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DEGRUYTER
Francis Oppong*, Sybrand Johan van der Spuy, Theo von Backström, Abdullatif Lacina Diaby.
AN OVERVIEW OF MICRO GAS TURBINE ENGINE PERFORMANCE INVESTIGATION Abstract: 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 electrical 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 performance of its individual components. The effectiveness of MGT is limited due to non-linear behaviour of the engine elements. As at present, a substantial number of research and papers have been published in this area to assess and improve the performance of MGTs. This paper intends to present an outlook of MGTs system performance analysis and improvement evaluation-based methods available in open literature. Keywords: Micro gas turbine, performance, thermal
efficiency
treatises in which micro gas turbine performance analyses has been described and reviewed to different extents.
1.1 Micro Gas Turbine Engine The micro gas turbine engine shown in Fig. 1 is a single
spool (shaft) rotary engine that extracts energy from the flow of micro combustion gases to produce thrust or power. MGTs have thermal efficiency of between 10% and 30% [11], [12] and power and thrust in the ranges of 15-200 kW [13] [14] and 30-200 N [15] at a rotational speed of 20 000-150 000 rpm [1]. The engine consist of a centrifugal compressor at the engine upstream, equipped with a radial or crossover vanned diffuser, an annular straight through, or reverse flow combustor fixed between the compressor and the axial flow turbine, and a fixed convergent propelling nozzle at the downstream of the engine.
1 Introduction The use of micro gas turbine (MGT) engine is increasingly becoming popular technology in the commercial aviation and hobby industries. They are employed in unmanned aerial vehicles (UAVs) used in missions such as national security, telecommunications, real-time reconnaissance, remote sensing, crime fighting, disaster management, agriculture and election monitoring [1]–[3]. MGTs are also used in hybrid electric vehicles 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 energy costs and emissions [7]–[9]. MGTs can be regarded as a prospective and compact competitor to the other propulsion system power supplies such as reciprocating engines and battery cells [5], [10]. These engines have interrelated components, which have nonlinear characteristics. Therefore, the overall engine performance depends on the individual engine element's performance. The present paper includes numerous
Correspondent Author: Francis Oppong, University of Stellenbosch, South Africa, Department of Mechanical and Mechatronic Engineering. Email:
[email protected]
Fig. 1: Micro gas turbine engine
During engine operation, the centrifugal compressor compresses and increases the pressure and temperature of the incoming air from its surroundings. The static pressure of the air rises in 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 produce high temperature, pressure, and velocity gases. The turbine expands the high temperature gas from the combustion process to produce mechanical shaft power
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to drive the compressor. The convergent exhaust propelling nozzle accelerates the exhaust gases from the turbine to create thrust for propulsion.
2. MGT Thermodynamic Working Cycle Micro gas turbine operates on the principles of Brayton open-air cycle. The cycle is represented in temperatureentropy (T-s) and pressure-volume (p-v) diagram as shown in Figs. 2 and 3. In a Brayton, open cycle the engine working fluid exit/exhaust 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.
isentropic work done at station 5-8 in the nozzle and finally constant pressure heat rejection from 8-0 into the atmosphere.
2.1 MGT Cycle Analysis 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 upstream) of the engine to the hot section (combustor upstream to nozzle outlet/downstream) of the engine. From Figs. 2 and 3, the compressor pressure ratio and efficiency are determined using these relations:
c
p p
03
00
(1) 1 1
00
T c T T
c
Fig. 2: T-s diagram
03 00 (2) where, πc is the compressor total-total pressure ratio πc is the compressor inlet stagnation pressure πc is the compressor downstream stagnation pressure T00 is the compressor upstream stagnation temperature T03 is the compressor outlet stagnation temperature ηc is the compressor total-to-total efficiency
Knowing the compressor parameters, the combustion outlet stagnation temperature is estimated as:
T T 04
Fig. 3: p-v diagram
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,
03
fLHV cc
c
pg
(3)
where, T04 is the combustor outlet stagnation temperature ηcc is the combustion efficiency f is the fuel to air ratio LHV is the heating value of fuel cpg is the specific heat of gas. Applying the principles of work and speed compatibility, the turbine downstream stagnation temperature is determined using equation (4):
3
.
