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Performance Characterisation of an Air-breathing
thruster neutralised by an Air-breathing Cathode at
NewOrbit Space
IEPC-2024-385
Presented at the 38th International Electric Propulsion Conference, Toulouse, France
June 23-28, 2024
Alexander Schwertheim∗
, Ruslan Rakhimov†
, Juan M. Arteaga‡
,
Chengyu Ma§
, and Anatolii Papulov¶
NewOrbit Space Ltd, Reading, Berkshire, RG45AF, UK
We present an Air-Breathing Electric Propulsion (ABEP) System designed to provide
full drag compensation for Ultra-low Earth Orbit (ULEO) applications. The propulsive
section of the ABEP tested here includes an air-breathing Radio Frequency (RF) gridded
ion engine, an air-breathing RF cathode, and two dedicated Radio Frequency Generators
(RFGs). The thruster performance is measured by operating the thruster and the accom-
panying cathode on mixtures of oxygen and nitrogen. The thruster was operated on a 1:1
ratio of O2:N2by mass over combined mass flow rates of 9-14sccm at RF input powers
ranging from 150-324 W. For all measurements, the cathode was operated on 0.35 sccm
of oxygen and 0.35 sccm of nitrogen at 120 W of input RF power. Using indirect thrust
measurements, we observe thrust up to 15.1 mN and thruster specific impulses of 6380 s,
resulting in a maximum thrust-to-power ratio of 13.4mN/kW. These data suggest that
with the addition of our air-intake, this system is capable of fully compensating for drag
under certain ULEO conditions.
∗Propulsion Test Engineer, Alexander@NewOrbit.space
†Chief Technical Officer
‡Electrical & Electronics Engineer
§Mechanical Engineer
¶Chief Excecutive Officer
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The 38th International Electric Propulsion Conference, P. Baudis Convention Center, Toulouse, France, June 23-28, 2024
Copyright 2024 by the Electric Rocket Propulsion Society. Al l rights reserved.
I. Background
Many are eyeing Very-low Earth Orbits (VLEO) as the next step in the commercialisation of space. These
are orbits below 400 km which in comparison to traditional LEO altitudes greatly increase the achievable
resolution for Earth observation and reduce the minimum latency and transmission power requirements for
telecommunication. Typical LEO spacecraft cannot operate at VLEO altitudes for extended periods due
to the effect of aerodynamic drag. The atmospheric density in these orbits is several orders of magnitude
greater than at LEO and can fluctuate greatly with location, solar activity, and geomagnetic activity.1This
increased density dictates that a typical spacecraft at VLEO altitudes experiences a retarding drag force in
the order of millinewtons, which, if left unmitigated, would lead to orbital decay and re-entry within several
days or weeks of operation. This decay can be postponed by compensating atmospheric drag using a suitable
propulsion system. This principle was successfully demonstrated by the European Space Agency’s Gravity
Field and Steady-State Ocean Circulation Explorer (GOCE)2and the Japanese Super Low Altitude Test
Satellite (SLATS).3These satellites demonstrated that stable VLEO orbits as low as 250 km can be achieved
by using onboard electric propulsion systems to offset the atmospheric drag.
For VLEO missions such as GOCE and SLATS, where drag is compensated with thrust, the mission lifetime
and hence the feasibility of such missions are strictly limited by the amount of propellant (here xenon) that
can be stored onboard. Once all propellant has been consumed, drag can no longer be compensated, and
the mission ends in re-entry. This demands that a considerable fraction of the mass and volume budget of
VLEO spacecraft must be allocated to propellant storage.
All of the benefits of VLEO operation are amplified when flying lower still. We refer to orbits below
250 km as Ultra-Low Earth Orbits (ULEO). These altitudes are half that of traditional LEO, which dictates
Earth observation resolutions are doubled, and latency of telecommunications is halved. At ULEO altitudes
less radiation reduces the need for radiation hardening, thus reducing the time and resources required for
hardware development.4However attractive these orbits may be, the drag caused by the atmosphere at these
altitudes is an order of magnitude above that which was experienced at VLEO by GOCE and SLATS. The
drag on any object orbiting in ULEO determines that any space debris created here will rapidly re-enter,
making these orbits highly sustainable and self-cleaning.5Such high drag forces demand propellant storage
requirements which make ULEO altitudes unattainable for stable operation using traditional electric propul-
sion technologies.
