Feasibility of a Helium Closed-Cycle Gas Turbine for UAV Propulsion

Abstract

Featured Application
UAV Propulsion System.

Abstract
When selecting a design for an unmanned aerial vehicle, the choice of the propulsion system is vital in terms of mission requirements, sustainability, usability, noise, controllability, reliability and technology readiness level (TRL). This study analyses the various propulsion systems used in unmanned aerial vehicles (UAVs), paying particular focus on the closed-cycle propulsion systems. The study also investigates the feasibility of using helium closed-cycle gas turbines for UAV propulsion, highlighting the merits and demerits of helium closed-cycle gas turbines. Some of the advantages mentioned include high payload, low noise and high altitude mission ability; while the major drawbacks include a heat sink, nuclear hazard radiation and the shield weight. A preliminary assessment of the cycle showed that a pressure ratio of 4, turbine entry temperature (TET) of 800 °C and mass flow of 50 kg/s could be used to achieve a lightweight helium closed-cycle gas turbine design for UAV mission considering component design constraints.

Full text

applied
sciences
Article
Feasibility of a Helium Closed-Cycle Gas Turbine
for UAV Propulsion
Emmanuel O. Osigwe 1, 2, * , Arnold Gad-Briggs 1,2 and Theoklis Nikolaidis 2


Citation: Osigwe, E.O.; Gad-Briggs,A.;
Nikolaidis, T. Feasibility of a Helium
Closed-Cycle Gas Turbine for UAV
Propulsion. Appl. Sci. 2021,11, 28.
https://dx.doi.org/10.3390/app11
010028
Received: 30 November 2020
Accepted: 18 December 2020
Published: 22 December 2020
Publisher’s Note: MDPI stays neu-
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Creative CommonsAttribution (CCBY)
license(https://creativecommons.org/
licenses/by/4.0/).
1EGB Engineering, Southwell NG25 0BB, UK; a.a.gadbriggs@cranfield.ac.uk
2Centre for Propulsion Engineering, School of Aerospace Transport and Manufacturing (SATM),
Cranfield University, Cranfield MK43 0AL, UK; t.nikolaidis@cranfield.ac.uk
*Correspondence: eosigwe@cranfield.ac.uk
Featured Application: UAV Propulsion System.
Abstract:
When selecting a design for an unmanned aerial vehicle, the choice of the propulsion
system is vital in terms of mission requirements, sustainability, usability, noise, controllability,
reliability and technology readiness level (TRL). This study analyses the various propulsion systems
used in unmanned aerial vehicles (UAVs), paying particular focus on the closed-cycle propulsion
systems. The study also investigates the feasibility of using helium closed-cycle gas turbines for UAV
propulsion, highlighting the merits and demerits of helium closed-cycle gas turbines. Some of the
advantages mentioned include high payload, low noise and high altitude mission ability; while the
major drawbacks include a heat sink, nuclear hazard radiation and the shield weight. A preliminary
assessment of the cycle showed that a pressure ratio of 4, turbine entry temperature (TET) of 800
◦
C
and mass flow of 50 kg/s could be used to achieve a lightweight helium closed-cycle gas turbine
design for UAV mission considering component design constraints.
Keywords: UAV; propulsion system; closed-cycle engines; gas turbines; mission requirement
1. Introduction
In the context of this paper, the term unmanned aerial vehicles (UAV), referred to
as uncrewed aerial vehicles, are aircraft that use aerodynamic forces to provide lift either
directly or by forward motion, with no human on board and controlled autonomously
or remotely. Hence, throughout this paper, the term uncrewed aerial vehicle is used
interchangeably with remotely piloted vehicles (RPV) and uncrewed air system (UAS).
UAV, UAS and RPV have been around since World War II, and there is growing interest
in the use of uncrewed aircraft for special mission requirements [
1
], due to the advent of
better technologies to support the intelligence of controlled UAVs. To this end, there are
projections for significant growth in the use of UAVs in both civil and military applications
due to the added economic and technological value related to them [
2
,
3
]. Despite these
projections, there are still many challenges facing the industry, with the salient ones being
the choice of the propulsion system, the energy source, the conversion system, the fluid
media, and the thrust converter. The considerations of energy source, conversion system,
fluid media and thrust converter are linked to the choice of the propulsion system.
The choice of the propulsion system is vital in terms of sustainability, usability, noise,
controllability, reliability, and technology readiness level (TRL). Unlike the crewed aircraft
vehicles, the design for UAVs will also need to meet long-endurance criteria, a duty cycle
portrayed by heavyweight and high-altitude flight, compactness, and payload require-
ments. Hence, a propulsion system that does not meet these demands will not make any
economic sense and will be less applicable for either military or civil usage. With sus-
tainability, the goal will be to have a system that would not have a severe impact on the
environment, as well as achieving the objective of a net-zero carbon footprint.
Appl. Sci. 2021,11, 28. https://dx.doi.org/10.3390/app11010028 https://www.mdpi.com/journal/applsci

