arcjet mass, including gas generator, propellant valve, and mounting structure, was assumed to 

be a funct,ion of input power to account for changes in required radiating surface area. The systein 

modeling identified 3 candidate power levels - -1.5 kW, -5 kW, and -10 kW per thruster/ppu. The 

present 1 kW NASA-LeRC arcjetI3 is capable of operation up to approximately 2 kW. Consequently, 

this point design of 0.84 kg was used to estimate the 1.5 kW thruster mass. The thruster masses for 

the 5 and 10 kW power levels were estimated to be 1.70 and 2.95 kg, respectively. 

2. The arcjet discharge supply mass (Md,) in the power processor (in kilograms) was scaled based on the 

following equation 

10 
Md. = 0.455 + 4.095. W6 (3) 

where P is the input power into the power processor in kW. This equation is based on a single point 

design of a 1.4 kW flight-type arcjet power processor14. 

3. A power processor efficiency of 90% for ppu input power levels of -1.7 kW was assumed. The power 

processor efficiency was assumed to be 93% at -5 kW and 94% at -10 kW. Line losses of 0.5% from 

the solar array to the ppu were assumed. 

11 
Table I. Propulsion System Model Description 

Thruster 

Efficiency V,, V R ratio 

THRUST MODULE 

-- Thruster 

-- Gimbal (fmed fraction of thruster mass) 

-- Thermal Control (Power Processor only) 

-- Power Processor (PPU) 

