![Compressor map [22]](https://www.researchgate.net/profile/Francis-Oppong/publication/317312410/figure/fig3/AS:500786317545472@1496408170014/Compressor-map-22_Q320.jpg)
![Turbine map [22]](https://www.researchgate.net/profile/Francis-Oppong/publication/317312410/figure/fig4/AS:500786319036418@1496408170086/Turbine-map-22_Q320.jpg)
![Compressor stage [41]](https://www.researchgate.net/profile/Francis-Oppong/publication/317312410/figure/fig5/AS:500786316955648@1496408170136/Compressor-stage-41_Q320.jpg)
Content uploaded by Francis Oppong
Author content
All content in this area was uploaded by Francis Oppong on Jun 02, 2017
Content may be subject to copyright.
Correspondent Author: Francis Oppong, University of Stellenbosch,
South Africa, Department of Mechanical and Mechatronic Engineering.
Email: oppong_francis@outlook.com
Micro gas turbine (MGT) engines are ex-
tensively used in Remote Control (RC) model airplanes
or unmanned aerial vehicles. They are also used in
small combined heat and power generation plants and
hybrid electric vehicle applications as well as auxiliary
power units (AUPs) for modern aircrafts. Notably, the
overall performance of the engine depends on the per-
formance of its individual components. The effective-
ness of MGT is limited due to non-linear behaviour of
the engine elements. Component matching analysis of
MGT engine showed that the inefficiencies of the en-
gine gas generator units affect the mutual performance
of one another and the overall engine output. Thus, in-
sufficient turbine work and low component pressure ra-
tio and efficiency results in excessive fuel consumption
and high turbine inlet and exhaust gas temperatures.
Moreover, heat transfer and/or thermal losses, low
Reynolds number as well as combustion instabilities
such as low combustion efficiencies, flame stability,
and low-pressure drop accounts for MGTs poor perfor-
mance. The objectives of the study are to provide a fun-
damental overview of the methodologies used in the
modelling of micro gas turbine performance. The liter-
ature includes works on various aspects of micro gas
turbines. The engine thermodynamic cycle analyses,
performance modelling with component maps and the
drawbacks associated with the downscaling of conven-
tional gas turbine into smaller engines were described.
The second part of the paper discussed the various
methodologies that have been employed in investigat-
ing and improving MGTs performance output. The pa-
per concludes with summary observations in relation to
micro gas turbine performance modelling as well as
some opportunities, which merit further research and
the challenges to be resolved.
Micro gas turbine, performance, thermal
efficiency, output
The use of micro gas turbine (MGT) engines is be-
coming increasingly popular in the commercial avia-
tion and hobby industries. They are employed in un-
manned aerial vehicles (UAVs) used in missions re-
lated to national security, telecommunications, real-
time reconnaissance, remote sensing, crime fighting,
disaster management, agriculture and election monitor-
ing [1]–[3]. MGTs are also used in hybrid electric ve-
hicles and small combine heat and power generating
plant applications as well as auxiliary power units
(APUs) for modern aircraft [4]–[6]. They are suitable
for such applications due to their high power to weight
ratio, multi-fuel capability and simple design, low en-
ergy costs and emissions [7]–[9]. MGTs can be re-
garded as a prospective and compact competitor to the
other propulsion system power supplies such as recip-
rocating engines and battery cells [5], [10]. These en-
gines have interrelated components, which have non-
linear characteristics. Therefore, the overall engine per-
formance depends on the individual engine element's
performance.
Some of the technical barriers that impede the en-
gine performance include better understanding of the
aerodynamics of the engine components. The flow is
characterised by viscous, compressible, and turbulent
phenomena. This includes thermo-mechanical charac-
teristics of the engine allied with relatively high oper-
ating temperatures, very low engine component pres-
sure ratios and efficiencies, and high rotational speeds
[11]. Additionally, heat transfer, tip clearance losses,
and low Reynolds number related problems due to the
engine size affect the engine output. Lee et al. [12] have
analysed micro gas turbine component characteristics
changes on the engine performance using experimental
testing. Further, Yoon et al. [13] developed a model to
investigate the performance deterioration of MGT ele-
ments. They evaluated the components characteristic
effects on the engine performance output. Experimental
and theoretical measurements of MGT thermo-me-
chanical parameters have been investigated by Bakalis
and Stamatis [14]. In the investigation by Xie et al.
[15], a three-dimensional fluid thermal structural cou-
pled analysis for a radial inflow micro gas turbine is
conducted. Tip clearance losses contributed to inhomo-
geneous flow in the impeller blades. The hot turbine
heat flux heats the cold air in the compressor via con-
duction, hence reducing the mass flow rate through the
engine, efficiency, and pressure rise.
Combustion interrelated drawbacks such as effec-
tive design of combustion chambers to achieve flame
stability and recirculation zones in the chamber also
contributes to MGTs low performance outputs. The im-
plementation of different and new designs of combus-
tors has been investigated as well as the use of alterna-
tive source of fuels to assess the engine output. Yang et
al. [16] have examined the practicality of biogas on the
performance of micro gas turbine annular combustor
using experimental and numerical investigations. They
measured thermal and electrical efficiency of 23% and
10%. According to the writers, the numerical analysis
showed that raising the combustor outlet temperature
could improve the engine performance. C. Yang et al.
[17] have showed the effects of using fuels with low
heating values on the performance of an annular MW-
54 micro gas turbine. They sought to confirm that there
are no hot spots in the liner walls and measure the max-
imum exhaust temperatures when the engine operates
under variable conditions. From the above brief discus-
sions, the geometry of MGTs and its accompanied de-
ficiencies in modelling the engine performance remain
as a challenge to researchers.
The main goal of this paper is to present detailed
overview of the performance investigation techniques
employed in modelling the performance output of small
gas turbine engines. This involve references to numer-
ous treatises in which micro gas turbine performance
analyses has been described and reviewed to different
extents.
The micro gas turbine engine shown in Fig. 1 is a
single spool (shaft) rotary engine that extracts energy
from the flow of combustion gases to produce thrust or
power. MGTs have a thermal efficiency of between
10% and 30% [12], [18] and power and thrust in the
ranges of 15-200 kW [11] [19] and 30-200 N [20] at a
rotational speed of 20 000-150 000 rpm [1]. The engine
consists of a centrifugal compressor at the front of the
engine, equipped with a radial or crossover vanned dif-
fuser, an annular straight through, or reverse flow com-
bustor fixed between the compressor and the axial flow
turbine, and a fixed convergent propelling nozzle at the
back of the engine.
During engine operation, the centrifugal compres-
sor compresses and increases the pressure and temper-
ature of the incoming air from its surroundings. The
static pressure of the air rises in the diffuser whereas the
velocity is reduced as it passes through the diverging
passages (vanes). Fuel is added to the low velocity air
in the combustion chamber to burn continuously to pro-
duce high temperature, pressure, and velocity gases.
The turbine expands the high temperature gas from the
combustion process to produce mechanical shaft power
to drive the compressor. The convergent exhaust pro-
pelling nozzle accelerates the exhaust gases from the
turbine to create thrust for propulsion.
A micro gas turbine operates on the principles of a
Brayton open-air cycle. The cycle is represented in a
temperature-entropy (T-s) and pressure-volume (p-v)
diagram as shown in Figs. 2 and 3. In a Brayton, open-
air cycle the engine working fluid exits/exhausts into
the atmosphere after expansion in the turbine and the
exhaust propelling nozzle.