Appl. Sci. 2021,11, 28 2 of 18
The complexity of selecting and exploiting the propulsive system capability to meet
all assumed requirements is not an easy task, but one that requires continuous optimisation.
Within this context, several propulsion systems have been considered as part of ongoing
innovative research programs. A broad spectrum of these initiatives includes open, closed
and semi-closed propulsion systems such as the use of reciprocating engines (diesel en-
gines), Rankine, Stirling, and Brayton cycles (Turbojet, Turboprop, Turbofan), as well as
battery and fuel-cell, powered propulsion systems. For example, in the work of Rogers, [
4
],
the author proposed the use of a small turbofan or turbojet for UAV propulsion with the
aim that a decrease in both the engine size and frontal area would decrease fuel consump-
tion and improve engine affordability. The work addresses the drawback of remarkably
high fuel consumption associated with the use of turbojet or turbofan. An example of a
Turbojet UAV is the Bombardier CL-289 [
1
] developed by Dornier GmbH in Munich as a
tri-national project between Canada, France and Germany with output thrust of 236 kN
and the Rolls-Royce RA-4a Global Hawk [1] an example of Turbofan UAV.
Similarly, Andrei et al. [
5
] discussed the modelling challenges for turbojet UAV using a
one-dimensional correlation model to access the performance and operational requirements
for a UAV mission. There are also several proposals to use fuel-cell and battery-powered
propulsion recorded in the works of Osenar et al. [
6
] and Gonzalez-Espasand et al. [
7
].
The stated advantages of the fuel cell-based system include higher efficiency and zero-
emission compared to fossil fuel technology. However, the high cost of production and
increased sensitivity to fuel contamination could create additional complexity. As the drive
for a reduction in the cost of lithium is sustained and technology to reduce the mass-to-
energy ratio to less than a quarter, battery-powered propulsion system could lead the UAV
missions. Additionally, Tacconi et al. [
8
] compared the use of a semi-closed cycle to open
Brayton cycles for medium-altitude UAV mission. The result from their work showed a
slight performance improvement with the semi-closed cycle compared against the large
open cycles. This is because the semi-closed cycle offered a compact and efficient design
for medium-altitude UAV missions [9].
The use of reciprocating internal combustion engines (ICE) has also been documented
in the works of Harper and Jansen [
10
], Adamski [
11
] and Marcellan, [
12
], which is the most
common form of the propulsion system used for the UAVs, ranging between
1–150 kW
.
The reciprocating engines are widely used because of weight and high power to weight
ratio, but its disadvantage is hinged on its noisy operation, vibrational challenges at high
revolution per minute and cooling requirement. The reciprocating piston engine has
many pressurised seals that often fails, leading to significant loss of power or engine
failure. An example of the UAVs with reciprocating ICE includes the General Atomics
MQ-1 Predator developed by General Atomics in Poway, US. The UAV weighing 68 kg
(engine weight) with 115 horsepower output [
1
]. Others are the Northrop Grumman
RQ-5A Hunter developed in Fort Hood, Texas, Insitu ScanEagle A [
13
] and Honeywell
Micro Air Vehicle (MAV). Table 1provides an overview of the various UAV propulsion
without mentioning the close-cycle options. The comparison of the closed cycle propulsion
system is highlighted in Section 2of this paper.

Appl. Sci. 2021,11, 28 3 of 18
Table 1. Unmanned aerial vehicle (UAV) propulsion systems overview [12].
Power Plant Altitude
(km) Speed SFC @ Cruise
(kg/kWh)
Power (or
Thrust) (kW)
P/W
(kW/kg)
Electric Motors
Battery
0–15 <Mach 0.6 N/A N/A 0.1–1
Fuel Cell N/A <1
Solar-powered
photovoltaic cell
0.090 (where the
reference fuel is H2) N/A N/A
Reciprocating
Engines
Two Stroke
0–9 <Mach 0.6
0.4–1.2 1–150 0.8–2
Four Stroke 0.3–0.4 0.4–1
Rotary 0.35 15–70 N/A
Gas Turbines
Turbojet 3–20
>370 km/h
N/A
>4 kN
(0.13–4.45 kN
for
target/cruise
missiles)
N/A
Turbofan N/A
Turboprop 7–15 <Mach 0.6 0.3–0.5 >200 3.5–4.8
Turboshaft N/A 4.3–9
Whilst each of these propulsion systems mentioned has its merits and demerits,
which has been discussed extensively in references [
6
,
11
,
14
,
15
], the focus of this paper is on
the closed-cycle propulsion system. The reason for choosing the closed-cycle and especially
the closed-Brayton cycle is because there is no valuable literature to the best of the author’s
knowledge that addresses the possibility of using a helium closed-cycle gas turbine for
UAV application.
In this paper, the authors give a brief performance comparison of the different closed-
cycle engines with regards to UAV propulsion capability and also discusses the potentials
of a closed-Brayton cycle, whilst highlighting various fluid media application and previous
work by other researchers. The novelty of this paper is on the feasibility of helium closed-
cycle gas turbine UAV propulsion, highlighting both the performance and component
design considerations for high altitude and high payload mission requirement. This work
did not consider complex Computational fluid dynamics analysis (CFD) of the system
behaviour as this will be covered in a future publication. Additionally, the scope of this
paper does not cover the transient behaviour of the system as this will be covered in
future work.
This paper is broken down into five sections, with Section 1providing the background
and rationale behind this work as well as the research novelty. Section 2describes the closed-
cycle UAV propulsion and performance comparison of the closed-cycle engines as well as
brief limitations encountered in the use of the cycles for the propose UAV mission. Section 3
provides a conceptual overview of the helium closed-cycle gas turbine for UAV application
whilst showing both performance and component design characteristics. Section 4provides
a discussion of some results presented. In contrast, Section 5gives a conclusion of the work
whilst highlighting some of the research limitation and proposal for future work.
2. The Closed-Cycle UAV Propulsion System
The concept of using the closed-cycle engines to power UAVs is not new, as there
are several studies on the feasibility of these type of propulsion systems for aircraft and
UAVs which developed in the second world war [
10
]. Some of the closed-cycle propulsion
systems reportedly used for UAV mission include the Rankine, Stirling and Brayton
cycle [
15
]. Experience from both experimental and prototype models has shown that the