-- Thruster Structure 

-- Propellant dismbution 

Mass flow 

kgh 

INTERFACE MODULE 

1050 

1150 

1325 

1450 

1650 

2050 

-- Thrust System Controller 

-- Converter 

-- Reconfiguration Unit 

-- Thermal Control 

-- Housing structure 

.80 

31 

.78 

.79 

.79 

.79 

-- Propellant 

-- Tankage 

-- Tankage Structure 

"Dry" Interface Module Mass 

Propellant Module Mass 

I 

Table II. Projected 30 cm Xenon Ion Thruster Performance 

Beam Current, A 

1.13 

1.39 

1.91 

2.35 

3.15 

5.15 

Thrust, N 

.052 

.067 

.098 

.126 

.180 

.329 

1.14 

1.52 

2.27 

3.03 

4.55 

9.04 

3179 

3347 

3526 

3712 

3960 

4414 

.7 1 

.73 

.75 

.76 

.77 

.79 

, 1.66e-6 

I 2.05e-6 

2.82e-6 

3.46e-6 

4.w-6 

7.e-6 

12 
Table III. Xenon Ion Propulsion System Options 

Number of Thrusters 

Specific Impulse, s 

Total Thrust, N 

System Efficiency 

Dry Masskg 

Specific Mass,kg/kW 

10 I 5 I I Total Power, kW 

4 

3179 

.208 

.65 

208 

41.6 

4 

3526 3712 

.392 .378 

.69 .70 

234 191 

23.4 19.1 

32 1 

3960 4414 

.360 .329 

.70 .72 

147 101 

14.7 10.1 

I 10 I 5 I I Total Power, kW 

Thrust, N 

Table V. Projected Arcjet Thruster Performance 

Thruster Mass flow 

Input Power, kW Isp, s Eficiency kgls 

.195 

.428 

.763 

1.49 500 .32 3.97e-5 

4.63 750 .34 5.81e-5 

9.35 900 .36 8.&-5 

13 

.I98 

.423 

.810 

1.49 523 .34 3.86e-5 

4.63 758 .34 5.69e-5 

9.35 800 .34 1.03e-4 
Table VI. Arcjet Propulsion System Options 

500 750 

584 .428 

.29 .31 

49 38 

9.8 7.6 

10 I 

3131 

523 758 

.593 .423 

.30 .31 

49 38 

9.8 7.6 

Total Power, kW 5 

Propellant "3 I N2H4 "3 

- 

N2H4 

1 

900 

.763 

.34 

57 

5.7 

- 

6 

500 

1.168 

.29 

87 

8.7 

- 

2 

750 

.855 

.3 1 

65 

6.5 

- 

6 

523 

1.187 

.30 

87 

8.7 

- 

2 

758 

.846 

.31 

65 

6.5 

Number of Thrusters 

Specific Impulse, s 

Total Thrust, N 

System Efficiency 

Dry Masskg 

Specific Masskg/kW 

Table Vn. Arcjet Ropulsion System Component Masses 

Total Power, kW 5 10 

- 2 

4.6 

18.8 

21.0 

1.4 

19.0 

- 

- 1 

2.3 

9.4 

10.5 

0.7 

15.0 

- 

- 6 

6.8 

26.9 

29.2 

2.1 

22.0 

Number of Thrusters I 

Thruster/Gimbal,kg 

Thermal Contro1,kg 

Power Prwessor,kg 

Thruster Structure,kg 

Dry Interface Modulekg 

3 

3.4 

13.4 

14.6 

1.0 

16.0 

4.0 

16.1 

17.4 

1.2 

18.0 

Total Dry Mass,kg -49 -38 

- 

-87 

- 

-57 

14 
Table WI. Ion and Arcjet Propulsion System and Mission Parameters for Extended-Duration Mission 

(a). Xenon Ion Propulsion 

(Mission Mode I - -Zero Launch Vehicle Payload Margin) 

