final technical report

Running Head: PROPULSION SYSTEMS FOR INTERPLANETARY TRAVEL

Comparing Propulsion Capabilities of the Variable Specific Impulse Magnetoplasma

Rocket and the NASA Evolutionary Xenon Thruster for Interplanetary Travel:

Christopher W. Hays

Embry-Riddle Aeronautical University

December 9, 2016

PROPULSION SYSTEMS FOR INTERPLANETARY TRAVEL 2

TABLE OF CONTENTS

Abstract……………………………………………………………………………………………4

Introduction

Overview…………………………………………………………………………………..5

Problem…………………...……………………………………………………………….5

Purpose…………………………………………………………………………………….6

Scope……………………………………………………………………………………....7

Methodology………………………………………………………………………………7

Data Analysis

VASIMR

Overview………………………………………………………………………….9

Thrust Capability………………………………………………………………...10

Thrust Efficiency……………………………………………….………………..13

Discussion of Power………………………………………………………….….15

Longevity and Degradation……………………………………………………...18

Evaluation……………………………….……………………………………….19

NEXT

Overview…………………………………………………………………………19

Thrust Capability…………………………………………………………...……20

Thrust Efficiency……………………………………………………..………….23

Discussion of Power……………………………………………………………..25

Longevity and Degradation………………………………………………………26

Evaluation………………………………………………………………………..28

Comparison……………………………………………..………………………………………..28

Conclusion…………………………….…………………………………………………………30

Recommendations…………………………………………………………………….……….…31

Appendices

Appendix A……………………………………………………….….……………....…..32

Appendix B……………………………………………………………………….......….33


LIST OF FIGURES AND TABLES

Figure 1: Visualization of VASIMR Operation Sequence………………………………………..9

Figure 2: Thrust Capability vs. RF Power Output Graph………………………………………..11

Figure 3: Chart of Thrust/Power Ratio vs. the Specific Impulse…………………………...……12

Figure 4: Graph of Thrust Efficiency vs. Exhaust Velocity……………………………………..14

Figure 5: Graph of Thrust Efficiency vs. Specific Impulse……………………………...………15

Figure 6: Helicon, ICH, Helicon Power Consumption Display……………………………….…16

Figure 7: Solar vs. Nuclear Power Comparison Chart…………………………………………...17

Figure 8: NEXT Discharge Chamber Operation Sequence Visual………………………………20

Figure 9: Graph of Total Impulse vs. Elapsed Time……………………………………………..21

Figure 10: Thrust to Power Input Relationship – BOL…………………………………………..22

Figure 11: Thrust to Power Input Relationship – EOL…………………………………………..22

Figure 12: Chart of the Required Power for NEXT Variations to Mars…………………………25

Figure 13: Grid Aperture vs. Hours of Operation………………………………………………..26

Figure 14: Graph of the Propellant Throughput vs Hours of Operation…………………………27

Figure 15: Thrust Comparison Graph………………………………………...………………….29

Figure 16: Thrust Efficiency Comparison Graph………………………………………………..29

Figure 17: Required Power Comparison Graph…………………………………………………30

Table 1: Characteristics of NEXT System Based on Mass Throughput Table………………….24


ABSTRACT

Electrical propulsion has become of extreme interest to scientist and engineers, and has shown

promise in becoming the primary propulsion type for interplanetary travel. Two electrical

propulsion systems currently in development that have shown promise are the Variable Specific

Impulse Magnetoplasma Rocket (VASIMR) and the NASA Evolutionary Xenon Thruster. This

report delves into the thrust qualities, power consumption and sources, and overall reliability to

determine the most effective thruster for interplanetary travel, and aims to initiate the discussion

into the most effective interplanetary electrical propulsion system. Researching current

developments for both systems and translating that information into side-by-side graphs, the

systems are compared based on their functionality of the qualities listed above. Through this

research, it is found that the VASIMR provides greater thrust (6 N) compared to the NEXT (.237

N), however the NEXT requires less power to operate (6.9kW to the VASIMR’s 200kW), and

the efficiencies of each engine were roughly equivalent (71% to 72%). Based on this criteria,

longevity, and proven record, the NEXT system was chosen to be the more effective thruster at

the present time. However, once each system proves itself in a space environment, they will need

to be reevaluated based on their performance in a space scenario for more accurate conclusions.


Introduction

Overview

According to NASA, Dr. Robert H. Goddard’s first liquid-fuel rocket engine launch on

March 16, 1926, “was as significant to history as that of the Wright brothers at Kitty Hawk”

(Garner, 2016). Since Goddard’s historic first flight, rocketry has become a primary focus of

many federal and private organizations for its uses in weaponry and scientific exploration.