pa
mc T T T T . mc 05
04
a
g
03
pg
00
m
(4)
where, T05 is the combustor outlet stagnation temperature ηm is the shaft mechanical efficiency cpa is the specific heat of air cpa is the mass flow rate of air ṁg is the mass flow rate of gas With known turbine parameters, the engine exhaust velocity is determined.
u 2c T T pg
e
04
8
(5)
where, ue is the exhaust velocity T8 is the nozzle outlet static temperature Having estimated the nozzle exhaust velocity the engine thrust is found.
F
net
.
m
a
u u e
0
(6)
where, u0 is the freestream velocity Fnet is the net thrust The net power is calculated
P W W net
T
C
(7)
where, WT is the turbine work Wc is the compressor work Pnet is the net power
3. MGTs Performance Challenges Although, MGTs have outstanding advantages over its competitors in the aviation and power environment, 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], [16]. Micro turbomachinery’s are inherent with low
Reynolds number due to smallness of the engine, therefore resulting in highly viscous forces in the engine [17]. This can cause considerable loss of pressure and temperature and other thermodynamic variation in the engine, hence reducing the efficiency and power output of the engine [1], [18]. The engine low Reynolds number can cause flow separation and transition on the compressor [19], and consequently affects the compressor efficiency and pressure ratio [1], [20]. In addition, the 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 [21]. Xiang et al. [17] assessed the effects of Reynolds number on MGT miniaturised compressor. It was found that the compressor efficiency and pressure ratio decreases with decreasing Reynolds number. This is due to increase entropy change, hence increasing the viscous losses. The engine internal heat transfer influences the engine power output and thermal efficiency. Whereas MGTs high power to weight presents significant advantages, the higher surface to volume ratio creates heat transfer complexities in the engine [16], [22]. Therefore, resulting in thermal losses in the engine combustion chamber, hence, limiting the combustion efficiency and flammability as well as the overall operating efficiencies of the engine [18]. A detailed analysis of internal heat transfer of small aero engine using 3D Navier-Stoke solver has been provided by Verstraete et al. [23]. The heat flux from the turbine to the compressor and the components adiabatic and diabatic performance were used as performance inputs. They examined the impact of recuperative and non-recuperative cycles on power output and efficiency of the MGT. According 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. [24] and Verstraete et al. [25] numerically 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 large temperature difference between 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. [26] implemented a computational fluid dynamics (CFD) based comparison study of adiabatic and non-adiabatic flow characteristics to evaluate the flow phenomenon and to quantify the physical mechanisms owing to heat transfer effect in MGT compressors. The results on the impact of heat transfer 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 temperatures of 450K and 600K, showing
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that heat addition is a major contributing factor of MGT performance deterioration. The heat transfer led to increase in relative Mach number at the impeller hub region, which resulted in flow separation. The analysis showed that heat addition increases angle of tip clearance flow and therefore reduces the slip factor. They proposed that tip clearance flow blockage and relative Mach number should be accounted for the correction of slip factor in compressor impeller design. In general, effective and efficient cooling significantly increases the turbine inlet temperature (TIT) of the engine and consequently improves the engine performance output. Conventional larger gas turbines cool the turbine blades using bled air from the compressor [27]. Smaller gas turbines hardly accommodate this cooling technique and are susceptible to lower turbine inlet temperature. The turbine inlet temperature depends on the turbine blade alloy stress rupture and low cycle fatigue strengths, and cooling options [28]. Research thus far shows that MGTs turbine inlet temperatures are restricted between 600-800°C [12], [29]. In order to accomplish high temperatures at the turbine inlet high temperature alloys or ceramics are required for the turbine design. Higher tip clearance losses are intrinsic in MGTs due short compressor and turbine blade heights [20]. The higher the tip clearance the lower the component efficiencies [30] and the higher performance deterioration and fuel consumption of gas turbines [13]. MGTs compressor has low-pressure ratios 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 turbo-machines always have lower efficiencies than larger geometrically similar machines [31]. 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, therefore having high engine losses [16]. Additionally, excessive frictional losses in the engine bearings adversely affect the engine cycle efficiency.