An alternative to using stored propellant for electric propulsion has been proposed to enable extended
ULEO operation: Air-Breathing Electric Propulsion (ABEP). The core components and configuration of
such a system are shown in Fig.1. An atmospheric intake system is positioned on the ram-facing body of
Thruster
Direction of Flight
Intake
Atmospheric
Particles Accelerated
Ions
Spacecraft
Bus
Energy via Solar Arrays Cathode
Electric propulsion system
Figure 1. The core components of a typical air breathing electric propulsion system.
2
The 38th International Electric Propulsion Conference, P. Baudis Convention Center, Toulouse, France, June 23-28, 2024
Copyright 2024 by the Electric Rocket Propulsion Society. Al l rights reserved.
the spacecraft and acts to collect the rarefied atmosphere which is incident on the spacecraft. The collected
atmospheric particles are then supplied to the electric propulsion system, which comprises a thruster and (in
some cases) a cathode. The electric propulsion system acts to ionise the collected particles before accelerat-
ing and ejecting them using the power supplied by the spacecraft bus. As with a stored propellant electric
propulsion system, the thrust produced by the ABEP system must be greater than the drag on the entire
spacecraft. By continually harvesting propellant in situ, the lifetime of an ABEP system is not limited by
propellant storage. This greatly increases mission lifetime, with the only limit becoming the degradation
of the spacecraft components. Since its conception, there have been many attempts to demonstrate ABEP
systems in the laboratory.6–13 These attempts have achieved varying degrees of success, yet none have
demonstrated full drag compensation. In this context, full drag compensation refers to the case in which the
thrust generated exceeds the drag of the spacecraft, which is a necessity for steady-state ULEO operation.
II. Context
NewOrbit Space Ltd is a spacecraft manufacturing startup founded in the UK in 2021, enabling our cus-
tomers true ULEO access below 200 km, through the development of ABEP-capable platforms. Our ABEP
system is developed in-house specifically for this purpose and comprises an atmospheric intake system, a
Radio Frequency (RF) Gridded Ion Engine (GIE) as the thruster, an RF cathode as the neutraliser, and
the associated electrical and fluidic support subsystems. Each of these components is the result of several
stages of testing and iterative design based on performance measured in our space laboratory in Reading,
UK. Internal experiments have shown that when combining the intake system, thruster and cathode into
our complete ABEP system and operating under certain ULEO-like conditions, the system is now able to
generate more thrust than the simulated drag for such a spacecraft.
The NewOrbit intake system is able to supply mass flow rates and pressures high enough to allow the
ignition and operation of an air-breathing GIE. Where some published ABEP experiments report supply-
ing the intake with mass flow rates in excess of 4.5 mgs−1(213 sccm) in order to operate the thruster, the
NewOrbit propulsion system is able to operate with mass flow rates below 0.22 mgs−1(10 sccm) injected to
the intake.
In this study, we focus on the development of the electric propulsion system, which includes thruster, cathode
and radio frequency generators (RFGs), each of which are developed in-house and detailed in the following
section. In section IV we describe a test campaign of our electric propulsion system conducted within our
space-system testing laboratory’s vacuum facility. These tests include indirect thrust measurements of the
GIE when neutralised by our cathode with both components operating on oxygen and nitrogen only. For
these experiments, both thruster RFG and cathode RFG were mechanically coupled to their respective de-
vices and operated under vacuum for the duration of the experiments. Results are presented in section V
and discussed in section VI.