Appl. Sci. 2021,11, 28 4 of 18
closed-cycle engines are well suited for the UAV applications in several ways because of
their advantage of containing fluid media, which eliminates the noise associated with the
intake and exhaust systems. The high-power conversion efficiencies and power to weight
ratio provided by the closed-cycle engines have demonstrated renewed interest in adapting
the system for space power and UAV application. The outside air density does not have any
significant effect on its output power/thrust requirement, and the operating pressure can
be several times higher than the reciprocating engines which allow for a compact design.
A better contender to the closed-cycle engines in terms of low noise is the fuel-cell
or battery-powered propulsion system. The closed-cycle engine is also suitable for high
altitude UAV missions, as its operation is isolated from the effect of pressure drop occurring
outside the system. Hence, it could benefit from improved efficiency at high altitude be-
cause of the gain in the heat exchanger temperature difference across the system. Although,
there could be some adverse effect in terms of heat transfer, in general, the temperature
difference at high altitude compensates for this shortfall providing an improvement in
the cycle efficiency. Other benefits of the closed cycle engines are identified based on the
various types discussed in Section 2.1 to Section 2.3.
In general, the major drawbacks to the use of the closed-cycle engines are its high heat
rejection requirement and the additional weight from the heat exchangers and condenser,
which could have its effect in a low altitude UAV mission. Other demerits have been
highlighted for the individual closed-cycle propulsion systems. The differences between
the Rankine, Stirling and Brayton cycles are their extra components, the cycle operations,
and the fluid media.
To select any of the closed-cycle engines for a UAV, it must prove to have a competitive
advantage over the other internal combustion engines. It must also be safe and have good
flexibility in terms of fuel and working media selection. The working fluid flexibility is
important because of mission requirements, especially for freezing conditions which the
system must be designed to withstand. All closed-cycle engines compared to the open-
cycle ICE suffer from volume, weight, and capital cost disadvantages; hence, there is a
need to ensure that any selected closed-cycle system demonstrates unique advantages to
compete effectively with off the shelf ICEs. The consequence of increasing weight and
engine complexity could partially offset the benefit of specific fuel consumption (SFC).
For environmentally sensitive operations, the use of closed-cycle engines eliminates the
open-cycle system. The closed-cycle engines align with the goal of resource conservation;
however, the multiple heat exchangers make it more complex.
2.1. The Rankine Closed-Cycle Engines
The Rankine cycle is a vapour power cycle with constant pressure heat addition and
rejection process. This cycle type is famous for power generation, and with a TRL less than
6 for aircraft application as a Besler engine, was the first successfully flown Rankine aircraft.
The Besler engine operated on the Rankine cycle with a two-cylinder V-type engine that
generated up to 150 HP. A steam boiler powered this semi-closed engine which flew 200 feet
altitude and had a reduced noise operation compared with the reciprocating engines. It had
a reciprocating turbine and a small condenser for cooling the steam. However, maintaining
cooling was not a smooth task for this aircraft.
Although the first successfully flown Rankine occurred in 1933, its prospect quickly
diminished due to several constraints associated with the Rankine cycle such as weight,
lower efficiency compared with the competing engines and complex system integration
with the airframe due to heat exchanger requirement. Additionally, its response to changes
in throttle demand is relatively slow, and it has obvious cooling problems. These are among
the many reasons that there are not many ongoing pieces of research that explore the
potentials of steam Rankine cycle for UAVs mission, as known by the authors. However,
rather than using steam, there are propositions to use organic fluid for this type of UAV
altitude propulsion system [
10
,
15
,
16
]. Using organic fluid could potentially allow for
high altitude operation as organic working fluids can handle colder temperature without