The cycle is assumed to have isentropic air
compression, constant pressure heat addition,
isentropic gas expansion, and constant pressure heat
rejection. In practice, these processes are not isentropic.
As shown in Figs. 2 and 3, isentropic compression
occurs at point 0-3 in the compressor, constant pressure
heat addition at 3-4 in the combustion chamber,
isentropic expansion at position 4-5 in the turbine,
isentropic work done at station 5-8 in the nozzle and
finally constant pressure heat rejection from 8-0 into
the atmosphere
The engine performance or thermodynamic parameters
are evaluated using the turbomachinery transport
equations, thus mass, momentum, energy, and speed
compatibility equations. The cycle analysis starts at the
cold section (compressor inlet to combustor inlet) of the
engine to the hot section (combustor inlet to nozzle
outlet) of the engine. From Figs. 2 and 3, the
compressor pressure ratio and efficiency are
determined using these relations:
p
p
c
00
03
(1)
TT
Tc
c0003
1
00 1
(2)
where,
is the compressor total-total pressure ratio
is the compressor inlet stagnation pressure
is the compressor downstream stagnation pressure
is the compressor upstream stagnation temperature
is the compressor outlet stagnation temperature
is the compressor total-to-total efficiency
Knowing the compressor parameters, the combustion
outlet stagnation temperature is estimated as:
c
TT pg
fLHV
cc
0304
(3)
where,
is the combustor outlet stagnation temperature
is the combustion efficiency
is the fuel to air ratio
is the heating value of fuel
is the specific heat of gas.
Applying the principles of work and speed compatibil-
ity, the turbine downstream stagnation temperature is
determined using equation (4):
m
pg
g
pa
ac
m
TT
c
m
TT .
.0003
0405
(4)
where,
is the combustor outlet stagnation temperature
is the shaft mechanical efficiency
is the specific heat of air
is the mass flow rate of gas
With known turbine parameters, the engine exhaust ve-
locity is determined.
TT
cu pge805
2
(5)
where,
is the exhaust velocity
is the nozzle outlet static temperature
Having estimated the nozzle exhaust velocity the en-
gine thrust is found:
uu
m
Fe
a
net 0
.
(6)
where,
is the freestream velocity
is the net thrust
The net power is calculated
WW
PCT
net
(7)
where,
is the turbine work
is the compressor work
is the net power
The compressor and turbine maps shown in Figs 4 and
5 are used to model the engine performance at the de-
sign and off-design points. The maps consist of effi-
ciency and pressure ratio graphed against mass flow
rate. Off-design evaluation is considered as the reliable
and effective means for modelling, investigating, and
predicting gas turbine operating range [21]. Accord-
ingly, the impact of a change in the engine’s nonlinear
thermodynamic behaviour characteristics on the engine
performance is examined at the off design state.
Ahmed [22] emphasise that the engine performance at
off-design conditions should be satisfactory and safe ir-
respective of its non-linear behaviour. Although gas
turbines are designed to operate effectively at specific
design points, the engine should also work acceptably
at off-design conditions. The off-design performance of
a gas turbines involves performance predictions, condi-
tion monitoring and degradation analysis [23]. The en-
gine diagnosis and degradation analysis are normally
executed based on the performance predictions [24].
The engine performance can be evaluated at off design
using the following methods: component matching,
stage stacking, gas path analysis, computational fluid
dynamics [25]. These modelling methods are mathe-
matically formulated with linear and non-linear equa-
tions into computer programs for performance simula-
tions.
Although, MGTs have outstanding advantages
over its competitors in the aviation and power environ-
ment, however, miniaturization of conventional larger
gas turbines into MGTs leads to engine efficiency and
performance loss due to low Reynolds number, poor
heat transfer, large (compressor and turbine) tip losses,
and other mechanical and geometrical limitations [10],
[26]. Micro turbomachinery inherently have low Reyn-
olds numbers due to the small size of the engine, there-
fore resulting in relative viscous forces dominating the
inertia forces in the engine [27]. This can cause consid-
erable loss of pressure and temperature and other ther-
modynamic variation in the engine, hence reducing the
efficiency and power output of the engine [1], [14]. The
engine’s low Reynolds number can lead to flow sepa-
ration and transition in the compressor [28], and conse-
quently affects the compressor efficiency and pressure
ratio [1], [29]. In addition, the relatively high viscous
effects can slow down the mixing of fuel and air, and
hot and cold gases in the combustor, therefore, reducing
combustion residence time [30]. Xiang et al. [27] as-
sessed the effects of Reynolds number on an MGT’s
miniaturised compressor. It was found that the com-
pressor efficiency and pressure ratio decreases with de-
creasing Reynolds number. This is due to an increased
entropy change, hence increasing the viscous losses.
The engine internal heat transfer influences the en-
gine power output and thermal efficiency. Whereas an
MGT’s high power to weight ratio presents significant
advantages, the higher surface to volume ratio creates
heat transfer complexities in the engine [26], [31], re-
sulting in thermal losses in the engine combustion
chamber, hence, limiting the combustion efficiency and
flammability as well as the overall operating efficien-
cies of the engine [14]. A detailed analysis of internal
heat transfer of small aero engine using 3D Navier-
Stoke solver has been provided by Verstraete et al.
[32]. The heat flux from the turbine to the compressor
and the components’ adiabatic and diabatic perfor-
mance were used as performance inputs. They exam-
ined the impact of recuperative and non-recuperative
cycles on power output and efficiency of the MGT. Ac-
cording to the authors, the results show that decreasing
cycle efficiency is almost independent of compressor
efficiency but increases for a regenerative cycle.
Verstraete et al. [33] and Verstraete et al. [34] numeri-
cally studied the heat transfer inside an MGT and its
effects on the engine performance. A conjugate solver
was used to quantify and investigate the heat transfer
and the different mechanisms that contribute to it. The
study showed that a large temperature difference be-
tween the turbine and compressor, in combination with
the small dimensions, results in a high heat transfer
causing a drop in efficiency of both components. Sun
et al. [35] implemented a computational fluid dynamics
(CFD) based comparison study of adiabatic and non-
adiabatic flow characteristics to evaluate the flow phe-
nomenon and to quantify the physical mechanisms
leading to heat transfer in MGT compressors. The re-
sults revealed that the compressor efficiency, mass flow
rate and pressure ratio decreased by 7.6%, 1.5%, 4.4%
and 20.8%, 3.7% and 14.5% for isothermal wall temper-
atures of 450K and 600K, showing that heat addition is
a major contributing factor of MGT compressor perfor-
mance deterioration. The heat transfer led to an increase
in relative Mach number at the impeller hub region,
which resulted in flow separation. The analysis showed
that heat addition increases the angle of tip clearance
flow and therefore reduces the slip factor. They pro-
posed that tip clearance flow blockage and relative
Mach number should be accounted for in the correction
of the slip factor in compressor impeller design.