Currently, liquid bi-propellant engines, similar to the one Goddard used, are the primary choice

for interplanetary space missions to other parts of the solar system. While these propulsion

systems are primarily chosen for interplanetary space missions, other propulsion systems have

been and/or are currently being researched for such purposes. Electrical and nuclear propulsion

systems are of special interest to replace the current chemical rockets for in-space transportation.

Although none have been used on a large, propulsive scale, Hall-Effect thrusters and resistojet

rockets have been used for station-keeping and have proved the effectiveness of electrical

propulsion in space. This raises the question: Can an electrical propulsion system ever be used as

a primary propulsion system for longer-range missions in space? Two projects are currently

underway to develop a long-range electrical propulsion system – NASA Glenn Research

Center’s NASA Evolutionary Xenon Thruster (NEXT) and Ad Astra’s Variable Specific Impulse

Magnetoplasma Rocket (VASIMR).

Problem

For years, chemical rockets have been used to propel man and machine into space. The

mammoth Saturn V rocket that put Apollo 11 on the moon was a multi-stage liquid fuel rocket

that stood 363 feet tall and used 5.6 million pounds of propellant (Boeing, 2016). The current

plans for multi-stage chemical rocket boosters to propel man to Mars dwarf the Saturn V.


Predicted to stand over 400 feet tall (Musk, 2016), the Interplanetary Transport System (ITS)

will be the largest rocket booster ever assembled. However, the larger the rocket, the greater the

cost. Elon Musk, CEO and Founder of Space Exploration Technologies, explains this increased

cost to the boosters will be offset by the reusability of the spacecraft. He claims that using the

craft for 12 round-trips to Mars will bring the cost down from $10 billion per person to

approximately $200,000 per person (2016). Although this plan will reduce the cost, fuel

efficiency and consumption problems remain. Electrical propulsion systems seem to be a

suitable replacement for modern chemical rocket engines, and may yet be able to outperform

them. Unfortunately, the majority of modern electrical propulsion systems do not have the thrust

capacity to propel larger spacecraft more than station-keeping. The NEXT and VASIMR are two

current attempts to develop a thruster capable of propelling larger spacecraft out into the solar

system. Although the NEXT and VASIMR have a higher thrust than other electric propulsion

systems, they do not have the capability to launch directly from Earth.

Purpose

The report dives into the discussion of the capabilities and detriments of developmental

stage electrical-propulsion systems to solve current feasibility issues for long-range space flight.

In this report, thrust capability and thrust efficiency are compared in how they propel spacecraft

through interplanetary missions. Although NEXT is further along in development and currently

being designed for commercial uses, both propulsion systems are far enough along in their

development for their thrust capabilities to be compared effectively. Key operational differences

of ion thrusters (NEXT) and electromagnetic propulsion (VASIMR) is discussed at a level for

those in the science community, specifically space flight researchers and developers, to

understand and further analyze the potential in different environments and scenarios.


Scope

The search for propulsion systems equipped for interplanetary travel may include a wide variety

of topics and criteria that differs with every mission. For these reasons, the main body of the

analytical report will be limited to the following topics:

1. Thrust Capabilities

2. Thrust Efficiencies

3. Operating Power

4. Life-Span

Each of these topics are consistent aspects when planning a mission and are key components to

any propulsion system’s design. Considering humanity’s push for interplanetary travel and

expansion into the solar system, these components will become of greater concern to scientific

researchers.

Due to the complexity and scale of the development of each of these propulsion systems,

no testing and research will be done in-house. All information needed will be gathered from

outside sources who had the funding and means to conduct testing on the propulsion systems.

However, all the information gathered will be presented side by side to visualize the differences

between the propulsion systems and why one may be better than the other, given certain

situations.

Methodology

A comparative analysis is conducted to determine the effectiveness of the NEXT and

VASIMR propulsion systems based on certain scenarios. Many of these scenarios will include

topics ranging from variable thrust capacity and thrust efficiency, power required to operate,

payload maximums and overall potential for utilization in interplanetary missions. Because little


to no physical research is conducted specifically for this topic, a heavy reliance will be on

outside sources. Some outside sources are used to develop a background for the coming research

and to provide context for any reader after the completion. The Ad Astra Rocket Company

provides an in-depth overview of the VASIMR propulsion system and thoroughly explains the

principles of operation and power sources available for the thruster. From this source, the basics

of operation are explained and further expanded upon by more scientific sources such as the

“Performance Studies of the VASIMR VX-200,” written by B.W. Longmier and company in the

49th AIAA Aerospace Sciences Meeting and Exhibit, Florida. The journal articles containing

pertinent scientific information, like the one above, are primarily used for the analytical portions

of the report and examining the aspects of each propulsion system. The American Institute of

Aeronautics and Astronautics puts on many conferences that directly influence the electrical

propulsion systems and some of the conferences directly influence this paper. Written by Pratik

Saripalli, a Graduate Research Assistant at the University of Maryland – College Park, the

“NEXT Performance Curve Analysis and Validation” was presented at the 52nd AIAA/SAE/ASEE

Joint Propulsion Conference in 2016. The article models the thrust capability through multiple

throttle curves that are dependent on the power Input and mass flow rate. The original research

for the report involves pulling data from numerous articles that were directly related to the

testing of the propulsion systems. All the data gathered from these articles are compiled and

graphed to easily visualize the key differences in the two propulsion systems.