Appl. Sci. 2021,11, 28 5 of 18
freezing, and the heat required to power the system is incredibly low. One perceived
advantage of this cycle with organic media is that a small high-pressure electric fluid
pump could entirely replace its compressor as the pump work required is relatively small
compared to other propulsion systems. This is because the compression process occurs in
the liquid phase. Numerical analysis in reference [
16
] shows that high-temperature dry
fluid with trans-critical cycles such as Toluene or cyclohexane could improve the system
efficiency by ~8%.
Interestingly, organic fluid could achieve high efficiency with mid-low temperatures
(between 100–200
◦
C). Operating an organic Rankine cycle (ORC) at temperatures above
300
◦
C could be limited by thermal stability [
17
]. The typical efficiency for this cycle would
be between 20% and 27%. One option to manage to weight issue for the Rankine cycle
design is to ensure the small mass flow rate.
2.2. The Stirling Closed-Cycle Engines
The Stirling cycle engine is a closed-cycle heat engine that undergoes an isothermal
compression and expansion process while rejecting heat at constant volume. This engine
has the potential to operate at high efficiency and capable of utilising a variety of heat
sources and fuels. Other advantages of the engine include low noise and vibration, espe-
cially for the free-piston Stirling engine, clean-burning and low emission, working fluid
flexibility and high reliability. In 1986, McConaghy demonstrated the capability of using
Stirling engines for a UAV mission. His 0.350 kg Stirling engine operating on helium
was able to produce 0.025 kW of power to fly small UAV for 8 min [
15
]. Besides the Mc-
Conaghy UAV test, the US Navy has also implemented several designs for UAV underwater
operations [18].
Designing a Stirling engine could be overly complicated because of the dynamic
behaviour of the mechanism and heat exchangers, as it is not easy to build a heat exchanger
to operate continuously at high temperature. Increasing the working temperature and
pressure increases the power density. However, the maximum temperature is constrained
by material technology, which is currently within the range of 800–900
◦
C. Good Stirling
engine performance requires that the total gas volume be minimised. Maximum perfor-
mance is also achieved with gases that have low specific heat, such that the maximum
possible pressure increase is obtained for a given amount of heat input. Ideally, helium and
hydrogen would be well suited for this application; however, the challenge will be how
to manage seal leakage effectively [
19
,
20
]. One major difference between Stirling and the
internal combustion engine is the rapid variation of power required and the instantaneous
control of fuel supply to meet this demand. For the Stirling engines, there is usually a
substantial time lag between fuel input and power output which make less competitive
to the Brayton closed cycle. Another disadvantage would be the turbomachinery design,
as its TRL is far below its Brayton closed-cycle counterpart. The Stirling engine, similar to
any other reciprocating engine, can never be as light as a Brayton closed-cycle gas turbine
but can compete only by achieving higher efficiency.
2.3. The Brayton Closed-Cycle Engines
The Brayton closed-cycle engines have the same thermodynamic process (isothermal
compression and expansion) as the Brayton open-cycle (turbojet and turbofan). The differ-
ence is that the working media is isolated from the environment within a control volume
and the additional component (heat exchanger) needed to cool the working fluid before
recirculating it through the compressor. The closed-cycle gas turbine could either be cou-
pled directly or indirectly to its heat source, giving it that flexibility in terms of heat source
choice selection. Similar to other closed-cycle engines, it offers compactness and high
reliability as there is no moving part in the machine in high temperature and reduction of
thermal stresses. Another advantage is the flexibility in terms of working fluid selection,
adapting to a variety of high energy and be able to operate efficiently over a wide power
range [21].

Appl. Sci. 2021,11, 28 6 of 18
The Brayton closed-cycle provides a better advantage for small UAV than the open cy-
cle as the efficiency at low-pressure ratio favours the closed Brayton cycle (CBC). The Bray-
ton closed-cycle has a lightweight advantage and higher efficiency in large size compared
to other closed cycle engines. Hence, the CBC engine has many characteristics that are
required for UAV missions. The system is suitable for a wide variety of missions, such as
lunar, deep space, and earth orbit missions. In the 1950s, a program for Aircraft Nuclear
Propulsion (ANP) was initiated by the Atomic Energy Commission in the US to redesign
B-36 which was later renamed NB-36 [
22
–
24
]. By the 1960s, progress was made with
both direct and indirect closed-cycle engines tested, using air as the working fluid [
23
].
Similarly, after extensive experimentation with different engines and transfer systems,
the Soviet Engineers developed the Tupolev (Tu) Tu-95LAL and 34 more research flights
completed between the 1950s and 1960s using the closed-cycle gas turbine technology [
25
].
Feasibility studies using closed-cycle gas turbine engines for UAVs have been documented;
the study [
15
] evaluated the benefit of the closed-cycle gas turbine at high altitude perfor-
mance, operation in polluted air environment using generic thermal heat sources.
The NASA-Lewis Research Centre developed a radioisotope closed-cycle gas turbine
power generation unit in the 1960s, which achieved an output power between 2 kW and
10 kW [
26
,
27
]. Similarly, the NASA Space Station Freedom selected a solar-powered CBC
for space power system application. To this end, there are several pieces of research and
development programs been deployed to show the technology readiness of the closed-cycle
gas turbine in aerospace application [15,24,28].
With the CBC, the layout configuration can be arranged as such that the performance
and thermo-economics are taken into consideration to get the right balance between system
performance and capital cost. Interestingly, the working fluid selection could have an
influence on the physical layout of the CBC engine [
21
]. Due to the self-contained nature of
the CBC, almost all permanent gaseous working fluid can be used, however, with UAVs,
unique criteria are considered such as thermal stability, inflammability, component material
compatibility, and behaviour of fluid at freezing conditions. Several working fluids that
can be considered range from monoatomic inert gases and mixtures, however, this paper
will focus on the use of helium. Helium was selected due to its thermophysical properties,
allowing for compact small UAV design, and would demonstrate high efficiency at a
low-pressure ratio to compete with other propulsion systems.
The major challenge for the CBC design would be the heat source. As previously
mentioned, this could be coupled directly or indirectly, and this can run on fuel ranging
from nuclear fuel, natural gas, or solar. The most promising of these fuels are nuclear and
solar; however, they both have their drawbacks. For this work, the authors have considered
nuclear reactor coupled directly to the CBC.
While alternate propulsion systems may exhibit some of the attributes required for
small UAV missions, this paper provides a balanced view of the merits and demerits of
using helium closed-cycle gas turbine for UAV propulsion in terms of cycle performance
and system design.
3. The Helium Closed-Cycle Gas Turbine UAV Propulsion
The helium closed-cycle gas turbine operates as a closed Brayton engine described in
Section 2.3. The thermo-economic conceptual reason to use the helium closed-cycle gas is
hinged on the need for an aircraft with a high payload, long-duration endurance, and unique
missions for either low altitude high-speed penetrations or high-altitude mission.
To this end, the closed-cycle gas turbine discussed in this paper is coupled directly
to a small, very high-temperature reactor (VHTR) cooled by helium working fluid with
a core outlet temperature simulated between 700 and 950
◦
C. With helium as the coolant
and cycle working fluid, it brings several benefits such as chemical inertness, single-phase
cooling and neutronic transparency, which makes reactivity effect with structural material
low [
29
]. The choice of working fluid also has effects on the turbo set (compressor and
turbine) in terms of the number of stages for the attainment of the required pressure ratio,