In general, effective and efficient turbine cooling
significantly increases the allowable turbine inlet tem-
perature (TIT) of the engine and consequently im-
proves the engine performance output. Conventional
larger gas turbines cool the turbine blades using bleed
air from the compressor [36]. Smaller gas turbines can-
not accommodate this cooling technique and are forced
to run at lower turbine inlet temperatures. The turbine
inlet temperature depends on the turbine blade alloy
stress rupture and low cycle fatigue strengths, and cool-
ing options [37]. Research thus far shows that MGTs
turbine inlet temperatures are restricted between 600-
800°C [12], [38]. In order to accomplish high temper-
atures at the turbine inlet, high temperature alloys or
ceramics are required for the turbine design.
Higher tip clearance losses are intrinsic in MGTs
due to relatively short compressor and turbine blade
heights [29]. The higher the tip clearance the lower the
component efficiencies [39] and the higher perfor-
mance deterioration and fuel consumption of gas tur-
bines [11]. An MGT’s compressor has low-pressure ra-
tios due to smaller blade heights. The compressor
pressure ratio influences the specific fuel consumption
of the engine.
The manufacturing of a smaller gas turbine is similar
to that of a larger gas turbine. However, in spite of the
careful attention to detail at the design stage and during
manufacture, small turbomachines always have lower
efficiencies than larger, geometrically similar machines
[40]. The manufacturing accuracy attained for larger
gas turbines is impossible for MGTs; hence, the engine
would not perform as expected. They have high relative
surface roughness from the design process and there-
fore high engine losses [26]. Additionally, excessive
frictional losses in the engine bearings adversely affect
the engine cycle efficiency.
Different research projects have been undertaken on
micro gas turbines with the purpose of improving their
performance. This includes MGT compressor and tur-
bine stage design improvements (new designs and flow
physics analyses), new combustor designs, and com-
bustion analysis. The intent of these projects was to op-
timise the engine performance parameters such as the
specific fuel consumption, pressure ratio, cycle peak
temperature, system pressure losses, turbine and com-
pressor efficiencies and the engine thrust and power.
Some of the techniques considered for the performance
improvements are discussed below.
Fig. 6 shows the centrifugal compressor stage of a mi-
cro gas turbine engine.
The compressor stage consists of an impeller equipped
with radial (wedge) diffuser (shown in Fig 6) to reduce
the velocity of the airflow and increase the static pres-
sure. These mini centrifugal compressors intrinsically
have low isentropic efficiency (50-70%) and pressure
ratio (2-5%) due to low Reynolds numbers, blade tip
leakage, growth of boundary layer and separation on
the blades [1], [42], [43]. In recent years, research has
been directed towards increasing the performance of
MGT compressors through redesign and through com-
putational modelling and simulations. Amirante et al.
[44] studied and optimised the Pegasus small turbojet
engine compressor intake. They sought to analyse and
improve the effects of boundary layer growth and ve-
locity profile using ANSYS FLUENT incorporated
with the progressive optimisation technique. The nu-
merical result disclosed that the inducer inlet flow field
is dominated by flow separation covering almost 8% of
the inlet section. Therefore, for proper design of the in-
let the airflow rate should be increased to improve the
velocity profile and hence reduce flow losses. The op-
timisation, which was aimed at mitigating the intake
pressure losses and recirculation bubble, increased the
intake mass flow rate by 2% to improve the velocity
profile. They concluded that the numerical results and
experimental data match well.
In 2007, Ling et al. [45] designed and improved the
compressor stage of the KJ66 micro jet engine using
ANSYS CFX. The intent was to improve the efficiency
and performance output of the engine. They increased
the pressure ratio of the centrifugal compressor at a
lower mass flow rate. They contend that the new design
outperformed the baseline compressor. Aghaei et al.
[46] deals with the design and computational fluid dy-
namics analysis of a micro gas turbine compressor.
Their design involved theoretical and computational
fluid dynamics (CFD) modelling of the compressor
stage. A pressure ratio of 4 was obtained for the new
compressor stage design.
Jie & Guoping [47] discussed the re-design of an 11
cm diameter MGT compressor diffuser. They sought to
investigate the effect of cross-sectional area distribu-
tion along the flow path of the new diffuser's perfor-
mance. The new diffuser configuration was equipped
with main and splitter blades. The computational fluid
dynamics (CFD) and the experimental predictions
showed that the pressure coefficient and pressure re-
covery coefficient improved by 0.65 and 0.9. It was
found that the thrust of the engine increase by 11% and
the specific fuel consumption decreased by 9%. In the
study by Tough et al. [48], the performance of an MGT
compressor impeller was evaluated using different inlet
blade angles. The writers studied the effects of different
blade inlet angle and backward sweep on compressor
performance. Tough et al. [48] found that reducing the
backsweep and the inlet blade angle by 3° and 2° pro-
duced a stable operating range, high-pressure ratio, and
total efficiency.
Van der Merwe [49] designed and optimised the
centrifugal compressor impeller of the BMT 120KS en-
gine. His aim was to achieve a total-to-total pressure
ratio of 4.72 and an isentropic compressor efficiency of
79.8% at a mass flow rate of 0.325 kg/s. The new im-
peller performance was validated by comparing its
mean-line, experimental and CFD results. He showed
that the experimental and numerical data correlated
well. Krige [6] redesigned the BMT engine vaned dif-
fuser. He aimed to maximise the compressor stage pres-
sure recovery in order to increase the engine's total -to-
static pressure ratio, mass flow and thrust output. The
experimental and numerical examination showed that
the diffuser pressure recovery increased from 0.48 to
0.73. The static-to-static pressure ratio across the dif-
fuser increased from 1.39 to 1.44. De Villiers [50] em-
ployed a 1D (1-dimensional) mean-line code and CFD
software codes FINETM/Turbo and FINETM/Design3D
to design a centrifugal compressor stage for the BMT
engine. According to the author, the new compressor
stage yielded a total-to-static pressure ratio of 3.0, and
efficiency of 76.5% and a thrust output of 170 N at a
rotational speed of 119 000 rpm.
Alex & M [51] report on the design of an MGT
compressor stage implementing ANSYS FLUENT.
They used titanium, aluminium, and stainless steel al-
loy materials to design different compressor stages. Vi-
brational and stress analysis were investigated for the
various designs. The analysis showed that titanium al-
loy offers the best design in terms of safety factor. They
stipulated that the centrifugal stress distribution is con-
centrated at the blade hub. Recently, Burger [41] de-
signed and optimised a crossover diffuser for the BMT
engine. The crossover diffuser combined with Van der
Merwe's [49] impeller improved the compressor stage
total-to-static pressure ratio from 2.62 to 3.65. An inter-
polated thrust output of 200 N was predicted by the new
diffuser at the engine maximum speed of 120 000 rpm.
The performance of the KJ-66 MGT compressor
has been investigated by Xiang et al. [29] implementing
a steady and unsteady Reynolds Average Navier-Stoke
CFD solver. They intended to study the compressor
transient flow behaviour during operation. The simula-
tion results under predicted the total-to-total adiabatic
efficiency showing a decrease of the peak efficiency
from 0.73 to 0.55 while the total peak pressure ratio in-
creased from 1.54 to 1.96. However, they proposed that
the effects of friction and tip losses at the compressor
shroud should be further modelled for better perfor-
mance output.
Diener [52] presented the design of a mixed flow
impeller for the compressor stage of a 600N small tur-
bojet engine. A 1-dimensional and multi-point optimi-
zation was performed to assess the aerodynamic perfor-
mance of the stage. He concluded that a total-to-total
pressure ratio and isentropic efficiency increase of
30.6% and 5% were obtained for the compressor stage.