Data Analysis

VASIMR Overview

The VASIMR thruster - currently in development by the Ad Astra Rocket company –

operates as an electromagnetic thruster that converts a cold gas – argon, hydrogen, or xenon -

into plasma through the use of two radio wave (RF) couplers. The first coupler (helicon) heats

and ionizes the gas into plasma that is then transferred to the second coupler (Ion Cyclotron

Heating), which superheats the plasma and expels it through a magnetic nozzle. The magnetic

nozzle serves to convert the ions’ orbital energy into linear energy capable of propulsion. Figure

1 outlines the operation sequence of the VASIMR thruster.

Figure 1 displays the operation sequence of the VASIMR VX-200 Thruster (Longmier, Ballenger, Olsen, Squire, Díaz, 2011, p. 2)

Ad Astra claims this new thruster has “many unique advantages” (n.d.) which include reliability,

longevity, and a variable specific impulse to accomplish a wide array of objectives. Specifically,

its variable specific impulse – the ability to transition between a high-efficiency and a high-thrust

system – has been widely publicized by the company (n.d.). This feature will allow the VASIMR

system to adapt to specific mission requirements based on necessity and demand.


A test to analyze the exhaust plume of the VX-200 engine conducted in May 2010 by

Benjamin Longmier of the Ad Astra Rocket Company served as a significant source for the

VASIMR rocket engine thrust capabilities and thrust efficiencies. Before these properties of the

test are discussed, important figures of the test must be noted. In this specific test, the magnetic

field in the engine was peaked at 2 tesla while the helicon coupler operated at 28kW of power

and the ICH coupler operated at 90kW of power. In addition, argon gas was used as the

propellant and injected through the engine at a mass flow rate of 107mg/s (2011).

Thrust Capability

A key factor guiding any design of a propulsion system is its capability to produce thrust.

Presently, chemical rockets are the only thrusters powerful enough to propel missions into orbit

and beyond. Modern thrusters, such as the RL-10A engine powering the Atlas V Centaur Upper

Stage, are operating on the order of 9.92x104 N (Atlas V, n.d.). The magnitude at which the

electrical propulsion systems operate are not near enough to serve as a full-scale launch system.

As a result, all electrical-propulsion systems are currently, and for the foreseeable future,

exclusively used in-space flight.

The VASIMR propulsion systems achieves thrust through the expulsion of superheated

plasma through a magnetic nozzle. When the particles are discharged from the nozzle, they form

a plume of superheated plasma that results in the opportunity to analyze the thrust, force density,

and thrust/power (T/P) ratios of the entire propulsion system.

According to Andrew Ilin and his team at the Ad Astra Rocket Company, the VASIMR

thruster has the potential to reduce the travel time from over 6 months to just 39 days (2011, p.

1). However, in order for this to be accomplished, substantial amounts of thrust must be

produced. Figure 2 demonstrates the VASIMR VX-200’s thrust capacity at varying power levels.


Figure 2 displays the relationship between the thrust and the RF power output of the engine. (Longmier et al., 2011, pg. 22).

The thrust is observed to increase with the increasing power, until the maximum amount of

power is reached (200kW), at this point a thrust of just under 6 N is seen. The 6 N of thrust

produced is significantly more than the comparable thruster (NEXT) which generates only

237mN of thrust at full power. This difference in thrust allows for greater acceleration over

shorter time periods. The 39 days are estimated upon these numbers.

The thrust to power ratio (T/P) is the measure of thrust produced per unit power. In the

case of the VASIMR it is in micronewtons per kilowatt. Optimizing the T/P ratio of a system

can aid in reducing orbit transfer times which supports the efforts to reduce the Mars travel

time. Any reduction to the travel time to Mars would benefit the mission through cost reduction

and diminish radiation exposure on the crew.