Appl. Sci. 2021,11, 28 7 of 18
high cycle efficiency and machine size. The conceptual system configuration shown in
Figure 1consists of a simple single-shaft arrangement connecting the turbo set, the heat
sink (heat exchanger) and the VHTR.
Figure 1. Helium closed-cycle gas turbine configuration.
Understandably, the VHTR could pose several regulation issues due to radiation
hazards; however, the use of a unit shield and an inert gas such as helium could reduce the
radiation effect to satisfactory levels both on the ground and inflight, even under reactor
malfunction conditions. Evolutionary advances in nuclear propulsion—specifically in
reactor compactness and shielding—makes the application of a shielded closed system
possible and payload that could compete effectively with conventional reciprocating piston
engines and open-cycle gas turbines at ranges above 90,000 feet. The simple single-shaft
cycle arrangement described in Figure 1is different from the ERAST (Environmental
Research Aircraft and Sensor Technology) semi-closed cycle configuration discussed in
reference [30].
The altitude, flight Mach number, output thrust and payload requirement have been
specified in Table 2to establish a benchmark for the proposed UAV mission, The mission
requirement for the UAV is defined in terms of operational altitude, range, endurance
and payload. Although the cycle arrangement is still in the conceptual stage, the goal of
this work is to ascertain the feasibility of using the helium closed-cycle gas turbine for the
mission requirement. Table 1provides details of existing UAVs characteristics,
and Figure 2
provides the performance trend of the existing UAVs.
Table 2. Baseline performance data.
Description Values
Cruise Altitude (m) 30,500
Cruise Mach no 0.7
Take-off Mach no 0.4
Compressor pressure ratio 2–5
Core outlet temperature (◦C) 700–950
PC effectiveness (%) 87
Compressor efficiency (%) 90
Turbine efficiency (%) 92
Propeller efficiency (%) 90

Appl. Sci. 2021,11, 28 8 of 18
Figure 2.
Performance trend of existing unmanned aerial vehicles (UAVs) (
a
) Maximum take-off
weight against power (
b
) flight range against endurance (
c
) Span against maximum take-off weight
(d) endurance against maximum take-off weight [1,12].

Appl. Sci. 2021,11, 28 9 of 18
To this end, a pressure ratio of 4 was selected for the gas turbine to keep the turbo set
weight and size within reasonable values. For a simple helium closed-cycle gas turbine,
an increase in pressure ratio peaks between 4 and 6, from whence the cycle performance
begins to retrogress. The core outlet temperature of 800
◦
C was chosen to manage the
cooling requirement of the conceptual design and heat exchanger size. An unlimited range
would offer an attractive and competitive UAVs requirement for today. However, there are
constraints to achieving an unlimited range proposition which is linked to the fuel load
and system weight. One benefit of the closed-cycle system is the freedom in selecting the
cycle pressures and temperatures as this would allow for tailoring and system adaptability
to mission-specific criteria.
3.1. Engine Model Development
A comprehensive engine component model and thermodynamic calculations to define
the system performance was accomplished with an in-house tool developed by the authors,
which is described in details in reference [
31
]. The overall performance of the system is
a function of its components based on the baseline characteristics presented in Table 1.
Changing the operating variables at different altitude describes the off-design model and
component matching procedure. The working fluid model implemented in the in-house
tool is correlated with Cool Prop library [32].
3.1.1. Turbomachinery Model
Dimensionless parameters describe the behaviour of the turbomachinery component.
The performance characteristics are usually plotted as pressure ratio against dimension-
less or corrected mass flow (CMF), corrected speed (CS) and corrected enthalpy (CH).
The following expressions represent the dimensionless parameters:
CMF = W√θ
δ×sR
γ!,CS =N
√θRγ,CH =∆H
√θRγ(1)
where,
θ= T
Tref !,δ=P
Pre f
(2)
The temperature at the exit of the compressor is obtained from the isentropic relation-
ship between inlet temperature, isentropic efficiency, pressure ratio and the ratio of specific
heats, given by the expression:
TCout =TCin +TCin
ηc
PCout
PCin (γ−1
γ)
−1
(3)
where the compressor discharge pressure is derived from the given pressure ratio:
PRc=Pcout
Pcin
=f(CMF,CS)(4)
The Non-Dimensional Compressor Work (NDCW) is derived as:
NDCW =CW
δ√θ=f(CMF,CS)(5)
where the Compressor Work (CW), is a product of the mass flow rate, specific heat capacity
at constant pressure and the overall temperature rise in the compressor.
CW =WCPout Tcout −WCPin Tcin (6)