4. MGTs Performance Studies Different academic research projects have been undertaken on micro gas turbines with the purpose of improving the performance. This includes MGTs compressor and turbine stage design improvements (new designs and flow physics analyses), new combustor designs and
combustion analysis. The intent of these researches was to optimise the engine performance parameters such as the specific fuel consumption, pressure ratio, cycle peak temperature, system pressure losses, turbine and compressor efficiencies and the engine thrust and power. Some of the techniques considered for the performance improvements are discussed.
4.1 Compressor Stage Investigations Compressor is a turbo machinery device that increases the pressure of a fluid, thus it provides high pressure and high volume air. A centrifugal compressor is an aerodynamic type of compressor and is used for low power applications such as micro gas turbine applications. They also find wide usage in the petrochemical industry due to their smooth operation and reliability. Fig. 4 shows the centrifugal compressor stage of a micro gas turbine engine. The compressor consists of impeller (usually scaled from a turbocharger impeller) equipped with a radial (wedge) or crossover (curved) diffusers (shown in Fig 4) to reduce the velocity of the airflow and increase the static pressure.
Fig. 4: Compressor stage
These mini centrifugal compressors are inherent with low efficiencies due to low Reynolds Number, blade tip leakage, growth of boundary layer and separation on the blades, therefore having adverse effect on the engine performance output. In recent years, researches have been directed towards increasing the performance of MGT compressors through redesigning and through computational modelling and simulations. Amirante et al. [32] studied and optimised the Pegasus small turbojet engine compressor intake. They sought to analyse and improve the effects of boundary layer growth and velocity profile using ANSYS FLUENT incorporated with the progressive optimisation technique. The numerical 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 inlet the airflow rate should be increased to improve the velocity profile and hence reduces flow losses. The optimisation, which was aimed at mitigating the intake
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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. [33] 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. [34] deals with the design and computational fluid dynamics 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 [35] discussed the re-design of an 11 cm MGT compressor diffuser. They sought to investigate the effect of cross-sectional area distribution along the flow path of the new diffuser's performance. 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 recovery 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% decrease. In the study by Tough et al. [36], the performance of MGT compressor impeller was evaluated using different inlet blade angles. The writers studied the effects of different blade inlet angle and backward sweep on the compressor performance. Tough et al. [36] found that reducing the backsweep and the inlet blade angle by 3° and 2° produced stable operating range, high-pressure ratio, and total efficiency. Van der Merwe [37] designed and optimised a centrifugal compressor impeller of the BMT engine. 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 impeller 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 120 KS engine vaned diffuser. He aimed to maximise the compressor stage pressure 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 diffuser increased from 1.39 to 1.44. De Villiers [38] employed 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 ouput of 170 N at a rotational speed of 119 000 rpm compared to the original BMT engine. Alex & M [39] report on the design of MGT compressor stage implementing ANSYS FLUENT. They used titanium, aluminium, and stainless steel alloy materials to design different compressor stages. Vibrational and stress analysis were investigated for the various respective designs. The analysis showed that titanium alloy offers the best design in terms of safety of factor. They stipulated that the centrifugal stress distribution is concentrated on the blades hub. Recently, Burger [40] designed and optimised a crossover diffuser for the BMT engine. The crossover diffuser combined with Van der Merwe's [37] impeller improved the compressor stage total-to-static pressure ratio from 2.62 to 3.65. An interpolated 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. [20] implementing the steady and unsteady Reynolds Average NavierStoke CFD solver. They intended to study the compressor transient flow behaviour during operation. The simulation results under predicted the total-to-total adiabatic efficiency showing a decrease of the peak efficiency from 0.73 to 0.55 whiles the total peak pressure ratio increased 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 performance output.
4.2 Combustion Chamber Analysis The main function of a combustion chamber is to increase 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 [41]. 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 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. The CFD tool has been employed in recent combustor design and analysis for reliable modelling of the internal flow behaviour.