Figure 3 exhibits the optimum Thrust/Power ratio as a function of specific impulse of the VASIMR system. (Longmier, 2012, p. 7)

As seen in Figure 3, the greatest T/P ratio is greatest (51mN/kW) at a specific impulse of 1699

s. This data was collected during a campaign to optimize the T/P ratio performed by the Ad

Astra Rocket Company. However, the maximum T/P ratio does not occur at the maximum

efficiency of the system which occurs at 5000s of specific impulse. This contrast will not cause

in-flight problems, however, it may cause some logistical issues concerning mission priorities

and fuel weight and consumption. While the increased thrust to power ratio will decrease orbit

transfer times, the efficiency of the thruster will decrease. This is a direct result of the increase

in thrust necessary to achieve a greater change in velocity (ΔV) to accomplish the orbit transfer.

Observed in Figure 3, the 20mN/kW jump from 30mN/kW to 50mN/kW costs 3000s of specific

impulse. This significant gap causes a discrepancy in the functionality of the VASIMR as a high

efficiency, high thrust electrical propulsion system. It immediately becomes a high thrust or

high efficiency and may vary between the two dependent on the mission phase.


Thrust Efficiency

The efficiency of an engine aids in determining the overall effectiveness and quality of

the system based upon how adequately it consumes the fuel. The mass flow rate of the fuel

through engine (mflow) and the specific impulse (Isp) are key indicators of the efficiency of the

engine. The total thrust efficiency of the VX-200 can be calculated using the following equation:

𝜂𝑡 =𝑃𝑗𝑒𝑡

𝑃1𝑅𝐹+𝑃2𝑅𝐹 (1)

Where 𝑃1𝑅𝐹 and 𝑃2𝑅𝐹 are the helicon power and ICH power respectively and 𝐹𝑗𝑒𝑡 is described

by the equation:

𝑃𝑗𝑒𝑡 =𝐹2

2𝑚𝑓𝑙𝑜𝑤 (2)

F is the total force of the engine and 𝑚𝑓𝑙𝑜𝑤 is the mass flow rate of the gas through the pipe.

Using the equations outlined above, Longmier and his team measured a maximum thrust

efficiency of 72% with an 𝐼𝑠𝑝 of 5000s at a total RF coupled power of ~200kW (approximately

186kW of coupled helicon and RF power). This amount is significantly more than the predicted

operating efficiency of 60% at the same 𝐼𝑠𝑝 and power, as displayed in Figure 4. This means the

VASIMR is operating a much higher efficiency than expected and will incorporate less fuel mass

which will allow for a greater payload mass.


Figure 4 displays the efficiency relationship with exhaust velocity and the difference among 3 separate trials and theoretical values. (Longmier et al., 2011, pg. 23)

Figure 4 shows the relationship the thrust efficiency has with the exhaust velocity of the charged

argon ions exiting the engine. Noticeably and expectedly it increases with the increasing exhaust

velocity, which is also representative of the increasing coupled RF power. As the exhaust

velocity approaches 50 km/s, the efficiency begins to level off at 72%, however the graph still

shows the potential for an increase in efficiency as higher exhaust velocities are achieved. It has

yet to be seen whether the thruster can accomplish and maintain these high thrust efficiencies for

extended periods of time, like the NEXT.


Figure 5 displays the thrust efficiency with respect to the specific impulse (Longmier et al., 2012, p. 8).

Figure 5 shows the increase in thruster efficiency with the increasing specific impulse. The

figure shows that as the specific impulse of the thruster increases so does the efficiency. This is

expected based on the nature of specific impulse. This means while operating at higher

efficiencies the system could provide constant thrust over longer periods of time. For

interplanetary travel, this would allow for the spacecraft to reach very high speeds with lower

power outputs than comparable thrusters.

Discussion of Power

The typical operating power for the VASIMR system hovers at about 200kW. The

majority of this power is being consumed by the two couplers stimulating the ionization and

heating process for the gas particles passing through the VASIMR. Figure 6 details the power

consumption of the helicon and ICH couplers used in the VX-200 model.


Figure 6 displays the coupled RF power of the VASIMR system (Longmier et al., 2011, p. 13).

As seen in Figure 6, this power consumed by the helicon and ICH RF couplers is just shy

of 190kW. Of which, the helicon coupler employs approximately 30kW, leaving the remaining

160kW to the ICH coupler. This gap in power usage stems from the tasks each coupler

completes. Where the helicon ionizes the gas and heats it to a cold plasma state (5800K), the

ICH superheats this plasma to temperatures similar to that of the sun’s core (10 million K) which

requires a significant increase in power, as seen in Figure 6 (Ad Astra, n.d.). However, the

200kW of power required for the system is not the total amount of power needed by the thruster

to propel the craft through interplanetary space. Estimates made by Andrew Ilin et al. of the Ad

Astra company place the required power of the VASIMR thruster to be 12MW (megawatts) of

power for a four-month trip, and 200MW for a trip less than two months (2011). Assuming the

claims made for the system are true, the entire system is projected to consume over 200MW of

power along the expedition.