Appl. Sci. 2021,11, 28 10 of 18
To solve Equations (1)–(5), the input data is obtained from the design point inlet
conditions and compressor performance map.
Similar to the compressor, the performance characteristics of the turbine is described
by several dimensionless parameters. The non-dimensional parameters defined for the
compressor is similar to the turbine. The input data at the design point are the turbine inlet
conditions (mass flow rate, temperature, and pressure), and component efficiency.
The temperature at the outlet of the turbine is obtained from the following expression:
Ttout =Ttin −(Ttin ×ηt)
1−Ptout
Ptin (γ−1
γ)
(7)
The Non- Dimensional Turbine Work (NDTW) can be obtained from:
NDTW =TW
δ√θ=f(Ptout
Ptin ,CS)(8)
where TW is the turbine work, which is expressed as:
TW =WCPout Ttout −WCPin Ttin (9)
Implementing system components pressure losses will mean that the turbine pressure
ratio becomes:
PRt=Pcout
Pcin
(10)
The overall compressor and turbine design assumed a free vortex velocity distribution
with constant outer diameter annulus configuration. The annulus area for the turbo-set
inlet and outlet geometry is given as:
A=W√T
QoKbP=πD2
m
4(11)
where,
Kb= blockage factor
Q0= Non-dimensional mass flow
The number of stages is given by:
Number o f S tages =Cp∆T
∆h(12)
The stage loading and flow coefficient are given as:
Stage load =Cp∆T
U2
m
(13)
Sta ge f l ow coe f f icient =Va
Um(14)
The length of the turbo-set is calculated using Sagerser empirical relationship such
that compressor length is given as [33]:
Lc=Dm×[(0.2 +[(0.234 −(0.218 ∗hub_tip_ratio))] ×NS)] (15)
And the turbine length is given as:
LT=NSh Dt−Dh
0.96×hub_ti p rotor +Dt−Dh
0.96×hub_ti p stator
+(2NS−1)0.4 ×Dt−Dh
0.96∗hub_ti p rotori(16)

Appl. Sci. 2021,11, 28 11 of 18
3.1.2. Precooler Heat Exchanger Model
The precooler is modelled as a heat exchanger using the
E
-NTU method, and a
counterflow shell and tube configuration was assumed. The
E
-NTU method was used since
the inlet condition (temperature and pressure) of the fluid stream can be easily obtained
and simplifies the iteration involved in predicting the performance of the flow arrangement.
This method is fully described in references [
34
,
35
]. The approach also assumes that the
heat exchanger effectiveness is known, and the pressure losses are given.
Therefore, the effectiveness of the heat exchanger is the ratio of the actual heat transfer
rate to the thermodynamically limited maximum heat transfer rate available in a counter
flow arrangement.
ε=Qactual
Qmax
=Chot (Thotin −Thotout)
Cmin (Thotin −Tcol din )=Ccol d(Tcol dout −Tcol din )
Cmin (Thotin −Tcol din )(17)
where Cmin and Cmax are the smaller and larger of the two magnitudes of Chot and Ccold
Chot =(WCP)hot f l uid Stream,Ccold =(W CP)co ld f l uid Stream (18)
Cmin =Chot f or Chot <Ccold
Ccold f or Ccold <Chot
(19)
For counterflow shell and tube heat exchangers, number of transfer unit (
NTU
) is
given by:
NTU =LOGeh2−E(1+C∗−ηHe x
2−E(1+C∗+ηHex i
ηHex
(20)
where
C∗=Ca pacity rate ratio =Cmin
Cmax (21)
ηHex =C∗2+10.5 (22)
Assuming pressure losses, then the exit pressure at design point is given as:
Pout =Pin(1−∆P)(23)
where
∆P
is the percentage of pressure loss specified as the input for each component.
It is important to emphasis that the precooler utilizes air cooling. The code was modelled
to have the working fluid temperature after pre-cooling process equal to the compressor
entry temperature; although, in a real scenario, the temperature may vary a little.
3.1.3. The Thrust Converter System
The purpose of the thrust converter is to translate the torque produced by the gas
turbine into thrust.
Thus,
Pr=V∞×Thrust (24)
where,
Pr=Power requiredV∞=Free stream velocity
The thrust converter efficiency is given as:
ηthr =V∞×Thrust
Pr(25)
And thrust for steady flight is a function of weight, drag and lift coefficient given as:
Thr ust f or steady f l ight =CD
CL ×Weight (26)

Appl. Sci. 2021,11, 28 12 of 18
3.1.4. Weight Model
A significant design goal is to minimise the propulsion system weight. Hence the
overall system weight is an aggregate of the component weight.
Weightengine =WeightC+WeightT+WeightVHTR +WeightPC (27)
The weight of the turbomachinery component is estimated according to the NASA
weight analysis of turbine engine model [
36
] and the precooler weight model as described
in the work of Kakac and Liu [
34
] and Shan and Sekulic [
35
]. A breakdown of the weight
contribution shows that the significant contributors to the system are the nuclear reactor
and shield, and the heat sink.
3.1.5. Endurance and Range
Endurance is the maximum length of time the UAV can spend in cruise, which is a
function of the lift, drag, system weight and fuel load. The Breguet range equation was
used to calculate endurance and range [
37
], which depends whether the thrust is propeller
or jet driven. The relationship is expressed as:
Range =ηthr
Wf×CD
CL ×ln Take o f f Wei ght
Landing Weight (28)
Endurance =ηthr
Wf×p2ρSCL1.5
CD ×Landing Weight−0.5 −Takeo f f W eight−0.5(29)
4. Discussion
4.1. System Performance
The helium closed-cycle gas turbine looks suitable for UAV missions. It is compact,
efficient and offers high reliability since it requires only one rotating equipment (turbo set).
The choice of helium as a working fluid enhances the overall system design, and the
operating envelope does not adversely affect the working fluid. The top-level requirement
of high payload, high altitude, long-endurance, and range are representative of the mission
requirements which could be reasonably expected but sufficiently stringent to test the real
application of the closed-cycle gas turbine system for UAV propulsion.
The system mass flow rate, temperature, and pressure, as well as component char-
acteristic, were simulated for targeted thrust and system performance based on pressure
ratio, turbine entry temperature (TET), and altitude variation. The choice of mass flow has
a direct relationship with the thrust, the power required, weight and system size. In the
parametric assessment, a fixed mass flow of 50 kg/s was chosen as a representative for a
thrust and payload requirement. The mass flow value can be optimised in relationship
with TET to obtain a lightweight UAV. The proposal to use a single-shaft arrangement also
reduces the complexity of weight. The payload estimate refers to equipment for which the
UAV provides platform and transportation. In this context, the payload is a measure of the
size, weight, and power available to perform take-off, fly around and landing.
Figure 3shows the strong relationship between specific fuel consumption and the
TET and pressure ratio. At take-off, a simultaneous increase in the TET from 700 to 950
◦
C,
and pressure ratio both decreases the SFC requirement. However, there is a limit for an
optimum UAV mission requirement considering the implication of size for pressure ratio
and the cooling need for both turbine and the precooler heat exchanger. A temperature
between 700 and 800
◦
C seems to be conservative to high payload mission as the TET have
a direct effect on the power required and thrust.