Kock [53] designed and optimised a crossover diffuser
for Diener [52] impeller using mean line and numeri-
cal evaluations. According to Kock [53] the optimised
diffuser showed static-to-static pressure recovery in-
crease of 21.5% whereas the total-to-total isentropic ef-
ficiency and pressure ratio yielded increase of 8.75%
and 7.7%.
The main function of a combustion chamber is to in-
crease the maximum allowable temperature and energy
of the gas turbine working fluid. The fuel mixes with
air in the chamber and undergoes exothermic chemical
reaction to release thermal energy [54]. Fig 5 depicts a
micro jet engine combustor. It consists of inner and
outer liners with dilution holes, vaporizing tubes and
fuel injection needles. Combustor design is the most
complex among MGT components; hence, the design
of an effective and efficient combustor is indispensable
for the performance of the engine. Previous combustor
design procedures were based on experimental results
and empirically derived design rule. CFD has been em-
ployed in recent combustor design and analysis pro-
cesses for reliable modelling of the internal flow behav-
iour.
Micro gas turbines to date use a constant pressure com-
bustor. They are simple to implement and provide a rel-
atively steady, uniform gas flow to the turbine. Alt-
hough they have practical benefits, the flow behaviour
in the combustor is complex due to compactness and a
relatively high viscous environment. The combustion
efficiency, flame stability and ignition, wall cooling,
pressure loss and emission control are some of the com-
mon challenges faced in the design and performance of
this combustors [55]. These parameters ascertain the
engine operational range, durability, cost, maintainabil-
ity and emissions characteristics [22].
Ridzan.J.J.M et al. [56] have considered the aero-
dynamic flow analysis in a micro combustion chamber.
Their focus was to study the effects of swirl on the flow
inside the combustor. They employed four axial vane
swirlers in the order of 20°, 30°, 45°, and 60° with swirl
numbers of 0.27, 0.42, 0.74, and 1.285 respectively. The
investigation showed that high swirl vane angles in-
creases swirling in the combustion chambers, therefore
increasing turbulence strength, recirculation zone size
and the amount of recirculation mass at increasing pres-
sure loss in the combustor. Guidez et al. [57] described
the studies performed on combustion characteristics of
a miniature combustor. The aim was to examine com-
bustion stability and efficiency. Raman spectroscopy
and 1D Rayleigh scattering and standard thermody-
namic measurements were used to estimate the temper-
ature profile and main species concentration at the com-
bustor outlet. They measured a combustion efficiency
of 80%.
Chaudhari et al. [58] report the design and simula-
tion of a miniature annular combustion chamber using
ANSYS CFX. They evaluated the impact of flow pat-
terns and temperature distributions on the combustion
chamber liner walls. The numerical results showed that
combustion flames, which can cause combustor failure,
significantly affect the combustion chamber walls. Fur-
ther, Gieras & Stankowski [4] have performed compu-
tational modelling of the aerodynamic flow in the
GTM-120 micro jet engine combustor. They considered
the effect of engine downsizing on the mass flow, pres-
sure losses, and heat transfer in the combustor. The aim
of their research was to maximise thermal efficiency by
minimising fuel consumption and controlling emis-
sions. The simulations yielded a combustion pressure
loss of 10%. They emphasised that the size and shape
of diffuser channels (vanes) give rise to high speed and
non-uniformity of combustor aerodynamic flow, which
are an important source of pressure loss in the combus-
tor. The authors indicated that the consequences of ex-
cessive flow velocity in the diffuser and the annulus
might restrict the flow through the first row of holes in
the combustion chamber, therefore, deteriorating the
process of mixing fuel and air.
Krieger et al. [59] numerically and experimentally
evaluated a micro jet engine combustion chamber using
ANSYS FLUENT RANS turbulent solver. They ac-
cessed the flow physics of heat transfer and temperature
profiles in the combustion chamber liner walls. The
flame was stabilized in the burner using a swirler and
reverse flow configuration. The reverse flow configu-
ration enhanced cooling of the liner walls, as well as
preheating of incoming air and ensured flame stability.
The flow at the inlet of the combustor is influenced by
swirl due to loads on the diffuser and the vanes at the
end of the compressor. It was found that the swirler in-
creases air-fuel mixing processes. According to Krieger
et al. [59], the numerical and experimental results
showed good correlation. Armstrong & Verstraete [60]
assert that a constant volume combustor increases the
pressure in the chamber and reduces engine fuel con-
sumption as well as increasing the engine cycle thermal
efficiency. They describe the re-design of a constant
volume combustor for a micro jet engine employing a
choked nozzle guide vane, which functions as a flow
restrictor, incorporating a mechanical valve at the inlet
of the chamber. The new design improved the specific
fuel consumption and thrust by 27% and 35% for a
combustion pressure ratio of 1.1 and 1.25. It was found
that the inlet valve discharge coefficient influences the
average combustion pressure ratio. Discharge coeffi-
cients below 3.0 are required to increase the pressure
ratio.
Gao et al. [61] examined the performance of the
SR-30 micro turbine combustor. They analysed com-
bustion flow characteristics such as liner wall tempera-
ture, pressure and temperature distribution in the cham-
ber. The total air-to-fuel ratio in the chamber decreases
with increasing engine rotational speed, whereas the
average liner wall temperature increases due to the in-
sufficient cooling air after complete combustion. The
investigations showed that orifice distribution and com-
bustor geometry dominates the flow pattern inside
combustion chamber. According to Gao et al. [61], hot
areas and/or spots in the liner walls are dominant in the
primary and dilution zones. Suchocki et al. [62] ana-
lysed the combustion performance characteristics of the
GTM-140 mini turbojet engine combustor employing
computational investigations. They sought to under-
stand and identify the chemical process in the combus-
tor. The combustion was modelled using the k-ϵ
(RANS) Turbulence Model and Non-Premixed Model.
The performance investigations showed a pressure drop
of 9%. However, they proposed that modification of the
diffuser geometry should be considered for further de-
celeration of the velocity.
Fuchs et al. [9] have considered the flow character-
istics of a small aero engine combustor using experi-
mental and numerical evaluations. The computational
investigation was performed employing ANSYS CFX.
Their objective was to study the behaviour of fuel dis-
tribution and atomization in the engine combustor. The
engine was tested at 0°, 15°, and 30° pre-swirl angles.
Low air-fuel ratio's (AFR) resulted in high combustion
temperatures whereas high AFR showed low tempera-
tures. Basson [63] deal with the investigation of swirl
effects on the mass flow distribution and flow struc-
tures in a miniature combustion chamber. She assumed
constant pressure drop along the liner walls in the com-
putational modelling, however, the numerical and the
experimental results showed that the pressure varies
along the inner liner wall zones. The author concluded
that the numerical results correlated well with the ex-
perimental data.
The turbine extracts kinetic energy from the combus-
tion hot gases to produce mechanical shaft power to
drive the compressor. The turbine efficiency depends
on how well it extracts kinetic energy from the combus-
tion hot gases. Fig 8 shows the axial turbine stage of a
small jet engine.