A sufficient power source is required to output this substantial amount of power to the

system. Currently, the two power sources being discuss are solar and nuclear. Both are viable


options and have been proven to work on long-range space missions before. However, they each

have different performance capabilities and can be used based on mission specifications. Figure

7 below displays the trip time in days versus the power consumed by the system.

Figure 7 exhibits the decrease in travel time based on the power consumption of the system. An initial mass of 100mT and 60% power efficiency was assumed. Since the creation of this graph, the efficiency of the VASIMR system has been proven to operate at 72% lowering all travel times. The graph depicted only gas to 50MW because of the initial mass used (Ilin et al., 2011, p. 4).

Figure 7 displays the trip time to Mars as a function of the required power (MW) for both

solar and nuclear power. In general, the nuclear powered thruster seems to allow for quicker

travel times to and from Mars. Observed from the graph, any power change from 2.5MW-10MW

will draw a significance decrease in the travel, however, after 20MW of power the decrease in

time begins to level off. At approximately 50MW of power the travel time is predicted to be just

under 50 days for both power sources. The nuclear power source trends less than solar due to the

lack of power production loss. As a solar powered craft ventures further from the sun, it

experiences a significant loss in power. As a result, nuclear power has been the power source of

choice for many outer solar system missions.


Longevity and Degradation

As of today, no extensive longevity and degradation testing has been conducted on the

VASIMR engine. However, in July, at the Propulsion and Energy 2016 Forum and Exposition

Jared P. Squire et al. presented a paper detailing how they plan to test how effective the thruster

is over longer periods of time. According to Squire, initial testing plans to run the VASIMR for

at least 100 continuous hours at power levels greater than 100kW (2016). These experiments

come after a series of pulsing tests for power levels up to 200kW that were conducted to

determing the operational characteristics of the system. During this new series of tests, the

operation plans to find the thermal structure and determine a lifetime estimate for the entire

system. From this data, the longevity qualities of the thruster can be evaluated and will provide a

better estimate for how the thruster will perform in a space environment.

Ad Astra has made arrangements to conduct a series of iterative tests using VX-200SSa,

VX-200SSb, and VX-200SS models to continually expand the capability limits of the system.

Testing of the VX-200SSa is set to begin late 2016 into early 2017 to characterize thermal and

power performance (Squire et al., 2016). This initial phase of testing will allow for the

determination of thermal characterization and power performance and help set up the structure

for later testing iterations. The VX-200SSb will be tested subsequently in an effort to ensure the

vacuum and testing chamber is fully operational and can handle the extreme elements before

proceeding into full testing of the system. Using the information gathered from the first two tests,

the company will proceed with testing of the VX-200SS to determine its longevity in a

comprehensive atmosphere – full temperature, +100kW, +100 hours of operation. From here, the

quality of the thruster for lengthier time periods can be extrapolated. Once this series of testing is


complete, the longevity and degradation of the VASIMR thruster can be more thoroughly

estimated and evaluated.

Evaluation

To date, the VASIMR is the most powerful electrical propulsion system. Expelling 6 N of

thrust from its magnetic nozzle, it out performs other electrical propulsion systems by a full order

of magnitude. This increase in thrust means shorter travel times and allows for more frequent and

higher quality missions on the interplanetary scale. Along with the increased thrust levels, the

VASIMR has shown an increased thrust efficiency and has outperformed predictions. Topping

off at 72% efficient, the VASIMR topped the predicted values by almost 12%. However, the

power levels to sustain the 6 N thrust and 72% efficiency is approximately 200kW. Maintaining

this level of power could pose problems over long-term interplanetary missions, considering

those power levels have not been sustained for more than a few seconds. Because the VASIMR

has not conducted an in-space flight test, it is difficult to effectively compare it to any similar

thruster that is currently in development or on the market. What is known results from extensive

testing by the Ad Astra Rocket Company and NASA, however no definitive conclusions may be

made until it is known how the thruster system operates in a space environment.

NEXT Overview

According to NASA, they have “identified the need for a higher-power, higher-specific-

impulse, higher-thrust, and higher-throughput capable ion propulsion system” (Shastry, Herman,

Soulas, Patterson, 2015, p. 2) for which they are developing the NASA Evolutionary Xenon

thruster. The NEXT propulsion system – currently in development by the NASA Glenn Research

Center – is a form of ion thruster and a reiteration of the NASA Solar Technology Application

Readiness (NSTAR) which proved the effectiveness of ion thrusters on Deep Space 1 and


Voyager. Operating as an ion thruster means the NEXT accelerates Xenon ions through

processes of ionization and passes the ions through a magnetically charged nozzle, similar to the

VASIMR. The NEXT operates as a gridded ion thruster which services two grids – the screen

electrode (positive grid) and the accelerator electrode (negative grid). It begins by injecting

neutral particles and electrons into a discharge chamber filled with magnetic rings bordering the

perimeter. These magnetic rings generate a magnetic field in which the atom and electron

collide, ionizing the atom. From there, the positive ions are driven through the discharge

chamber by the magnetic field towards the electrode grids which accelerate the ions out of the

nozzle providing thrust (Mindek, 2015). Figure 8 describes the operating system of the discharge

chamber within the NEXT thruster.