System and Mission 

Parameters 

Launch Vehicle Option 

Delta I1 6920 

5 10 

No. of Thrusters 

EP System 

Design 

4 3 2 1 4 3 2 1 

Initial Thrust/Weight 

Max. Thrust Time, days 

Max. Total Impulse, N-s 

Max. A V, m/s 

EP System Dry Mass, kg 

Xenoflankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

Total Initial S/C Mass, kg 

5.78e-5 

674 

1.21e7 

3550 

208 

442 

650 

2950 

3600 

5.61e-5 

768 

1.33e7 

4000 

166 

467 

633 

2950 

3583 

5.47e-5 

909 

1.54e7 

4600 

123 

507 

630 

2950 

3580 

5.W-5 

1237 

1.92e7 

5800 

79 

567 

646 

2950 

3596 

1.09e-4 

361 

1.22e7 

3600 

234 

403 

637 

2950 

3587 

I .05e-4 

433 

1.41e7 

4200 

191 

444 

635 

2950 

3585 

1 .W-4 

538 

1.67e7 

5ooc 

147 

493 

640 

2950 

3590 

9.15e-5 

727 

2.07e7 

62M 

101 

546 

647 

2950 

3597 

System and Mission 

Parameters 

Launch Vehicle Option 

Atlas I 

5 10 

Design 

3 2 

1.37e-5 

2526 

4.28e7 

11200 

123 

1408 

1531 

2950 

448 1 

- 

1 

1.02e-5 

3156 

4.91e7 

13000 

79 

1447 

1526 

2950 

4476 

- 

4 

8.74e-4 

1167 

3.95e7 

10200 

234 

1301 

1535 

2950 

4485 

- 

3 

8.4 1 e-4 

1318 

4.3oe7 

11200 

191 

1352 

1543 

2950 

4493 

- 

2 

8 .O 1 e- 5 

1522 

4.73e7 

12400 

147 

1396 

1543 

2950 

4493 

- 

1 

7.31e-5 

1928 

5.48e7 

14400 

101 

1448 

1549 

2950 

4499 

~ 

Initial ThrusWeight 

Max. Thrust Time, days 

Max. Total Impulse, N-s 

Max. A V, ds 

EP System Dry Mass, kg 

Kenoflankage Mass, kg 

rota1 EP System Mass, kg 

Payload Mass, kg 

Total Initial S/C Mass, kg 

4.62e-5 

2047 

3.68e7 

9450 

208 

1344 

1550 

2950 

4500 

4.5oe-5 

223 1 

3.87e7 

10200 

166 

1356 

1522 

2950 

4472 

15 
System and Mission 

Parameters 

4 

8.5%-5 

1251 

4.24e7 

10800 

234 

1395 

1629 

2950 

4519 

No. of Thrusters 

EP System 

Design 

3 

8.25e-5 

1405 

4.59e7 

11800 

191 

1441 

1632 

2950 

4582 

Initial ThrustlWeight 

Max. Thrust Time, days 

Max. Total Impulse, N-s 

Max. A V, m/s 

EP System Dry Mass, kg 

Xenoflankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

Total Initial S/C Mass, kg 

System and Mission 

Parameters 

EP System 

Design 

[nitial Thrusweight 

Max. Thrust Time, days 

Max. Total Impulse, N-s 

Max. A V, m/s 

EP System Dry Mass, kg 

Xenoflankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

rotal Initial S/C Mass, kg 

Launch Vehicle Option 

Delta I1 7920 

4 

4.52e-5 

2195 

3.94e7 

loo00 

208 

1440 

1648 

2950 

4598 

3 

4.4oe-5 

2394 

4.16e7 

10800 

166 

1455 

1621 

2950 

457 1 

5 

1 

4.28e-5 

2695 

4.56e7 

11800 

123 

1502 

1625 

2950 

4575 

1 

3.9%-5 

340 1 

5.29e7 

13800 

79 

1560 

1639 

2950 

4589 

- 

10 - 

2 

7.87e-5 

1614 

5.02e7 

13000 

147 

1480 

1627 

2950 

4571 

- 

- 

1 

7.15e-5 

2062 

5.86e7 

15200 

101 

1549 

1650 

2950 

4600 

- 

- 

Launch Vehicle Option 

Ai 

4 

1.08e-5 

2958 

5.32e7 

12650 

208 

1941 

2149 

2950 

5099 

3 

3.97e-5 

3210 

5.57el 

13600 

166 

1951 

21 17 

2950 

5067 

2 

3.86e-5 

3605 

6.1Oe7 

14800 

123 

2010 

2133 

2950 

5083 

1 

l.55e-5 

4441 

6.91e7 

17000 

79 

2037 

21 16 

2950 

5066 

- 

4 

'.7le-5 

1704 

5.77e7 

13800 

234 

1900 

2134 

2950 

5084 

- 3 

7.42e-5 

1906 

6.22e7 

15000 

191 

1955 

2146 

2950 

- 5096 

L 

2 

7.07e-5 

2172 

6.76e7 

16400 

147 

1992 

2139 

2950 

5089 

1 

6.5Oe-5 

2677 

7.61e7 

18600 

101 

201 1 

2112 

2950 

5062 

16 

I 
(b). Hydrazine Arcjet Propulsion 

EP System 

System and Mission 

Parameters 

Power 

Initial Thrusweight 

Max. Thrust Time, days 

Max. Total Impulse, N 

Max. A V, m/s 

EP System Dry Mass, kg 

Hydrazinemankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

Total Initial S/C Mass, kg 

System and Mission 

Parameters 

~ 

Design No. of Thruster! 