The turbine stage consists of a rotor (left) and stator
(right). The incoming combustion hot gas is initially ac-
celerated through the stator vanes before entering the
rotor. Although axial turbines are the standard for small
jet engines, both axial and radial turbines can be used
for MGTs. The major challenges in turbine design are
effective and efficient cooling techniques, tip clearance
control, and blade materials suitable for high-tempera-
ture and corrosion [54]. The blades are subjected to
high temperatures, which can cause creep and corrosion
failure of the blade without effective cooling [64].
Wojciech & Michal [65] describe design-point and
off-design mean-line performance investigations and
numerical analysis of axial flow turbine stage of RC jet
engine employing CFD software NREC AXCENT.
They developed aerodynamic 2D and 3D entropy per-
formance contours for the turbine stage. They contend
that turbine tip clearance, Reynolds number, and spe-
cific speed coefficient influence turbine efficiency and
overall engine performance. Ensuring high Reynolds
number and specific speed coefficient as well as low tip
clearance increase the turbine work transfer. Verstraete
et al. [1] investigated and improved the overall perfor-
mance of the KJ-66 turbojet turbine stage. They
adapted a spherical dimple vortex turbine blade profile
for the axial turbine stage in order to obtain the required
performance improvement. The analysis disclosed that
flow separation and transition at low Reynolds number
is the cause of low turbine efficiency. They concluded
that the modified engine showed improvement in effi-
ciency compared to the baseline engine.
Basson [66] re-designed and examined new turbine
stages for the BMT 120 KS micro jet engine. He per-
formed empirical analytical and numerical analyses for
the new designs. Reverse engineering was used to pro-
duce 3D models of the existing axial flow turbine of the
engine. A finite element analysis was conducted to de-
termine the structural behaviour and stress distribution
of the axial flow turbine under different static and dy-
namic loadings. He concluded that the proposed de-
signs predicted an efficiency difference of 5%. Smit
[67] manufactured Basson [66] turbines, and obtained
experimental results for the newly designed turbines.
The experimental results showed disparities in the to-
tal-to-total temperature ratios due to non-uniform dis-
tribution of flow conditions at the immediate turbine
stage outlet; however, he suggested that extra measur-
ing instrumentation should be integrated into the meas-
uring of the turbine inlet and outlet temperatures.
Lakshmy [8] numerically designed and analysed a
MGT turbine stage. Different flow velocities and pres-
sures at different angle of attacks were used for the in-
vestigation. She claims that the engine thrust was in-
creased for the new design. The design and
performance analysis of a small axial flow turbine has
been conducted by Ennil et al. [68] using ANSYS
CFX. The turbine blade was optimised with the multi-
objective genetic algorithm with the aim to reduce the
turbine stage losses and increase the efficiency. He con-
cluded that the investigations yielded a total-to-total ef-
ficiency of 87.78% for the turbine stage.
Yedla et al. [69] have considered the design of an
axial turbine blade for a micro gas turbine engine em-
ploying numerical simulations. The writers stipulate
that the blade is susceptible to turbulence and shock
losses. Moodley [70] have described the design and
aerodynamic analysis of a single stage axial flow tur-
bine of an MGT. He performed a numerical analysis
using Concepts NREC Axial and AxCent design
softwares and compared the results to a meanline
analyis of the turbine blade geometry. A stress analysis
was conducted on the blades at design point operating
conditions with aerodynamic loading and free rotation.
A wave rotor engine uses pressure waves to transfer en-
ergy between the engine’s working fluids [71], there-
fore, increasing the pressure and temperature of the low
energy fluid of the engine. Fig 9 depicts the thermody-
namic cycle, thus the T-s diagram of a wave rotor en-
gine and engine without wave disk.
As depicted in Fig 9, the wave rotor technique al-
lows pressure gain in the combustion chamber, there-
fore increasing the engine’s thrust and power output as
well as the cycle thermal efficiency [72]. In a typical
wave rotor engine, the hot gases from the combustion
compresses the air coming out of the compressor [73].
The burnt gas is pre-expanded in the wave disk,
hence delivering hot gas of lower temperature to enter
the turbine inlet [71]. Nevertheless, the wave disk al-
lows higher temperatures in the combustion chamber,
thence, increasing the combustor downstream pressure
[75]. The wave rotor is embedded between the com-
pressor and the turbine and parallel to the combustion
chamber as shown Fig 10. The wave disk arrangements
are of two configurations, namely radial and axial.
Wilson & Paxson [76] investigated the performance of
a simple turbojet engine with the wave-rotor enhanced
technique. They obtained an increase of 1 to 2% for the
engine thermal efficiency and 10 to 16% for the specific
power. Snyder & Fish [77] discuss the performance
benefits of the wave rotor cycle in a small turboshaft
engine. The engine topped with the wave rotor showed
a considerable reduction of 22% in specific fuel con-
sumption (SFC) compared to the baseline engine.
Akbari & Müller [78] suggest that the wave-rotor
technique is the efficient approach to improve small jet
engine thermodynamic cycle performance. The authors
investigated and evaluated the performance of a small
turbojet engine at different thermodynamic conditions
with a four-port wave rotor. According to the authors,
the outcome of their investigation showed significant
improvement in the combustion pressures and temper-
atures compared to the baseline engine.
Iancu et al. [72] described the uses and advantages
of a wave rotor in MGTs. According to them, a wave-
rotor cycle increases the overall gas turbine compres-
sion and expansion ratio. They argued that the wave ro-
tor technique could achieve a 50 to 83% increase in
compression efficiency for ultra-micro and small gas
turbines.
Piechna & Dyntar [75] presented the numerical in-
vestigation of a two-dimensional model wave disk mi-
cro-engine taking into account unsteady, centrifugal
and Coriolis forces existing in such a configuration.
The main goal was to determine the engine gas com-
pression and torque generation. They concur that the
main challenge faced by an axial wave disk configura-
tion is scavenging in the channels; the use of a radial
wave rotor prevents the scavenging phenomenon.
Due to the complex and difficult geometry and flow
physics of the processes occurring in small turbojet en-
gines, various methods of system evaluation have been
used to investigate micro gas turbine engine perfor-
mance output. This includes computational (CFD), the-
oretical (simulation and analytical), and experimental
analysis of the processes occurring in the engine. The
CFD tool is an effective means of studying the mutual
behaviour of the flow physics of the engine elements.
Fig 9 illustrates the results of a computational flow
analysis of the flow phenomenon in a gas turbine unit.
Yin et al. [23] evaluated the performance of an
APU small gas turbine engine using experimental in-
vestigations and TURBOMATCH software simula-
tions. The engine exhaust temperature profile showed
high fluctuations in comparison with the exhaust pres-
sure. This is due to the presence of hot areas and/or
spots at the engine exhaust. They identified that the cy-
cle thermal efficiency drops as the ambient temperature
rises. Yin et al. (2003) stipulated that the simulation
showed good agreement with the experimental test.
Rahman & Whidborne [80] developed a model to
investigate the effects of bleed air on the performance
of the AMT Olympus single spool engine. The inter
component real time simulation was used for both open
and closed loop simulation of the engine. It was re-
ported that the simulated results and the experimental
data showed good correlation.
Trebunskikh et al. [79] assessed the KJ66 micro
turbine engine performance output employing FloEFD,
CFD software code. They analysed the internal flow
structure in engine from the inlet to the outlet of the en-
gine. The mass, momentum and energy equations were
used to examine the nature of the flow in the compres-
sor, combustion chamber, the turbine and the conver-
gent nozzle. The authors concluded that the numerical
results correlate well with the experimental test results.