III. Test Articles
The core components of the NewOrbit propulsion system under test for the present campaign are the air-
breathing RF GIE as the thruster, the air-breathing RF cathode as the neutraliser, and two radio frequency
generators, with one being optimised for the thruster and the second for the cathode. Note that the nature
of ABEP determines that thrust requirements and available mass flow rates are continuously changing, often
over a large range. This requires all components have the ability to be deeply throttleable during operation.
A. Thruster
The electric thruster in an ABEP system must ionise the particles captured by the intake and accelerate
them sufficiently to compensate for the drag of the entire spacecraft. Traditional RF GIEs designed for xenon
operation are used commercially for North-South Station Keeping (NSSK) due to their very high specific
impulse (Isp) and long lifetime.14 Two of the most successful VLEO missions were GOCE and SLATS, both
of which utilised Kaufman type GIEs operating on xenon for drag compensation.2, 3
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The 38th International Electric Propulsion Conference, P. Baudis Convention Center, Toulouse, France, June 23-28, 2024
Copyright 2024 by the Electric Rocket Propulsion Society. Al l rights reserved.
Radio frequency RF GIEs like the one developed by NewOrbit utilise a different ionisation mechanism
to Kaufman type GIEs, but the same ion acceleration mechanism. An RF GIE was selected over alternative
thruster technologies due to the absence of a cathode inside the discharge chamber, the high specific impulse
and the relatively high thrust level. The NewOrbit GIE differs significantly from a traditional xenon GIE
due to modifications required to operate on the oxygen-nitrogen mixture collected by our intake. Such mod-
ifications required a range of optimisations, including modification of the RF coil, the ion optic assembly,
and careful material selection and testing to withstand oxidation from this highly reactive propellant. The
GIE is mounted directly to the thruster RFG to reduce EMI and power losses. The thruster is shown inte-
grated with the RFG in Fig.2. When operating on oxygen-nitrogen mixtures, in recent exploratory testing
Figure 2. The NewOrbit RF GIE mounted on the radio frequency generator.
this generation of thruster has demonstrated thrust in the range of 1-15 mN, with a specific impulse above
8500 s, a mass flow range of 0.1-0.6 mgs−1and a total power consumption near 1500 W. In this article we
only included a subset of initial test points of the thruster, which shows a limited section of the achievable
operational space. Simulations suggest that this thruster should be able to achieve a specific impulse of up
to 10,000 s, and up to 22 mN of thrust.
B. Cathode
Our RF cathode is used for the ignition of the thruster and the subsequent neutralisation of the ion beam.
Traditional emitter-type neutralisers employed in many xenon electric propulsion applications rely on low
work-function materials such as LaB6, which are extremely sensitive to oxygen.15 An RF cathode was
selected as a much less sensitive neutraliser technology for our ABEP application. This device has undergone
several design iterations to reach the state shown in Fig.3. It has the ability to ignite and operate on oxygen,
nitrogen, and mixtures of the two with no need for the addition of any noble gases. Similar to the thruster, it
is shown here integrated with the dedicated radio frequency generator. This generation of cathode consumes
up to 120 W of power, with a mass flow rate range of 7x10−3mgs−1up to 3.5x10−2mgs−1and can extract
a current of up to 700 mA.
C. Radio Frequency Generators
To test the thruster and the cathode at various operating points, we have developed two RFGs that allow for
the coil current, frequency, and power to be altered in real time. The design of the RFG circuits was done
in line with the requirements from the cathode and the thruster, respectively. Firstly, the electromagnetic
environments of both the thruster and the cathode were characterised as two-port networks using a Bode
100 vector network analyser by Omicron Lab. Resonant power converters were then designed and built to
4
The 38th International Electric Propulsion Conference, P. Baudis Convention Center, Toulouse, France, June 23-28, 2024
Copyright 2024 by the Electric Rocket Propulsion Society. Al l rights reserved.
Figure 3. The Cathode with attached radio frequency generator.
drive the respective coils considering their expected loading conditions.