Figure 8 outlines the operating elements within the NEXT discharge chamber (Anthony, 2012)

Thrust Capability

The NEXT system has been shown to consistently output a thrust of 237 mN

(microNewtons). When compared to other booster systems, this level of thrust seems miniscule.


However, when paired with the run times of the system, velocities on the order of 321,000 kmh

can be achieved (Dunbar, 2008).

The total impulse of the NEXT greatly exceeds the total impulse of its predecessor

NSTAR as shown in Figure 9. This increase in the total impulse is the result of the necessity for

higher thrust boosters. The NSTAR only attained a total impulse of 7 MN-s, whereas the NEXT

achieves close to 17 MN-s. This difference in total impulse allows for the NEXT to propel larger

payloads faster and for longer periods of time. With this greater propulsion capability, the

payloads carried by electrical propulsion systems will become larger and eventually support

manned mission to extraterrestrial planets. Additionally, as seen in Figure 9 the NEXT thrusters

last over 20kH longer than the NSTAR. This supplementary operation time solves many issues

regarding space flight. It supplants the need for contingency plans by having extended range and

operational lifetime while allowing for extended missions into interplanetary space.

Figure 9 shows the relationship of the Total Impulse (MN-s) vs the Elapsed Time (Kh) of the NSTAR and the NEXT at various throttle levels to aid in visualization of the performance capabilities of the NEXT over the NSTAR (Shastry et al., 2015, p. 5).

A study conducted by Pratik Saripalli, a graduate research assistant at the University of

Maryland, alongside Eric Cardiff and Jacob Englander of the NASA Goddard Flight Center,

demonstrated the thrust capabilities of the system (mN) as a function of the input power from the


Power Processing Unit (PPU) in watts. Figures 10 and 11 are used for a comparison of the thrust

capabilities at the beginning of life (BOL) and end of life (EOL) of the system.

Figure 10 Displays the thrust capabilities of the NEXT as a function of power input for the beginning of life (Saripalli, Cardiff, Englander, 2016, p. 7)

Figure 11 Displays the thrust capabilities of the NEXT as a function of power input for the end of life (Saripalli et al., 2016, p. 8)

The thrust curves in Figures 10 and 11 show little change over the approximately

30,000hr flight time this attributes to the longevity and quality of the system. Figures 10 and 11

are an attribute to the longevity and effectiveness of the NEXT thruster as a system. It has proven

to be as equally effective after 30,000 flight hours as 0 flight hours, which shows greater

reliability for long-duration missions through interplanetary space. Maintaining the constant


thrust of 237mN allows for the accumulation of force to propel the spacecraft through

interplanetary space. This constant thrust allows for reliable acceleration for greater periods of

time which allows for greater velocities over long-range space missions.

Thrust Efficiency

According to NASA, the NEXT has become one of the most fuel-efficient thrusters in

space flight with a total of 51,200 hours of operation, only using 918.2 kg of Xenon as propellant

(Patterson & Pencil, 2014, p. 6). That is equivalent to .018kg of Xenon used per flight hour. It

also makes it the longest flight test in space propulsion at five and a half years. When this is

compared to conventional chemical rockets which use approximately 10,000 kg of propellant to

provide the same momentum over less time (Heidman, 2015).

Tests ran by Rohit Shastry show the maximum efficiency achieved by the NEXT is 71%

and decreases with thrust and power input, as shown in Table 1. It is common to see the thrust

efficiency rise with the decrease in thrust. However, that is not the case, according to Table 1. It

is observed that the thrust efficiency is more closely tied to the Input Power from the PPU then it

is to the thrust produced. The power input increase causes an increase in both the thrust and

thrust efficiency of the system.


Table 1 displays significant characterisitics for the NEXT thruster where the different shades of gray are representative of the mass Xenon throughput. (Shastry et al., 2015, p. 7).

The tests ran in Table 1 were conducted at different throttle levels and Xenon throughput

masses to determine their effect on the thrust capability and efficiency as well as many attributes

of the system. From this, the relationship between these attributes can be visualized and better

understood. According to this visual, the most efficient throttle levels are in the TL40 range

coupled with the 300kg and 450kg throughputs which operate consistently within the 71% range.