However, they reported that non-uniform fluid temper-
ature distributions at the combustor outlet, and wedge
diffuser inefficiencies contribute towards poor perfor-
mance of the engine.
Badami et al. [81] used ANSYS FLUENT to ana-
lyse the performance of the SR-30 micro jet engine.
The intent was to investigate the engine thermody-
namic cycle performance. According to them, the CFD
data showed consistent trend with the experimental val-
ues. The nozzle and turbine pressure ratios showed dif-
ferences of less than 5% and the mass flow rate and
thrust values showed percentage differences of 2% and
10% respectively.
Leylek [82] studied and investigated the perfor-
mance of the Olympus HP micro turbojet engine. He
employed a mean-line, through flow and CFD analysis
for the studies. The CFD analysis was performed using
Numeca CompAero and TurbAero. The main purpose
of the study was to understand thermodynamic model-
ling and performance characteristics of the engine. In-
frared modelling and thrust augmentation analysis were
also investigated. Leylek [82] blocked the engine con-
vergent nozzle and also reduced the nozzle area to
stress the engine to operate at different operating con-
ditions in order to assess the thrust output and compres-
sor and turbine operating maps. The author concluded
that the numerical results correlate well with the exper-
imental data.
Krivcov et al. [83] studied the performance of a small
gas turbine engine using ANSYS CFX and thermody-
namic model ASTRA software. They employed a gas-
dynamic modelling analysis. The investigation did not
account for the presence of leaks and heat transfer be-
tween the stream and the air-gas wall channels. The
study showed percentage difference of less than 7% be-
tween both simulations. They claim that good agree-
ment was obtained between the ANSYS CFX results
and the ASTRA thermodynamic 1D results for the en-
gine. Roberto Capata [5]
Oppong [25] investigated the performance of the
BMT 120 KS jet engine using computational, theoreti-
cal, and experimental evaluations. Flownex SE and
GasTurb simulation software were used for the numer-
ical analysis. The aim of the study was to examine
matching of the engine components. He performed a
thrust augmentation test by increasing and reducing the
engine turbine and nozzle exit areas. A Sensitivity anal-
ysis was conducted using different modified (rede-
signed) compressor and turbine stages of the engine.
The simulations revealed that the turbine capacity was
insufficient to drive the compressor of the engine, gen-
erating high exhaust gas temperatures (EGT) in an at-
tempt to increase the fuel consumption to increase the
turbine power. The experimental analyses indicated the
engine turbine inlet and exhaust showed high tempera-
ture fluctuations due to the presence of hot spots in the
turbine inlet and exhaust.
Elzahaby et al. [84] using theoretical simulations
and experimental investigations as well as the GasTurb
software simulations evaluated the thermodynamic pa-
rameters of a micro turbojet engine. The engine was
modelled using momentum and energy equations at the
various stations of the engine to estimate the flow pa-
rameters. The engine component maps combined with
the gas-dynamic equations were used to assess the en-
gine’s performance. The authors confirmed that the the-
oretical calculations and the experimental test yielded
good results when compared to the GasTurb simula-
tions.
The micro gas turbine engine is considered as one of
the key technologies in the hobby industry as well as
for small combined heat and power generation applica-
tion. This paper focuses on a review of treatises of
MGT performance studies. The following conclusions
are drawn:
Gas turbines performance is modelled and pre-
dicted at the off-design conditions or state using com-
ponent maps. Off-design offers better performance di-
agnostic and prediction of the non-linear behaviour of
the engine. Component matching performance model-
ling forms the basis of the other performance modelling
approaches. It is suitable for the non-linear simulation
of gas turbines due to its flexibility, consistency, sim-
plicity, and better accuracy [36]. The working principle
depends on compatibility of engine mass flow, work,
and rotational speed taking into account mechanical
losses[25].
Although, micro-fabrication has paved way for the
manufacturing of micro gas turbines, tip clearance
losses and geometrical limitations obstructs the smooth
and optimum performance of the engine. The same
and/or precise geometric resemblances cannot be estab-
lished for the scaled MGT engine and the larger gas tur-
bine due to increase relative surface roughness and
leakage losses because of increase relative tip clearance
of the downsized components.
The entropy change of small gas turbines increases
due to downsizing and/or scaling of larger gas turbines
into smaller engines. Therefore, the Reynolds number
of the engine decreases and increases the relative effect
of viscosity of the flow in the engine and consequently
limits the engine overall thermal efficiency and the per-
formance output of the engine. Research shows that
Reynolds number, heat transfer and thermal losses ef-
fects dominate in the gas generator units and signifi-
cantly affects the engine performance. Low Reynolds
number can lead to flow transition and separation in the
compressor, resulting in low compressor pressure ratio
and efficiency. Low Reynolds number also influences
the combustion residence time. However, published
studies on Reynolds number influence on MGTs per-
formance are few; therefore, this subject area of MGTs
performance evaluation still needs to be further inves-
tigated.
The rate of heat transfer is a significant trait of fluid
mechanics in micro gas turbine engines as these devices
operate in a completely different design space than
large-scale gas turbines. The heat flux from the hot tur-
bine to the cold compressor results in heat addition in
the cold compressor via conduction. This leads to large
temperature difference between the turbine and com-
pressor, subsequently resulting in reduction of the sys-
tem mass flow rate, efficiency, and pressure rise. Addi-
tionally, the heat transfer contributes to increase in the
compressor tip flow angle and hence reduces the slip
factor [35]. It has been proposed that the use of recu-
perator in micro gas turbines could alleviate some of
the shortcomings associated with heat transfer in MGTs
[10], [26], [32]. This will ensure low temperatures in the
system and therefore, reduce the heating effects in
MGTs.
The wedge type of diffusers contribute to a greater
loss of efficiency in the compressor unit [6], [79]. Also,
the curved or crossover diffuser is prone to flow sepa-
ration in the curved passages hence the volumetric flow
rate of the fluid have to be increased in order to de-
crease this effect [45]. Redesign or reconfiguration of
the compressor stage of the engine can substantially im-
prove the compressor efficiency. Mixed flow compres-
sor stage designs and other design configurations have
been adopted for the stage to improve the overall per-
formance of the engine. Efficiencies of between 75-
85% [1], [6], [41], [45], [50] and beyond can be
achieved by redesigning the compressor stage. Addi-
tionally, redesigning and optimization of the compres-
sor stage improves the stage pressure ratio and the fuel
consumption of the engine. While, redesign of the com-
pressor stage improves the compressor efficiency and
pressure ratio, it is appropriate to consider the turbine
stage parameters in the design process for smooth run-
ning of the engine. Thus, increasing the compressor
stage parameters without considering the turbine stage
could result in compressor-turbine mismatching. Con-
sequently, the turbine supplies insufficient work to the
compressor and could lead to excessive fuel consump-
tion and high exhaust temperatures.
Despite, a good understanding of the turbine stage
characteristics, effective and efficient material for the
design of the stage is still a challenge for micro gas tur-
bine design engineers. The material used to manufac-
ture MGT turbine blades is Inconel IN713LC, a nickel-
based super-alloy. The yield strength of the material de-
teriorates under high temperatures. Therefore, the tur-
bine blade inlet temperature is limited between maxi-
mum allowable temperatures of 1000 K and 1100 K to
give better operating performance and minimum risk to
the turbine blades [25], [66]. The combination of high
temperature alloys and other cooling techniques that
are desirable to withstand high temperatures and stress
could improve the turbine inlet temperature and the en-
gine output. In addition, redesigning the turbine stage
with high load and flow coefficients enhances the tur-
bine power and the overall power and thrust of the en-
gine [85].