Both RFGs employ Gallium Nitride transistors and achieve efficiencies from the DC source to the AC
output well over 90%, according to SPICE simulations. The expected temperatures from simulations are
congruent with those measured onboard during the experiments. The end-to-end efficiency from DC source
to plasma was estimated between 59.4% and 70.0% for the thruster operating on the entire designed range
of mass flow rate and power. According to simulations and characterisation with the vector network anal-
yser, most of the losses are due to inductive coupling and therefore eddy currents in the structures and shields.
The thruster RFG, which is seen integrated with the thruster in Fig.2, was designed and tested at an
input power of 150-400 W and driven at frequencies between 1-5 MHz. The cathode RFG, which is seen
integrated with the cathode in Fig.3, was tested at power levels of 80-240 W and operated at frequencies
between 3-11 MHz.
The RFGs were built with commercial off-the-shelf components and custom magnetics and were tested
in vacuum. The design features a custom casing that sinks the heat from the critical electronic components.
During testing, the back of the casing was fit to a water-cooled plate, which emulates heat transfer to the
spacecraft radiators. A thermal break was has been implemented between the structure of both the cathode
and the thruster and the respective RFG casings.
IV. Experimental Design
This test campaign surveyed the thrust and specific impulse generated by the propulsive section of
NewOrbit’s complete ABEP: the thruster and cathode both operating on a mixture of oxygen and nitrogen.
All experiments were conducted during April of 2024 within a vacuum facility at NewOrbit headquarters in
Reading, UK.
The facility comprises a chamber of 80 cm diameter and 1 m length, which is pumped by a 16 L/s dry
scroll primary/backing pump and a 2000 L/s turbo molecular pump. This facility is able to produce ul-
timate pressures lower than 1x10−5mbar. A dynamic pressure below 8x10−5mbar was maintained for all
experiments described here.
The complete electrical and fluidic setup is shown in Fig.4. For these experiments, propellants were supplied
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The 38th International Electric Propulsion Conference, P. Baudis Convention Center, Toulouse, France, June 23-28, 2024
Copyright 2024 by the Electric Rocket Propulsion Society. Al l rights reserved.
from compressed gas cylinders to simulate the air intake system. Four Bronkhorst EL-FLOW Select thermal
mass flow controllers (MFCs) supplied nitrogen with a purity of N4.8 and oxygen with a purity of N5.0 to
the thruster and cathode.
MFC MFC MFC MFC
O2N2O2N2
A
CR F G
Screen Grid
Cathode
Accel.
Grid
Beam
Target
T RF G Thruster
Facility
Isc IAc ICol
Collector
P SCDC in
P ST DC in
P Ssc P SAc
RF Coils
Figure 4. Full electrical and fluidic scheme showing Mass Flow Controllers (MFCs), power supplies (PS ) and
Radio Frequency Generators (RF Gs) described in the text.
The two radio frequency generators used for the cathode and thruster are denoted CRF G and T RF G
respectively. Each RFG was powered by a dedicated air-side Elektro-Automatik PSI9000T DC power supply
shown as P SCDC in and P ST DCin in Fig.4. For all experiments, both RFGs were mechanically integrated
onto their respective propulsion devices and operated in vacuum with a water-cooled plate used for thermal
stability. A Magna Power XR Series power supply was used to sustain the thruster screen grid potential,
shown as P SSc in Fig.4. The thruster acceleration grid used a Keysight N7500, denoted P SAc.
The conditions at ULEO orbits can change drastically and extremely quickly. Of critical importance to
an ABEP system is the ability to operate over a wide range of conditions, including a wide range of O2:N2
ratios. To test this, we monitor the thruster stability when changing the O2:N2ratio while the device is
extracting ions. Note that our intake system has been shown to recombine atomic species such that we are
able to replicate this flow using O2and N2.
The cathode was operated with the RFG at a fixed DC-input power of P SCDC in=120 W and a fixed mass-flow
rate of 0.35 sccm of N2and 0.35 sccm of O2to give an overall mass flow rate of 0.7sccm (1.54x10−2mgs−1).