Analysis of the propellant used conducted by Saripalli shows the amount of fuel needed

for a trip to Mars is between 140-150kg dependent on high impulse or high thrust stipulations.

When paired with the trip times of 350-450 days for high thrust or high impulse scenarios

(2016), the amount of fuel used per hour of flight is .013kg/hr for high impulse flights and

.018kg/hr for high thrust flights. Both of these are extremely efficient when compared to modern

propulsion systems traveling to Mars, so both types of flights would allow for lower amounts of

fuel and greater payload weights. It is almost irrelevant if a high thrust or high efficiency

configuration is used because the high thrust model will reach shorter destinations quicker,


whereas high efficiency models are better equipped for longer voyages due to their thrust

duration.

Discussion of Power

The NEXT thruster typically operates nominally between .5 and 6.9kW (Shastry et al.,

2015, p. 3). However, according to Saripalli, the peak power generated needs to be 10kW for a

trip to Mars for both high thrust and high impulse scenarios. This is confirmed by Figure 12

provided by Michael Patterson and Eric Pencil of the NASA Glenn Research Center, where the

minimum amount of power for a trip to Mars is 10kW. However, this problem should be solved

by using multiple NEXT thrusters, or further development of the PPU.

Figure 12 lays out the operational power range of the NEXT and provides sample missions for given operational powers (Patterson & Pencil, 2014, p. 12).

The PPU is the primary power source for many of the major components of the NEXT

and is currently designed to operate between .5 and 6.9kW. The PPU Power output (seen as

power input from other systems) has been used as the independent variable for many evaluations

and controls many of the aspects of the propulsion system.


Similar to the VASIMR, the NEXT can be powered by both solar and nuclear energy.

Each provides the system with sufficient amounts of power further away from the sun due to the

NEXT system’s lesser power requirements. However, nuclear is still the better option for

interplanetary travel based upon its continuous steady power output.

Longevity and Degradation

The longevity of any interplanetary propulsion vehicle is critical to the success of longer-

range space flights. According to Rohit Shastry of the NASA Glenn Research Center, the NEXT

thruster has logged over 50,170 hours of operation with minimal wear (2015, p. 1). This level of

longevity allows the NEXT thruster to outperform its competitors barring its limited thrust

capacity. These results were determined by the NEXT Long-Duration Test (LDT) conducted by

the NASA Glenn Research Center beginning in May 2005. The test was set to run until failure

but was ceased due to budget restrictions in April 2013.

Figure 13 shows the relationship between grid aperture diameter and hours of operation (Shastry et al., 2015, p. 22).

Figure 13 shows the relationship between the grid diameter as a function of the hours of

operation for the NEXT system. The grids of the system serve to propel ions through the nozzle

and increase the exhaust velocity to accumulate greater momentum. Each grid is specifically


designed to accelerate ions effectively and efficiently to nominal performance speeds, and any

deviation in the aperture diameter decreases the quality of the grids. As seen, the diameter of the

grid aperture of NEXT does not show significant increase over a prolonged time period, like it

does in the NSTAR. This improvement allows for greater enduring quality of operation in the

NEXT system and grants greater applications. This aids in the longevity and efficiency of the

NEXT system as a whole. The tighter the grid aperture is the more effective the thruster will be,

which means if the aperture radius can stay at specified sizes for extended periods of time allows

for the thruster to perform at optimal levels for longer.

With these applications, the NEXT must be capable of handling greater amounts of

propellant throughput. A goal of 450 kg of xenon propellant was set for the LDT, but it was

found that the thruster was capable of handling 902+kg of xenon propellant. This substantial

amount of propellant throughput increase raises the potential for long-duration space missions.

As shown in Figure 14, the propellant throughput of the NEXT system exceeds the throughput of

comparable spacecraft by a substantial margin when the systems operate for greater than 30

kilohours (kh). It outpaces the NSTAR by a margin of approximately 650 kg of throughput.

Figure 14 compares the propellant throughput of the NEXT system to other Spacecraft (Shastry et al., 2015, p. 5).


Evaluation

The NASA Evolutionary Xenon Thruster, a successor to the NSTAR, outperforms its

predecessor in almost all forms of functionality. The 237mN thrust is higher than most other

comparable electrical-propulsion systems except the VASIMR, and it consistently puts out an

efficiency of 71%. The uniqueness of the NEXT, however, stems from its ability to maintain

those levels of functionality for over 50,000 flight hours. This allows for the NEXT to operate

over extended periods of time and achieve exceedingly high velocities. Its longevity is currently

unmatched and top of its class, further allowing it to operate over extended periods of time with

little concern for wear and degradation.