Appl. Sci. 2021,11, 28 13 of 18
Figure 3.
Parametric analysis of cycle at Maximum Take-off Thrust, Weight and Mach number 0.4 at different turbine entry
temperatures (
a
) specific thrust against pressure ratio, (
b
) specific fuel consumption against pressure ratio (
c
) Thrust against
pressure ratio.
The parametric study shows the system performance peak pressure ratio of 4 and a
conservative TET of 800
◦
C was chosen as the final baseline for the analysis. A pressure
ratio of 4 was selected for the gas turbine to keep the turbo set weight and size within
reasonable values. An increase in power density for the helium closed-cycle gas turbine
provides a small system size and less cycle pressure losses compared with other propulsion
systems. However, a maximum take-off thrust of 110kN at a TET of 800
◦
C was chosen due
to cooling requirement and heat sink design constraints. In the preliminary design, one-third
of the pumping power is assigned to the reactor and one-sixth to the precooler. With much
high thermal conductivity and permissible flow velocity; helium heat exchangers are compact.
The problem with Ranking rejection at almost constant temperature is in contrast with the
closed-cycle gas turbine as heat is rejected over a wide range of temperature. The success
of the closed-cycle application for UAV is tied to the heat sink. The radiator area should be
compact and efficient at rejecting the waste for recycling of the helium working fluid into the
compressor. In the preliminary analysis, pressure losses were not considered.
Table 3shows the system performance at 30,000 m and a Mach number of 0.7.
Table 3. System performance at 30,000 m.
Descriptions Value
Altitude (m) 30,000
Mach number 0.7
Thrust (kN) 74.847
Specific fuel consumption (SFC) 3.03
Turbine entry temperature TET (◦C) 822
The control at different thrust requirement can be easily achieved with pressure level
control, heat source and by-pass control. For most cases, pressure level control is performed
at almost constant SFC and a constant temperature in all component. Parametric sizing
and weight estimates were made for the turbomachinery and precooler and reactor.

Appl. Sci. 2021,11, 28 14 of 18
4.2. Component Design Constraints
4.2.1. Turbo Set
The turbo set design relies on the technology available mainly in the area of the perfor-
mance, material, design and fabrication method. The choice of the turbo set (compressor
and turbine) type, and pressure ratio is an important design consideration for simplicity
and cost reduction. In the proposed model, a compressor pressure ratio between 3 and 5
seems to be a modest value because of the design difficulties associated with the thermo-
dynamic behaviour of the helium and the direct effect of the pressure ratio to compressor
size. The relative engine length tends to increase in a regular manner with pressure ratio.
However, the low molecular weight of helium causes a high acoustic velocity, hence, could
eliminate the Mach number constraints on the aerodynamic design of the turbo-set com-
ponents. Additionally, the enthalpy and specific heat change would allow for high blade
velocity (m/s) design to reduce the number of stages. The challenge then becomes how to
induce the highest possible velocities that the blade would allow since the stage loading
factor is inversely proportional to the square of the blade velocity.
Designing for a low noise propulsion system, one would prefer to slow the rotational
speed of the shaft while maintaining power so that mechanical noise and propeller tip
velocity are both decreased. The helium compressors design for the closed-cycle propul-
sion is characterized by high hub-to-tip ratios, low aspect ratios and small blade height.
The overall design of the compressor uses a free vortex velocity distribution with constant
inner diameter for all stages.
The properties of the helium would also affect the turbine design characteristics,
such as the compressor. Since it is desirable to have a high blade speed as possible to
reduce the number stages, the problem will be on the turbine first stage due to the rotor
blade temperature will be at a maximum. The overall design of the turbine is based on
free vortex with constant mean diameter for all stages. With a turbine entry temperature
between 700 and 850
◦
C, the turbine blade cooling requirement could be easily managed,
and the blade life could be met with a bleed flow of ~1%.
Another design challenge will be on the seal to reduce leakage, accommodate nonuni-
formities between mating parts and to accommodate differential expansion and thermal
distortion. The use of bearing stub shafts at each rotor end could eliminate unnecessary
weight while maintaining bending stiffness. Table 4shows the design characteristics for
both the compressor and turbine.
Table 4. Turbo set design characteristics.
Description Compressor Turbine
Inlet temperature (◦C) 300 800
Inlet pressure (kPa) 303.975 303.975
Mass flow (kg/s) 50 50
Rotational speed (rpm) 16,000 16,000
Pressure ratio 4 4
Number of stages 4 2
Um (m/s) 302 302
Temperature ratio 1.81 1.65
Length (m) 0.53 0.29
Relative Weight 0.6 0.7
4.2.2. Reactor
The reactor module comprises of the reactor, the primary circulator, the reactor shut-
down cooling system with the main cylindrical section housed in the reactor vessel. The de-
sign goal will be to minimise the reactor volume for a given reactor power. However,
this is limited to temperature materials, heat transfer rates and radiation considerations.