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Micro gas turbines to date uses constant pressure combustor. They are simple to implement and provide a relatively steady, uniform gas flow to the turbine. Although, they have practical benefits the flow behaviour in the combustor is complex due compactness and high viscous environment. The combustion efficiency, flame stability and ignition, wall cooling, pressure loss and emission control are some of the common challenges faced in the design and performance of this combustors [42]. These parameters ascertain the engine operational range, durability, cost, maintainability and emissions characteristics [43].
Fig. 5: Combustion chamber
Ridzan.J.J.M et al. [44] have considered the aerodynamic 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 increases swirling in the combustion chambers, therefore increasing turbulence strength, recirculation zone size and the amount of recirculation mass at increasing pressure loss in the combustor. Guidez et al. [45] described the studies performed on combustion characteristics of a miniature combustor. The aim was to examine combustion stability and efficiency. Raman spectroscopy and 1D Rayleigh scattering and standard thermodynamic measurements were used to estimate the temperature profile and main species concentration at the combustor outlet. They measured combustion efficiency of 80%.
Chaudhari et al. [46] report the design and simulation of a miniature annular combustion chamber using ANSYS CFX. They evaluated the impact of flow patterns 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. Further, Gieras & Stankowski [4] have performed computational modelling of the aerodynamic flow in the GTM-120 micro jet engine combustor. They considered the effect of engine downsizing on the mass flow, pressure losses, and heat transfer in the combustor. The aim
of their research was to maximise thermal efficiency by minimising fuel consumption and controlling emissions. 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 combustor. The authors indicated that the consequences of excessive 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. [47] numerically and experimentally evaluated a micro jet engine combustion chamber using ANSYS FLUENT RANS turbulent solver. They accessed 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 configuration 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 increases air-fuel mixing processes. According to Krieger et al. [47], the numerical and experimental results showed good correlation. Armstrong & Verstraete [48] assert that a constant volume combustor increases the pressure in the chamber and reduces engine fuel consumption 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 upstream 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 influence the average combustion pressure ratio and discharge coefficients below 3.0 are required to increase the pressure ratio. Gao et al. [49] examined the performance of the SR-30 micro turbine combustor. They analysed combustion flow characteristics such as liner wall temperature, pressure and temperature distribution in the chamber. 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 insufficient cooling air after complete combustion. The investigations showed that orifices distribution and combustor geometry dominates the flow pattern inside combustion chamber. According to Gao et al. [49], hot areas and/or spots in the liner walls are dominate in the
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primary and dilution zones. Suchocki et al. [50] analysed the combustion performance characteristics of the GTM-140 mini turbojet engine combustor employing computational investigations. They sought to understand and identify the chemical process in the combustor. The combustion was modelled using the k-ϵ (RANS) Turbulence Model and Non-Premixed Model. The performance investigations showed pressure drop of 9%, however, they proposed that modification of the diffuser geometry should be considered for further deceleration of the velocity. Fuchs et al. [9] have considered the flow characteristics of a small aero engine combustor using experimental and numerical evaluations. The computational investigation was performed employing CFD commercial software ANSYS CFX. Their objective was to study the behaviour of fuel distribution 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 temperatures.
4.3 Turbine Stage Assessment The turbine extracts kinetic energy from the combustion hot gases to produce mechanical shaft power to drive the compressor. The turbine efficiency depends on how well it extracts kinetic energy from the combustion hot gases. Fig 6 shows the axial turbine stage of a small jet engine..
Fig. 6: Turbine stage [13]
It consists of rotor (left) and stator (right). The incoming combustion hot gas is initially accelerated 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 a high-temperature and corrosion [41]. The blades are subjected to high temperatures, which can cause creep and corrosion failure of the blade without effective cooling [51]. Wojciech & Michal [52] 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 performance contours for the turbine stage. They contend that turbine tip clearance, Reynolds number, and specific speed coefficient influence turbine efficiency and overall engine performance. Ensuring high Reynolds number and specific speed coefficient as well as low tip clearance increases the turbine work transfer. Verstraete et al. [1] investigated and improved the overall performance 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 the efficiency compared to the baseline engine. Basson [53] re-designed and examined new turbine stages for the BMT 120 KS micro jet engine. He performed empirical analytical and numerical analyses for the new designs. Reverse engineering was used to produce 3D models of the existing axial flow turbine of the engine. A finite element analysis was conducted to determine the structural behaviour and stress distribution of the axial flow turbine under different static and dynamic loadings. He concluded that the proposed designs predicted efficiency difference of 5% between the various new designs. Lakshmy [8] numerically designed and analysed MGT turbine stage. Different flow velocities and pressures at different angle of attacks were used for the investigation. She claims that the engine thrust was increased for the new design. The design and performance analysis of a small axial flow turbine has been conducted by Ennil et al. [54] 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 concluded that the investigations yielded a total-to-total efficiency of 87.78% for the turbine stage.