The thruster was supplied with a 50:50 mixture by mass of oxygen and nitrogen, with a combined thruster
mass flow rate varying from 9 sccm to 14 sccm (0.197 mgs−1to 0.307 mgs−1). The RF power of the thruster
was controlled throughout the experiment by changing the DC-input power over the range of P ST DC in=140 W-
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324 W. Experiments were conducted with the screen voltage VSc=2000 V to demonstrate high specific impulse
mode, and at VSc =1600 V to demonstrate high-thrust mode. At each screen voltage, the thruster was oper-
ated at a fixed acceleration voltage of VAc=-300 V as the mass flow rate and RF power were changed.
The cathode was positioned 2 mm radially from the thruster with an angle of 40◦from parallel. This can be
seen in Fig.5. We explicitly ground the cathode structure, the thruster structure, both RFG housings and
the downstream beam-target to the facility. The current expelled from the ion collector within the cathode
was monitored using a Keysight U1252B Multimeter. After pumping down and completing electrical and
fluidic tests, the thruster and cathode were ignited and brought up to thermal equilibrium. The thruster
was brought to each set-point and allowed to reach stable operation before data was acquired.
Figure 5. Photographs of the experiment as shown from the front (A) and from above (B) showing the
thruster and cathode installed withing the facility. Some sensitive artefacts of the hardware have been have
been censored.
Indirect thrust measurements were performed to calculate the thrust produced from electrical measure-
ments of the thruster. This technique is widely used for GIE characterisation given its high accuracy, and
the lack of dedicated thrust measurement hardware.16 The general formula for thrust Tof a GIE is taken
from Goebel and Katz 2008,17 given by:
T=γr2mI
eIbpVSC ,(1)
where mIis the ion mass, assumed here to be 4.98x10−26 kg for our 50:50 mix of N2:O2by mass. eis
the charge of the electron, Ibis the beam current, which is the difference between the screen current and
the acceleration current: IB≡IS C −IAC , and VSC is the screen voltage. The correction term γaccounts
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Copyright 2024 by the Electric Rocket Propulsion Society. Al l rights reserved.
for a non-ideal beam divergence and the presence of multiple charged ions to which we give a conservative
estimate of γ= 0.97. We disregard any data for which the acceleration current increases above 4% of the
screen current.
Specific impulse is calculated in the usual manner:
Isp =T
˙mTg0
,(2)
for thruster mass flow rate ˙mTand acceleration due to gravity at the Earth’s surface g0.
The power consumption of the thruster, PT otal, is the sum of the DC power supplied to the RFG PT DCin ,
the screen power PSc, and the acceleration power PAc:
PT otal =PT DCin +PSc +PAc
=|VT DCin AT DC in|+|VSc ASc |+|VAcAAc |,(3)
where the subscripts T DC in denote thruster RFG DC input, S c is screen power supply, and Ac is the
acceleration power supply. Note that this does not account for the cathode power consumption.
We use the term thrust-to-power-ratio TTPR to describe the system-level cost of generating thrust in
mN/kW:
TTPR =T
PT otal
,(4)
V. Results
The thruster and cathode were able to consistently ignite and extract ions for the duration of the exper-
iment without the addition of noble gases. Changing the O2:N2ratio during operation had little effect on
their operation, with rapid changes from 100% oxygen to 100% nitrogen occurring without disruption to ion
extraction on the order of seconds. Changing from oxygen to nitrogen without altering other parameters had
a very minor impact on thrust and Isp, with oxygen demonstrating better performance by less than 10%.
The system is shown firing in Fig.6.
The operational range of the thruster is demonstrated by the thrust versus specific impulse plots at a
constant mass flow rate and at constant total power in Fig.7 for two screen voltages. For all these data,
the cathode was operated at PCDCin = 120 W and a mass flow rate of 0.35 sccm O2+ 0.35 sccm N2. As
expected, a greater specific impulse is observed at the higher screen voltages in all cases. The higher screen
voltage allows the thruster to operate at greater total powers, thereby achieving higher overall performance.