Appl. Sci. 2021,11, 28 15 of 18
To design a compact reactor for the UAV, one must also consider how to manage the fuel
burn-up, the fission products and elements.
The major single item weight of the reactor heat source is the weight of the reactor
shield, which is required to protect against the damaging effects of the nuclear radiation.
Although for a given reactor, the shield weight per unit power output tends to decrease as
the reactor output power increases, this would have an impact on the size of the reactor
and could add another constraint for a lightweight design. With new technology available,
there has been consistent progress made with compact reactor design and reduction in
the shield weight. A significant success can be achieved with justifiable efforts aimed at
minimising the weight penalty of the reactor and the overall system.
4.2.3. Precooler
The challenge with closed-cycle cycle system is the high heat sink requirement. This is
an addition to drag and get worse at the climb phase. The design of the heat sink is
critical to the success of the helium closed-cycle gas turbine. An involute spiral multi-
pass counter-cross flow heat exchanger can be used to achieve efficient cooling. The cold
side inlet Mach number and the number of tubes are essential factors to consider when
designing the precooler. A longer tube design will result in additional weight, and a smaller
tube diameter will perform better thermally. The heat rejection effectiveness improves
for the closed-cycle propulsion system with altitude because of increasing temperature
differentials.
5. Conclusions
The feasibility study on the helium closed-cycle gas turbine shows it could be suitable
for high payload, low noise, and high attitude mission requirement. The design require-
ment differs considerably from the ICEs, and other UAV closed-cycle propulsion systems.
However, the major drawback to the application of the helium closed-cycle propulsion
system includes weight associated with the reactor, and reactor shield, the heat rejection
and obvious stringent certification issues.
The goal of achieving a lightweight propulsion system is hinged on the system config-
uration and parameter selection. The preliminary assessment carried out showed that a
pressure ratio between 3 and 5 and TET between 700 and 800 ◦C would give an optimum
system performance while considering all possible constraints such as weight, size, mate-
rial and cooling requirement. The baseline representative assessment with a pressure ratio
of 4 and TET of 800 at MTOW and thrust shows a compact turbo set with four stages of
axial compression and two stages of the expansion to achieve a thrust of 110 kN. The value
may seem overly exaggerated since pressure losses and other component losses were not
considered during the assessment.
The scope of this paper does not cover a detailed geometry design and transient
behaviour of the system as this will be covered in future work, as well as CFD assessment.
Author Contributions:
Conceptualization, E.O.O. and A.G.-B.; methodology, E.O.O.; software,
E.O.O. and A.G.; validation, E.O.O., A.G.-B. and T.N.; formal analysis, E.O.O.; investigation, E.O.O.;
resources, A.G.-B., and T.N.; data curation, E.O.O.; writing—original draft preparation, E.O.O.;
writing—review and editing, E.O.O.; visualization, E.O.O.; supervision, T.N.; project administration,
A.G.-B.; funding acquisition, A.G.-B. All authors have read and agreed to the published version of
the manuscript.
Funding: This research received no external funding.
Institutional Review Board Statement: Not Applicable.
Informed Consent Statement: Not Applicable.
Data Availability Statement: Not Applicable.
Conflicts of Interest:
Authors confirm that there is no conflict of interest with regards to publishing
this article.

Appl. Sci. 2021,11, 28 16 of 18
Nomenclature
Notation
AFlow annulus area m2
CD Drag coefficient
CL Lift coefficient
CH Corrected enthalpy
CpSpecific heat capacity, J/kgK
CS Corrected speed
CW Compressor work, W
DmMean diameter, m
KbBlockage factor
LcCompressor length, m
LTTurbine length, m
NsRotational speed (rpm)
NDCW Non-dimensional compressor work
NDTW Non-dimensional turbine work
P Pressure, Pa
PrPower required, kW
SWing platform area, m2
SFC Specific fuel consumption
T Temperature, ◦C
TET Turbine entry temperature, ◦C
TW Turbine work, W
UmMean speed m/s
V∞Free stream velocity, m/s
W Mass flow kg/s
WfFuel flow, kg/s
ρDensity, kg/m3
ηthr Propeller efficiency
∆hEnthalpy change
∆PPressure change
∆TTemperature change
Abbreviations
ICE Internal Combustion Engine
CMF Corrected Mass Flow
C Compressor
CBC Closed Brayton cycle
CFD Computational fluid dynamics
DP Design point
OD Off-design
ORC Organic Rankine cycle
PC Precooler
PR Pressure ratio
SF Scaling factor
T Turbine
TRL Technology readiness level
UAV Uncrewed/unmanned aerial vehicles
VHTR Very high-temperature reactor
Subscript
cCompressor
tTurbine
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