Initial ThrusWeight 

Max. Thrust Time, days 

Max. Total Impul 

Max. A V, m/s 

EP System Dry Mass, kg 

Hydrazinemankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

Total Initial S/C Mass, kg 

Launch Vehicle Option 

Delta I1 6920 

5 I 10 

3 

1.65e-4 

47 

2.38e6 

710 

49 

599 

648 

2950 

3598 

~~ 

1 

1.18e-4 

97 

3.53e6 

1051 

38 

612 

650 

2950 

3600 

6 

3.3oe-4 

22 

2.24e6 

662 

87 

56 1 

648 

2950 

3598 

2 

2.35e-4 

46 

3.37e6 

1002 

65 

585 

650 

2950 

3600 

1 

!.25e-4 

52 

3.61e6 

1072 

57 

592 

649 

2950 

3599 

Launch Vehicle Option 

3 

L.32e-4 

116 

5.96e6 

1536 

49 

1500 

1549 

2950 

4499 

- 

5 

1 

9.4oe-5 

239 

8.72e6 

2245 

38 

1512 

1550 

2950 

4500 

Atlas I 

6 

!.64e-4 

57 

5.81e6 

1490 

87 

1461 

1548 

2950 

4498 

- 

11 

10 

2 

1.88e-4 

117 

8.55e6 

2198 

65 

1484 

1549 

2950 

4499 

1 

1.8Oe-4 

130 

9.08e6 

2334 

57 

1492 

1549 

2950 

4499 
System and Mission 

Parameters 

EP System Power 

Initial ThxusWeight 

Max. Thrust Time, days 

Max. Total Impulse, N-s 

Max. A V, m/s 

EP System Dry Mass, kg 

Hydrazineflankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

Total Initial S/C Mass, kg 

System and Mission 

Parameters 

Initial ThrusWeight 

Max. Thxust Time, days 

Max. Total Impulse, N-s 

Max. A V, m/s 

EP System Dry Mass, kg 

Hydrazineflankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

Total Initial SIC Mass, kg 

Launch Vehicle Option 

Delta I1 7920 

3 

1.29e-4 

124 

6.36e6 

1614 

49 

1600 

1649 

2950 

4599 

5 

1 

9.2Oe-4 

254 

9.29e6 

2358 

38 

1612 

1650 

2950 

4600 

- 

6 

2.58e-4 

61 

6.22e6 

1570 

87 

1563 

1650 

2950 

4600 

- 

10 

2 

1.84e-4 

125 

9.14e6 

2312 

65 

1584 

1649 

2950 

4599 

1 

1.76e-4 

138 

9.69e6 

2454 

57 

1592 

1649 

2950 

4599 

Launch Vehicle Option 

3 

1.16e-4 

163 

8.35e6 

1976 

49 

2101 

2050 

2950 

5100 

5 

1 

8.3Oe-5 

333 

1.22ei 

288( 

38 

2110 

2148 

2950 

5098 

Atlas II 

- 

6 

!.33e-4 

80 

8.2Oe6 

1932 

87 

2062 

2149 

2950 

5099 

- 

10 

2 

1.66e-4 

164 

1.2Oe7 

2836 

65 

2084 

2149 

2950 

5099 

- 

1 

1 59e-4 

182 

1.27e6 

3008 

57 

2093 

2150 

2950 

5100 

18 
Table IX. Ion and Arcjet Propulsion Systcm and Mission Parametm fur Mission with Specified Thrust Tim 

(a). Xenon Ion Propulsion 

(Mission Mode II) 