Most of the performance analyses on MGTs have
been considered on the aerodynamic flow behaviour
and physics in the combustion chamber due to the com-
plexities related to the flow dynamics. Numerical anal-
yses and computational modelling is suitable for deter-
mine the combustion chamber material, liner wall size
and improve the efficiency of micro combustion cham-
ber with reduction in heat losses. Constant pressure
combustors are mostly used in MGT applications [60].
These burners are susceptible to pressure loss. There-
fore, pressure gain techniques such as the use of con-
stant volume combustors and the implementation of
wave rotor methods are considered. The constant vol-
ume combustion chamber reduces the pressure drop in
the combustor and increases the combustor down-
stream pressure, hence contributing to overall increase
of the cycle efficiency and performance of the engine.
Likewise, the use of wave rotor method or approach in-
creases the combustion temperature and reduces the
combustion pressure loss. This configuration delivers
the maximum allowable temperature permissible to the
turbine inlet without infringing on the integrity of the
turbine material [86].
The combustion chamber walls and the combustor
outlet or turbine inlet are considerably vulnerable to
high temperature profiles and thermal losses. The high
temperature propagates into the turbine and adversely
influences the temperature measurements in the turbine
hot sections. The temperature measurement at the tur-
bine inlet and exhaust to assess the engine performance
remains a challenge due to the high temperature fluctu-
ations. Therefore, in order to measure reliable turbine
inlet and exhaust temperatures perfectly shielded ther-
mocouples are required. In addition, careful attention
must be paid to some unacceptable temperature gradi-
ents through the outlet section of the combustor as it
appears the most challenging problem to be dealt with
in terms of combustor fluid dynamics [87].
The fuel flow and nozzle area are the control inputs
usually used to change the state or operating regime of
the engine. Reducing the nozzle area increases the
thrust of the engine, whiles the fuel flow rate of the en-
gine also increases to compensate for the inefficiencies
of the engine-operating regime and/or the reduction in
turbine work. Increasing the fuel flow results in higher
exhaust gas temperatures. Therefore, the engine runs at
a higher turbine inlet for reduced nozzle area.
In conclusion, bleeding the compressor stage of the
engine to cool the turbine blades lowers the mass flow
through the engine, therefore the turbine will have to
operate at a lower mass flow to produce the power re-
quire to drive the compressor [80]. This results in
higher turbine inlet and exhaust temperatures due to an
increase in fuel consumption.
[1] D. Verstraete, P. Hendrick, V. Djanali, C. Gonzalez, J.
PowerMEMS 2010 Conference, Australia, 2010, pp.
14.
[2] G. B. Bruening, J. R. Snyder, and R. E. Frede
potential impact of utilizing advanced engine
technology for a combat capable unmanned air
J. Eng. Gas Turbines Power, vol. 123,
no. 3, pp. 508512, 2001.
[3]
hybrid propulsion system for a high-
configuration selection , aerodynamic study , and gas
35, 2014.
[4]
of an aerodynamic flow through a micro-turbine
engine J. Power Technol., vol. 92, no. 2,
pp. 6879, 2012.
[5]
J. Energy
Power Eng., vol. 9, no. 4, pp. 326335, 2015.
[6] tion of a micro gas
Thesis, University of Stellenbosch, 2013.
[7] - or Small- Gas
15, 2003.
[8]
Blades in Int. J. Trends
Eng. Technol., vol. 5, no. 2, pp. 170173, 2015.
[9] F. Fuchs, V. Meidinger, N. Neuburger, T. Reiter, M.
small jet engines - fuel distribution and atomization,
in International Symposium on Transport Phenomena
and Dynamics of Rotating Machinery, 2016, pp. 18.
[10]
Educ. Notes RTO-
EN-AVT-131, pp. 114, 2005.
[11]
and operation of a small turbojet-engine for research
Appl. Energy, vol. 84, pp. 11021116,
2007.
[12] J. J. Lee, J. E. Yoon, T. S. Kim, and J. L. Sohn,
ristics
J. Mech. Sci.
Technol., vol. 21, no. 1, pp. 141152, 2007.
[13] J. E. Yoon, J. J. Lee, T. S. Kim, and J. L. Sohn,
turbine and the use of neural network for predicting
J. Mech.
Sci. Technol., vol. 22, no. 12, pp. 25162525, 2008.
[14]
performance model calibration of a small turbojet
Proceedings of the Institution of
Mechanical Engineers, Part G: Journal of Aerospace
Engineering, 2011, pp. 123.
[15] -Thermal-
Structural Coupled Analysis of a Radial Inflow Micro
Gas Turbine Using Computational Fluid Dynamics
and Comput Hindawi Publ.
Corp. Math. Probl. Eng., pp. 111, 2014.
[16]
Identification and Performance Improvement to a
258,
2009.
[17] C. Yang, C. Lee
analyses and experiment investigations of an annular
micro gas turbine power system using fuels with low
Sci. China Ser. E Technol. Sci., vol.
52, no. 12, pp. 35653579, 2009.
[18] H. Asgari, X. Chen, an
Int. J. Model. Identif.
Control, vol. 20, no. 3, p. 253, 2013.
[19] Int.
J. Theor. Appl. Res. Mech. Eng., vol. 2, no. 3, pp.
116120, 2013.
[20] O. Lonard, J. P. Thomas, and S. Borguet, Ten
Years of Experience With a Small Jet Engine as a
J. Eng. Gas Turbines Power,
vol. 131, no. 1, p. 012303, 2009.
[21] C. Nicholas, Jet Propul A Simple Guide to the
Aerodynamic and Thermodynamic Design and
Performance of Jet Engines. 2009.
[22] E.-S. Ahmed, F, Aircraft propulsion and gas turbine
engines. CRC Press, Taylor & Francis Group, U.S.,
2008.
[23]
and Diagnostics of a Small Gas Tu
International Gas Turbine Congress,Tokyo., 2003,
pp. 15.
[24]
377,
2002.
[25]
s, University of
Stellenbosch, 2016.
[26]
microturbine based propulsion systems for civil
[27]
turbines: Numerical study of the effects of heat
transfer and Reynolds number on the performance of
Int. J. Comput. Mater. Sci. Eng.,
vol. 3, no. 2, pp. 115, 2014.
[28] N. Gourdai
Elsevier, vol. 106,
pp. 6778, 2015.
[29] -66
Steady and unsteady
Reynolds-averaged Navier-
Proceedings of the Institution of Mechanical
Engineers Part G: Journal of Aerospace Engineering.,
2016, pp. 114.
[30] C. A. Gonzalez, K. C. Wong, and S. Armfield,
micro-turbine engine
combustor using large eddy simulation and Reynolds
422,
2008.
[31]
Educ. Notes
RTO-EN-AVT-131, no. 1, pp. 118, 2005.
[32] T. Verstraete, R. A. Van den Braembussche, and Z.
The Fourth International Workshop on
Micro and Nanotechnology for Power Generation and
Energy Conversion Application, 2004, pp. 8487.