4.4 Wave Rotor Approach A wave rotor engine uses pressure waves to transfer energy between the engine’s working fluids [55], therefore, increasing the pressure and temperature of the low energy fluid of the engine. Fig 7 depicts the thermodynamic cycle, thus the T-s diagram of a wave rotor engine and engine without wave disk. As depicted in Fig 7, the wave rotor technique allow pressure gain in the combustion chamber, therefore increasing the engine’s thrust and power output as well as the cycle thermal efficiency [56]. In a typical wave
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rotor engine, the hot gases from the combustion compresses the air coming out of the compressor [57]. The burnt gas is pre-expanded in the wave disk, hence delivering hot gas of less temperature to enter the turbine inlet [55]. Nevertheless, the wave disk allow higher temperatures in the combustion chamber, thence, increasing the combustor downstream pressure [58]. The wave rotor is embedded between the compressor and the turbine and parallel to the combustion chamber as shown Fig 8. The wave disk arrangements are of two configurations, namely radial and axial configuration.
with a four-port wave rotor. According to the authors, the outcome of their investigation showed significant improvement in the combustion pressures and temperatures compared to the baseline engine. Iancu et al. [56] described the uses and advantages of wave rotor in MGTs. According to them wave-rotor cycle increases the overall gas turbine compression and expansion ratio. They argued that the wave rotor technique could achieve a 50 to 83% increase in compression efficiency for ultra-micro and small gas turbines. Piechna & Dyntar [58] presented the numerical investigation of a two-dimensional model wave disk micro-engine taking into account unsteady, centrifugal and Coriolis forces existing in such configuration. The main goal was to determine the engine gas compression and torque generation. They concur that the main challenge faced by axial wave disk configuration is scavenging in the channels; therefore, the use of radial wave rotor prevents the scavenging phenomenon.
4.5 Overall System Analysis
Fig. 7: Wave rotor thermodynamic cycle [59]
Due to the complex and difficult geometry and flow physics of the processes occurring in small turbojet engine’s, other methods of system evaluation have been extensively used to investigate micro gas turbine engine performance output. This includes computational (CFD), theoretical (simulation and analytical), and experimental analysis of the processes occurring in the engine. The CFD tool is the effective means of studying the mutual behaviour of the flow physics of the engine elements. Fig 9 illustrates computational flow analysis of the flow phenomenon in a gas turbine unit.