3526 

6.le-5 

208.3 

3.5e6 

1120 

123 

116 

239 

2950 

3189 

41 1 

1311 

1411 

1911 

Total Power, kW 

3960 

5.8e-5 

208.3 

3.2e6 

1053 

79 

96 

175 

2950 

3125 

475 

1375 

1475 

1975 

Number of Thrusters 

Isp,s 

Initial Thrust/Weight 

Thrust Time,,days 

Total Impulse, N-s 

A v, 4s 

EP System Dry Mass, kg 

Xenoflankage Mass, kg 

rota1 EP Mass, kg 

Payload Mass, kg 

rota1 SIC Mass, kg 

Launch Vehicle Payload 

Margin, kg 

Delta II 6920 

4 

3179 

6.3e-5 

208.3 

3.7e6 

1155 

208 

137 

345 

2950 

3295 

305 

1205 

1305 

1805 

- 

3 

3347 

6.2e-5 

208.3 

3.6e6 

1138 

166 

126 

292 

2950 

3242 

358 

1258 

1358 

1858 

5 

2I 4 

3526 

l.le-4 

208.3 

7.le6 

2121 

234 

232 

466 

2950 

3416 

184 

1084 

1184 

1684 

3 

3712 

l.le-4 

208.3 

6.8e6 

2088 

191 

214 

405 

2950 

3355 

245 

1145 

1245 

1745 

10 

2 

3960 

l.le-4 

208.3 

6.5e6 

2026 

147 

191 

338 

2950 

3288 

312 

1212 

1312 

1812 

1 

4414 

L.oe-4 

208.3 

5.9e6 

1887 

101 

156 

257 

2950 

3207 

393 

1293 

1393 

1893 
(b). Hydrazine Arcjet Propulsion 

~~ 

Number of Thrusters 

Isp, s 

Initial Thrust/Weight 

Thrusting Time, days 

Total Impulse, N-s 

A V, ds 

EP System Dry Mass, kg 

Hydrazineflankage Mass, kg 

Total EP System Mass, kg 

Payload Mass, kg 

Total Initial S/C Mass, kg 

Launch Vehicle Payload Margin, kg 

Delta II 6920 

Atlas I 

Delta II 7920 

Atlas I1 

3 

523 

1.7e-4 

41.7 

2.1e6 

643 

49 

537 

586 

2950 

3536 

64 

964 

1064 

1564 

5 

1 

758 

1.3e-4 

41.7 

1 Se6 

484 

38 

264 

302 

2950 

3252 ~ 

348 

1248 

1348 

1848 

- 

6 

523 

2.9e-4 

41.7 

4.3e6 

1161 

87 

1074 

1161 

2950 

- 4111 

-51 1 

389 

489 

989 

- 

10 

2 

758 

2.k-4 

41.7 

3.Oe6 

914 

65 

528 

593 

2950 

- 3543 

57 

957 

1057 

1557 

- 

1 

800 

!.3e-4 

41.7 

2.9e6 

885 

57 

479 

536 

2950 

3486 

114 

1014 

1114 

1614 

Figure 1. - Cross section of ring-cusp ion thruster. 

20 
"." 

* 

0 

c 

P) 

0 

5: 

i5 0.7 

- 

L 

L 

(I) 

a 

c 

IL 

2000 3000 4000 5000 

Specific Impulse, seconds 

Figure 2. - Ion thruster efficiency versus specific impulse with xenon propellant. 

0.5 

I 1000 v, 1'A 

electric field strength 

/limit 

2000 3000 4000 5000 

Specific Impulse, seconds 

Figure 3. - Ion thruster thrust versus specific impulse. 

Approximate operating envelope for 30 crn ring-cusp xenon ion thruster technology. 

21 
LC = constrictor length = 0.26 mm . e--. 

Dc = constrictor diameter 0.84 mm 

\ 

LPRWELLANT LIE IN-LIE 

ORIFICE 

Figure 4. - Cutaway view of arcjet thruster with typical dimensions. 

0 PRE-TEST MTA 

0 DATA AFTER 144 CYCLES 

0 DATA KTER 335 CYCLES 

A 

20 22 24 26 28 30 32 34 

pni,(-)x 10-3 

Figure 5. - Performance measurements before, during, and after a 1000 hr/500 

cycle automated arcjet lifetest. 

22 
300 

200 

100 

0 

3000 3500 4000 4500 

Specific Impulse, seconds 

Figure 6. - Ion propulsion system dry mass versus specific impulse. 

Number of thrusters is indicated. 

0 5kW 

0 lOkW 

4-Engine Sys 3-Engine Sys 2-Engine Sys 1-Engine Sys 

Figure 7. (a) - Dry mass distribution of 5 kW ion propulsion systems. 

Thruster structure 

Dry interface module 

Thermal control 

Thruster/gimbal 

tn Power processor 

23 
300 

200 

100 

0 

80 - 

70 - 

60 - 

50 - 

40 - 

T 

11911 

Thruster structure 

Dry interface module 

Thermal control 

Thruster/gimbal 

Power processor 

4-Engine Sys 3-Engine Sys 2-Engine Sys 1-Engine Sys 

Figure 7. (b) - Dry mass distribution of 10 kW ion propulsion systems. 

6 

O0 3 

m2 

0 

01 

1' 

30:. I., . I. I., . 400 500 600 700 800 900 1000 

Specific Impulse, seconds 

0 N2H4,SkW 

N2H4, lOkW 

0 NH3,5kW 

"3, lOkW 

Figure 8. - Arcjet bropulsion system dry masses versus specific impulse. 