A wider range of operation is observed for high mass flow rates, as the plasma can be sustained over a wider
range of powers. For constant power operation, the widest operational range is found at moderate powers.
The TTPR is presented as a function of the total thruster power in Fig.8. Results for two screen volt-
ages are presented: VS c=1600 V and 2000 V. Higher TTPR is achieved with VSc at 1600 V, and as expected,
the TTPR tends to decrease as the RF power, and hence the total thruster power, increases. Fig.9 shows
the total thruster power as a function of mass flow rate at a given thrust.
VI. Discussion and Future Work
Let us consider an ABEP ULEO mission in terms of momentum exchange. In order to compensate fully
for drag, the momentum of the particles colliding with the spacecraft must be less than the momentum that
the thruster imparts on the ejected particles:
vA˙mA≤vex ˙mT,(5)
where:
8
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Copyright 2024 by the Electric Rocket Propulsion Society. All rights reserved.
•vAis the velocity of the incoming atmospheric particles, which is the orbital velocity of the spacecraft
at ULEO, approximately 7800 m·s−1.
•˙mAis the mass flow rate of incoming atmospheric particles which impact the entire satellite and solar
arrays.
•vex is the mean exhaust velocity of particles expelled by the thruster.
•˙mTis the mass flow rate of the thruster, which is the mass flow rate successfully collected by the
intake.
Figure 6. The thruster and cathode operating on atmospheric propellants. The thruster is shown here
operating on 100% N2(A), an equal fraction of N2and O2by volume (B), and on 100% O2(B). In all cases,
the thruster was operated at a mass flow rate totalling 10 sccm, with a constant thruster DC input power
of PT DCin = 240 W, screen voltage of Vsc = 1600 V, and cathode DC input power of PCD Cin = 120 W. The
cathode consumed 0.35 sccm of O2and 0.35 sccm of N2. All photographs were taken with a white balance of
4200 K.
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Copyright 2024 by the Electric Rocket Propulsion Society. All rights reserved.
Figure 7. Thrust as a function of specific impulse at a constant flow rate (top) and total power (bottom).
Data at a screen voltage of VSc = 1600 V is shown on the left, and at VSc = 2000 V is shown on the right.
We can determine ˙mTas a fraction of the total incoming mass flow rate from the product:
˙mT=α η ˙mA,(6)
where αis the fraction of total incoming particles that enter the intake, and ηis the collection efficiency of the
intake, namely the fraction of particles entering the intake that are successfully collected. Our simulations
have indicated that up to 50% of incoming particles do not enter the intake, but instead impart a drag on
lateral surfaces such as solar arrays. For this reason, we assume α= 0.5. Our intake system is able to provide
a collection efficiency of η= 0.35, which is comparable to those found in the literature.10, 18 We substitute
equation 6 into 5, to give us the following requirement for drag compensation:
vex ≥44571 ms−1(7)
or when expressed as a specific impulse:
Isp ≥4543 s(8)
10
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Figure 8. Thrust-to-power ratio at a screen voltage of 1600 V (left) and 2000 V (right) as a function of total
thruster power at a given mass flow rate.
Figure 9. Total power as a function of mass flow rate at a given thrust for screen voltages of 1600V (left) and
2000 V (right).
Under these assumptions, any specific impulse greater than 4543 s will generate more thrust than drag, al-
lowing for sustained ULEO operation through ABEP. As shown in Fig.7, this specific impulse requirement is
met across much of the operational space surveyed in this campaign. We are currently undertaking further
developments, which we anticipate will enhance the performance of our electric propulsion system, thereby
expanding the operational range of our ABEP platform in ULEO.
We aim to demonstrate ABEP in ULEO with our technology demonstration mission next year. Before
this the subsequent generation of our ABEP system will undergo testing in the third quarter of this year
under contract with ESA (contract number: 4000144449) to assess system performance and monitor erosion
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within a larger vacuum facility in our space laboratory.
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