Fig. 8: MGT topped with wave rotor [55]
Wilson & Paxson [60] 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 [61] discuss the performance benefits of the wave rotor cycle in a small turboshaft engine. The engine topped with wave rotor showed a considerable reduction of 22% in specific fuel consumption (SFC) compared to the baseline engine. Akbari & Müller [62] 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
Fig. 9: CFD analysis of micro gas turbine internal flow structure [63]
Yin et al. [64] evaluated the performance of an APU small gas turbine engine using experimental investigations and the TURBOMATCH software simulations. The engine exhaust temperature profile showed high fluctuations in comparison with the exhaust pressure. This is due to the presence of hot areas and/or spots at the engine exhaust. They identified that the cycle thermal efficiency drops as the ambient temperature
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rises. Yin et al. (2003) stipulated that the simulation showed good agreement with the experimental test obtaining an accuracy of 2%. Rahman & Whidborne [65] 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. The simulated results and the experimental data showed 5% difference and 10% for nominal and excessive bleed extraction. Trebunskikh et al. [63] assessed the KJ66 micro turbine engine performance output employing FloEFD, CFD software code. They analysed the internal flow structure in engine from the upstream to the downstream of the engine. The mass, momentum and energy equations were used to examined the nature of flow in the compressor, combustion chamber, the turbine and the convergent nozzle. The authors concluded that the numerical results correlate well with the experimental test results. However, they reported that non-uniform fluid temperature distributions at the combustor outlet, and wedge diffuser inefficiencies contribute towards poor performance of the engine. Badami et al. [66] used ANSYS FLUENT to analyse the performance of the SR-30 micro jet engine. The intent was to investigate the engine thermodynamic cycle performance. According to them, the CFD data showed consistent trend with the experimental values. The nozzle and turbine pressure ratios showed differences of less than 5% and the mass flow rate and thrust values showed percentage differences of 2% and 10% respectively. Leylek [67] studied and investigated the performance 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 TurbAero. The main purpose of the study was to understand thermodynamic modelling and performance characteristics of the engine. Infrared modelling and thrust augmentation analysis were also investigated. Leylek [67] blocked the engine convergent nozzle and also reduced the nozzle area to stress the engine to operate at different operating conditions in order to assess the thrust output and compressor and turbine operating maps. It was reported that the engine showed 9-100% increase in the thrust for the reduced nozzle area. The author concluded that the numerical results correlate well with the experimental data. Krivcov et al. [68] studied the performance of a small gas turbine engine using ANSYS CFX and thermodynamic model ASTRA software. They employed gas-dynamic modelling analysis. The investigation did not account for the presence of leaks and heat exchange between the stream and the air-gas wall channels. The
study showed percentage difference of less than 7% between both simulations. They claim that good agreement was obtained between the ANSYS CFX results and the ASTRA thermodynamic 1D results for the engine. Oppong [69] investigated the performance of the BMT 120 KS jet engine using computational, theoretical, and experimental evaluations. Flownex SE and GasTurb simulation software were used for the numerical analysis. The aim of the study was to examine matching of the engine components. He performed thrust augmentation test by increasing and reducing the engine turbine and nozzle exit areas. Sensitivity analysis was conducted using different modified (redesigned) compressor and turbine stages of the engine. The simulations revealed that the turbine capacity was insufficient to drive the compressor of the engine. Therefore generating high exhaust temperatures EGT in an attempt to increase the fuel consumption to increase the turbine power. The experimental analyses indicated the engine turbine inlet and exhaust showed high temperature fluctuations due to the presence of hot spots and/or area in the turbine inlet and exhaust. Elzahaby et al. [70] using theoretical simulations and experimental investigations as well as the GASTURB software simulations evaluated the thermodynamic parameters 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 parameters. The engine component maps combined with the gas-dynamic equations were used to assess the engine performance. The authors confirmed that the theoretical calculations and the experimental test yielded good results when compared to the GASTURB simulations.
5. Conclusions Micro gas turbine engine is widely seen as one of the key technology in the hobby industry as well as for small combined heat and power generation application. This paper focuses on the review of treatises of MGT performance studies. The following conclusions are drawn: The entropy change of small gas turbines increases due to downsizing and/or scaling of larger gas turbine into smaller engine. Therefore, the Reynolds number of the engine decreases and increases the viscosity of the flow in the engine and consequently limits the engine overall thermal efficiency and the performance output of the engine. The wedge type of diffusers is found to contribute to the greater loss of efficiency in the compressor unit.