Comparison

The NEXT and VASIMR thrusters are both in-development, electric-propulsion systems

with aims to increase the amount of thrust produced by non-chemical thrusters. Each system has

amassed its own qualities and stands to support itself in the realm of interplanetary travel.

Analyzing the thrust capabilities of each system in Figure 15, it is observed that the VASIMR

outclasses the NEXT by roughly an order of magnitude (6 N to .237 N). This difference in thrust

allows for the VASIMR to propel larger payloads with greater acceleration, allowing for earlier

arrival times at destinations. However, according to Figure 17, the power required to maintain

the VASIMR is significantly more than the NEXT, which may cause problems as the system

travels further away from Earth. For example, the VASIMR has only been seen to operate at

maximum power for a few seconds where the NEXT has been proven to operate at maximum

power for over 50,000 hours of operation. The thrust efficiencies of both thrusters, viewed in

Figure 16, are essentially equivalent - 71% for the NEXT and 72% for the VASIMR - for shorter

missions, however, the efficiency levels for the VASIMR may come in to play for longer

extended interplanetary travel missions based on fuel consumption.


Figure 15 graphs the thrust capabilities in Newtons of both the NEXT and VASIMR propulsion systems so they may be compared is a side-by-side way (Original Research, Longmier et al., 2011, p. 22; Saripalli et al., 2016, p. 7).

Figure 16 graphs the thrust efficiencies in a percentage for both the NEXT and VASIMR propulsion systems so they may be compared is a side-by-side way (Original Research, Longmier et al., 2011, p. 23; Shastry et al., 2015, p. 7)


Figure 17 graphs the required power in kilowatts for both the NEXT and VASIMR propulsion systems so they may be compared is a side-by-side way (Original Research, Longmier et al., 2011, p. 13; Shastry et al., 2015, p. 3).

Based on these findings, it is difficult to determine the more effective thruster. Each have

their own advantages over the other. However, once advancements in VASIMR technology

becomes more readily available, and the system is proven in a space environment, it will become

a very effective system for space flight.

Conclusion

In conclusion, NEXT and VASIMR have each made significant strides in the appropriate

directions to ensure the further development and integration of electrical propulsion systems into

the realm of interplanetary travel. Each system has been developed to ensure all aspects of an

effective in-space propulsion are met and exceeded. The better of the two systems is going to be

highly dependent on the specifications of the mission. However, as the VASIMR is still in

development and has not undergone any in-space test flights, it still needs a few design

adjustments and developments until it can be made ready for space flight. Until then, the NEXT

is the better system for in-space flight based upon its longevity and proven record of reliability.


Recommendations

When a decision must be made between the VASIMR and NEXT propulsion systems, the

three following criteria must be considered:

1. High Thrust or High Impulse

2. Power Requirements

3. Time

Analyzing these criteria will aid in the decision by narrowing down the mission specifications to

determine the appropriate thruster. Missions with lower power requirements would involve the

NEXT system, whereas a mission with high power requirements would involve the VASIMR

system. Along the same lines, time sensitive missions would depend on the VASIMR for its

higher thrust capabilities. However, to fully compare the propulsion systems, a cost analysis of

each would be conducted to determine the cost effectiveness, price per launch, and overall

operations cost of the systems. These topics will be further analyzed in another report centered

solely on the expenditures of each system.


APPENDIX A

Abbreviations

BOL – Beginning of Life

EOL – End of Life

ICH – Ion Cyclotron Heating

ITS – Interplanetary Transportation System

LDT – Long Duration Test

NASA – National Aeronautics and Space Administration

NEXT – NASA Evolutionary Xenon Thruster

NSTAR – NASA Solar Technology Application Readiness

PPU – Power Processing Unit

RF – Radio Wave

TL – Throttle Level

VASIMR – Variable Specific Impulse Magnetoplasma Rocket

Nomenclature

𝐹 - total force

Isp – specific impulse

mflow – mass flow rate

𝑃2𝑅𝐹- ICH power output

𝑃1𝑅𝐹- Helicon power poutput

𝑃𝑗𝑒𝑡- thruster jet power

T/P – Thrust to Power Ratio

𝜂𝑡- thruster efficiency


APPENDIX B

Creation of Tables

The original research tables were assembled from in-depth testing reports to determine

the thrust and power characteristics of both the NEXT and VASIMR systems. Benjamin

Longmier, of the Ad Astra rocket company, conducted tests to determine the thrust capability,

and power requirements for the VASIMR rocket and the values found in his study were used in

the table. The thrust capability values for the NEXT system came from a study conducted by

Pratik Saripalli of the University of Maryland – College Park, while the other values came from

a study conducted by Rohit Shastry of the NASA Glenn Research Center.


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final technical report

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