Number of thrusters is indicated. 

24 
I 

100 

LI) 

Y 80 

60 - 

40 - 

20 - 

O+ -r T T 

(49) 

-r 

611OkW 2110kW l/lOkW 315kW 1/5kW 

Figure 9. - Dry mass distribution of arcjet propulsion systems. 

3000 

2000 

1000 

0 

(1 928) 

(2677) 

120621 

.r 

DELTA I1 6920 ATLAS I DELTA II 7920 ATLAS II 

Thruster structure 

Dry interface module 

Thermal control 

Thruster/girnbal 

Power processor 

1 OkW 1 -engine Xe ion 

1 OkW 1 -engine N2H4 

arcjet 

Figure 10. - Max. thrusting time vs. launch vehicle option for extended-duration mission. 

Mission Mode I 

25 
20 

118.6) I 

u 

Q 

a 

z 

e 

Y 

0 

Q 10 - - c 

n 

E 

- E' 

X 

I" 

(14.4) (1 5.2) 

1 OkW 1 -engine Xe ion 

1 OkW 1 -engine N2H4 

arcjet 

T l- lDELTA II 6920 ATLAS I DELTA II 7920 ATLAS II 

Figure 11. - Max. delta-v vs. launch vehicle option for extended-duration mission. 

Mission Mode I 

Total propulsion system mass 338 kg. 

PayloadIpower 

PropellanVtankage 

Thruster/gimbal 

0 Dry int. module/ 

Thermal controll 

Thruster structure 

f3m 

Figure 12. - Spacecraft mass distribution with 2-engine 10kW Xe ion option. 

5000 hour thrusting time. Total s/c mass 3288 kg. 

Mission Mode II 

26 
Total DroDulsion svstern mass 1 161 ka 

Paybadlpower 

PropellanVtankage 

WU 

Thermal control 

Dry int. module/ 

ThrusterlgimbaV 

Thruster structure 

Figure 13. - Spacecraft mass distribution with 6-engine lOkW N2H4 arcjet option. 

1000 hour thrusting time. Total s/c mass 41 11 kg. 

Mission Mode It 

21 
National Aeronautics and 

Space Administration 

1. Report No. 

NASA TM-102337 

Report Documentation Page 

2. Government Accession No. 

M.J. Patterson and F.M. Curran 

9. Performing Organization Name and Address 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 

12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, D.C. 20546-0001 

5. Supplementary Notes 

3. Recipient’s Catalog No. 

5. Report Date 

6. Performing Organization Code 

8. Performing Organization Report No. 

E-5046 

10. Work Unit No. 

506-42-3 1 

11. Contract or Grant No. 

13. Type of Report and Period Covered 

Technical Memorandum 

14. Sponsoring Agency Code 

Prepared for the 1989 JANNAF Propulsion Meeting, Cleveland, Ohio, May 23-25, 1989. 

6. Abstract 

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have 

been evaluated. Analyses were conducted to determine first-order propulsion system performance and system 

component mass estimates. Overall mission performance of the electric propulsion systems was quantified in 

terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated 

onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ionpropelled spacecraft options, launched on a DELTA I1 6920 vehicle, range from approximately 8,600 hours for a 

4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse 

values and maximum delta-v’s range from 1.2~10~ to 2.1~10’ N-s, and 3550 to 6200 m/s, respectively. 

Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA I1 6920 vehicle 

range fccm approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the singleengine 5 kW system. Maximum total impulse values and maximum delta-v’s range from 2.2~10~ to 3.6~10~ N-s, 

and approximately 662 to 1072 m/s, respectively. 

7. Key Words (Suggested by Author@)) 

Arcjet 

I Ion thruster 

Primary space propulsion 

18. Distribution Statement 

Unclassified - Unlimited 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No of pages 22. Price’ 

Unclassified Unclassified 

‘For sale by the National Technical Information Service, Springfield, Virginia 221 61 NASA FORM 1626 OCT 86