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Also, the curved or crossover diffuser is inherent with flow separation at the curved passage’s hence the volumetric flow rate of the fluid have to be increased in order to decrease this effect [33]. Research, thus far shows that Reynolds number effects dominate in the gas generator units (compressor, combustor, and the turbine) whereas heat transfer and thermal losses significantly affect the combustion chamber performance. Accordingly, most of the performance analyses on MGTs have been considered on the aerodynamic flow behaviour and physics in the combustion chamber. Constant pressure combustors are mostly used in MGTs application [48]. These burners are susceptible to pressure loss. Therefore, pressure gain techniques such as the use of constant volume combustors and the implementation of wave rotor methods are considered. The constant volume combustion chamber reduces the pressure drop in the combustor and increases the combustor downstream pressure, hence contributing to overall increase of the cycle efficiency and performance of the engine. Likewise, the use of wave rotor method or approach increases 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 [71]. 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 turbine inlet and exhaust to assess the engine performance remains as a challenge due to the high temperature fluctuations. Therefore, in order to measure reliable turbine inlet and exhaust temperatures perfectly shielded thermocouples are required. In addition, careful attention must be paid to some unacceptable temperature gradients through the outlet section of the combustor as it appears the most challenging problem to be dealt with in terms of combustor fluid dynamics [72]. 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 engine 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 lowers the mass flow through the engine, therefore the turbine will have to operate at a lower mass
flow to produce the power require to drive the compressor [65]. This results in higher turbine inlet and exhaust temperatures due to increase in fuel consumption.
References [1]
[2]
[3]
[4]
[5]
[6]
[7] [8]
[9]
[10]
[11]
[12]
[13]
[14]
[15]
[16]
[17]
[18]
D. Verstraete, P. Hendrick, V. Djanali, C. Gonzalez, J. Ling, K. C. Wong, and S. Armfield, “Micro Propulsion Activities at the University of Sydney.,” in PowerMEMS 2010 Conference, Australia, 2010, pp. 1–4. G. B. Bruening, J. R. Snyder, and R. E. Fredette, “The potential impact of utilizing advanced engine technology for a combat capable unmanned air vehicle (UAV),” J. Eng. Gas Turbines Power, vol. 123, no. 3, pp. 508–512, 2001. R. Capata, L. Marino, and E. Sciubba, “ARTICLE A hybrid propulsion system for a high-endurance UAV : configuration selection , aerodynamic study , and gas turbine bench tests,” vol. 35, no. April 2013, pp. 16– 35, 2014. M. Gieras and T. Stankowski, “Computational study of an aerodynamic flow through a micro-turbine engine combustor,” J. Power Technol., vol. 92, no. 2, pp. 68–79, 2012. Roberto Capata, “Experimental tests of the operating conditions of a micro gas turbine device,” J. Energy Power Eng., vol. 9, no. 4, pp. 326–335, 2015. D. S. Krige, “Performance evaluation of a micro gas turbine centrifugal compressor diffuser,” Masters Thesis, University of Stellenbosch, 2013. T. W. Simon and N. Jiang, “Micro- or Small- Gas Turbines,” pp. 1–15, 2003. M. Lakshmy Devi S, “Design and Analysis of Turbine Blades in a Micro Gas Turbine Engine,” Int. J. Trends Eng. Technol., vol. 5, no. 2, pp. 170–173, 2015. F. Fuchs, V. Meidinger, N. Neuburger, T. Reiter, M. Zündel, and A. Hupfer, “Challenges in designing very small jet engines - fuel distribution and atomization,” in International Symposium on Transport Phenomena and Dynamics of Rotating Machinery, 2016, pp. 1–8. R. Decuypere and D. Verstraete, “Micro turbines from the standpoint of potential users,” Educ. Notes RTOEN-AVT-131, pp. 1–14, 2005. H. Asgari, X. Chen, and R. Sainudiin, “Modelling and simulation of gas turbines,” Int. J. Model. Identif. Control, vol. 20, no. 3, p. 253, 2013. J. J. Lee, J. E. Yoon, T. S. Kim, and J. L. Sohn, “Performance test and component characteristics evaluation of a micro gas turbine,” J. Mech. Sci. Technol., vol. 21, no. 1, pp. 141–152, 2007. E. Benini and S. Giacometti, “Design, manufacturing and operation of a small turbojet-engine for research purposes,” Appl. Energy, vol. 84, pp. 1102–1116, 2007. T. Shukla, “Micro gas turbine Engine: A Review,” Int. J. Theor. Appl. Res. Mech. Eng., vol. 2, no. 3, pp. 116–120, 2013. O. Léonard, J. P. Thomas, and S. Borguet, “Ten Years of Experience With a Small Jet Engine as a Support for Education,” J. Eng. Gas Turbines Power, vol. 131, no. 1, p. 012303, 2009. A. Marcellan, “An exploration into the potential of microturbine based propulsion systems for civil Unmanned Aerial Vehicles,” 2015. J. Xiang, J. U. Schlüter, and F. Duan, “Miniature gas turbines: Numerical study of the effects of heat transfer and Reynolds number on the performance of an axial compressor,” Int. J. Comput. Mater. Sci. Eng., vol. 3, no. 2, pp. 1–15, 2014. D. P. Bakalis and A. G. Stamatis, “Data analysis and performance model calibration of a small turbojet engine.,” in Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, 2011, pp. 1–23.