[33] T. Verstraete, R. A. Van den Braembussche, and Z.
European Conference on
Computational Fluid Dynamics, 2006, pp. 115.
[34] T. Verstraete, Z. Alsalihi, and R. A. Van den
J. Turbomach., vol.
129, pp. 835841, 2007.
[35] -adiabatic fl
ow characteristics of micro imPropuls. Power
Res., vol. 3, no. 1, pp. 18, 2014.
[36] H.Cohens, G.F.C.Rogers, and
H.I.H.Saravanamuttoo, Gas turbine theory, Fourth
Edi. Longman Group Limited, England, U.K, 1996.
[37] A. P. Camacho, J. J. I. Wang, and J. M. Protz,
ts on The Efficiency of Micro Scaled
4, 2011.
[38] A. Calabria, R. Capata, M. Di Veroli, and G. Pepe,
-micro gas turbine devices ( 1 - 10
kW ) for portable power generation at university of
481489, 2013.
[39] S. L. Dixon, Fluid mechanics and thermodynamics of
turbomachinery seventh edition fluid mechanics and
thermodynamics of turbomachinery, Seventh Ed.
Elsevier Inc. USA, 2014.
[40] Investigation and Design
of Low Reynolds Number Micro-
Technology., 1997.
[41]
Aerodynamic Crossover Diffuser for Micro Gas
Thesis, University of
Stellenbosch, 2015.
[42] M. Sajesh, P. . Jyothi Sankar, K. Vishnu Prakash, and
[43] T. Witkowski, S. White, C. O. Dueñas, P. Strykowski,
and T. Sim
SR- Proceedings of the 2003
American Society for Engineering Education Annual
Conference & Exposition, 2003, pp. 121.
[44] R. Amirante, L. A. Catalano, A.Dadone, and V. S. E.
gn optimization of the intake of a small-
CMES, vol. 18, no. March, pp.
1730, 2007.
[45]
16th Australasian Fluid Mechanics Conference, 2007,
pp. 961966.
[46]
CFD analysis of centrifugal compressor for a
[47] -
Chinese J. Aeronaut., vol.
23, no. 3, pp. 298305, 2010.
[48]
of the micro-
ICCES,
vol. 14, no. 1, pp. 122, 2010.
[49] B. B. van der Me
Compressor Impeller for Micro Gas Turbine
Stellenbosch, 2012.
[50]
Masters Thesis, University of Stellenbosch, 2014.
[51]
Int. J. Mech.
Eng. Robot. Res., vol. 3, no. 4, pp. 493506, 2014.
[52] -Flow
Compressor Impeller for Micro Gas Turbine
Stellenbosch, 2016.
[53] -over Diffuser for a
University of Stellenbosch, 2017.
[54] S. Farokhi, Aircraft propulsion, Second Edi.
JohnWiley & Sons Ltd, U.K., 2014.
[55] A. Lefebvre, D. Ballal, and D. Bahr, Gas turbine
combustion: alternative fuels and emissions, Third
Edit. CRC Press, Taylor & Francis Group, U.S., 2010.
[56] Ridzan.J.J.M, EldrainyY.A., and Jaafar.M.N.M,
63, 2008.
[57] J. Guidez, P. Roux, N. Poirson, E. Jourdanneau, M.
miniaturised combustor for application to micro gas
Prog. Propuls. Phys., vol. 1, pp. 469480,
2011.
[58] K. V Chaudhari, D. B. Kulshreshtha, and S. A.
Combustion Chamber with Kerosene as Fuel for 20
Int. J. Eng. Res. Appl., vol.
2, no. 6, pp. 16411645, 2012.
[59] G. C. Krieger, A. P. V. De Campos, F. L. S. Filho, and
chamber for a micro-gas turbine fuelled with natural
J. Brazilian Soc. Mech. Sci. Eng., vol. 34, no. 4,
pp. 441449, 2012.
[60] -Constant
Fourth Australasian Unmanned Systems Conference,
2014, pp. 17.
[61] X. Gao, F. Duan, S. C. Lim, and M. S. Yip,
al and numerical study of flow structure
in Proceedings of the 15th International Heat Transfer
Conference, 2014, pp. 114.
[62] T. Suchocki, P. Lampart, and P. Klonowicz,
GTM-140 turbojet
De Gruyter, vol. 5, pp. 478484, 2015.
[63]
on the Flow through a Micro-Gas Turbine Combustion
Stellenbosch, 2016.
[64] T. Giampaolo, The Gas turbine handbook: Principles
and Practice, Third Edit. CRC Press, Taylor & Francis
Group, U.S., 2006.
[65] -point, off-design
meanline analysis and CFD computations of the axial
turbine to micro gas turbine eng
2009.
[66] -
University of Stellenbosch, 2014.
[67]
stage of a specific micro-
[68] A. S. B. Ennil, S. Mahmouda, R. K. Al-Dadaha, and
A. M. Al-
Proceedings of 22nd
TheIIER International Conference, London, United
Kingdom, 2015, pp. 1619.
[69]
Int. J. Innov. Eng.
Technol., vol. 7, no. 3, pp. 154159, 2016.
[70]
of Johannesburg, 2016.
[71] P. Akbari, R. Nalim, and N. M
ller, Performance
Enhancement of Microturbine Engines Topped With
J. Eng. Gas Turbines Power, vol. 128,
no. 1, p. 190, 2006.
[72] F. Iancu, J. Piechna, E. Dempsey, and N. Müller,
- The Fifth
International Workshop on Micro and
Nanotechnology, 2005, pp. 9396.
[73] E. Dempsey, N. Müller, P. Akbari, and M. R. Nalim,
- 4th International Energy
Conversion Engineering Conference and Exhibit .,
2006, pp. 114.
[74]
The 11th
International Conference on Vibration Engineering,
2005, pp. 17.
[75]
the wave disk micro- Int. J. Gas
Turbine, Propuls. Power Syst., vol. 2, no. 1, pp. 18,
2008.
[76]
Performance Enhancement Through Use of a Wave-
[77]
Rotor Topped D
in The International Gas Turbine and Aeroengine
Congress & Exhibition, 1996, pp. 18.
[78]
of Small Gas Turbine Engines Topped With Wave
39th AIAA/ASME/SAE/ASEE Jt. Propuls.
Conf. Exhib. 20-23, no. July, pp. 111, 2003.
[79]
simulation of micro-
[80]
investigation into the effect of engine bleed on
performance of a single-
Proceedings of the Institution of Mechanical
Engineers, Part G: Journal of Aerospace Engineering,
2008, vol. 222, no. 7, pp. 939949.
[81] M. Badami, P. Nuccio, and A. Signoretto,
-scale
Energy Convers. Manag., vol. 76,
pp. 225233, 2013.
[82]
modeling
[83] -
Dynamic Modeling of Gas Turbine Engine
449,
2014.
[84] A. M. Elzahaby, M. K. Khalil, and H. E. Khalil,
nd Experimental Analysis of a Micro
Int. J. Sci. Eng.
Res., vol. 7, no. 1, pp. 404410, 2016.
[85] I. Lewis R, Turbomachinery performance analysis.
Elsevier Science & Technology Books, 1996.
[86] ernal Combustion Wave
The
International Gas Turbine Congress, 2003, pp. 16.
[87]
NATO Res.
Technol. Organ., pp. 156, 2005.