11
[19]
[20]
[21]
[22]
[23]
[24]
[25]
[26]
[27]
[28]
[29]
[30]
[31]
[32]
[33]
[34]
[35]
[36]
[37]
[38]
N. Gourdain, “Prediction of the unsteady turbulent flow in an axial compressor stage . Part 2 : Analysis of unsteady RANS and LES data.,” Elsevier, vol. 106, pp. 67–78, 2015. J. Xiang, J. U. Schluter, and F. Duan, “Study of KJ-66 micro gas turbine compressor : Steady and unsteady Reynolds-averaged Navier-Stokes approach.,” in Proceedings of the Institution of Mechanical Engineers Part G: Journal of Aerospace Engineering., 2016, pp. 1–14. C. A. Gonzalez, K. C. Wong, and S. Armfield, “Computational study of a micro-turbine engine combustor using large eddy simulation and Reynolds averaged turbulence models,” vol. 49, pp. 407–422, 2008. R. A. Van Den Braembussche, “Micro Gas Turbines – A Short Survey of Design Problems,” Educ. Notes RTO-EN-AVT-131, no. 1, pp. 1–18, 2005. T. Verstraete, R. A. Van den Braembussche, and Z. A. Von, “Heat and Power Balance in Micro Gasturbine Rotors,” in The Fourth International Workshop on Micro and Nanotechnology for Power Generation and Energy Conversion Application, 2004, pp. 84–87. T. Verstraete, R. A. Van den Braembussche, and Z. A. Von, “A Conjugate Heat Transfer Method Applied to Turbomachinery,” in European Conference on Computational Fluid Dynamics, 2006, pp. 1–15. T. Verstraete, Z. Alsalihi, and R. A. Van den Braembussche, “Numerical Study of the Heat Transfer in Micro Gas Turbines,” J. Turbomach., vol. 129, pp. 835–841, 2007. Q. Sun, Q. Li, A. Arshad, and Z. Li, “Non-adiabatic fl ow characteristics of micro impeller,” Propuls. Power Res., vol. 3, no. 1, pp. 1–8, 2014. H.Cohens, G.F.C.Rogers, and H.I.H.Saravanamuttoo, Gas turbine theory, Fourth Edi. Longman Group Limited, England, U.K, 1996. A. P. Camacho, J. J. I. Wang, and J. M. Protz, “Scaling Effects on The Efficiency of Micro Scaled Gas Turbines.,” pp. 1–4, 2011. A. Calabria, R. Capata, M. Di Veroli, and G. Pepe, “Testing of the ultra-micro gas turbine devices ( 1 - 10 kW ) for portable power generation at university of Roma 1 : First tests results,” vol. 2013, no. May, pp. 481–489, 2013. S. L. Dixon, Fluid mechanics and thermodynamics of turbomachinery seventh edition fluid mechanics and thermodynamics of turbomachinery, Seventh Ed. Elsevier Inc. USA, 2014. A. Mehra, “Computtational Investigation and Design of Low Reynolds Number Micro-Turbomachinery,” Master’s Thesis, Massachusetts Institute of Technology., 1997. R. Amirante, L. A. Catalano, A.Dadone, and V. S. E. Daloiso, “Design optimization of the intake of a smallscale turbojet engine,” CMES, vol. 18, no. March, pp. 17–30, 2007. J. Ling, K. C. Wong, and S. Armfield, “Numerical investigation of a small gas turbine compressor,” in 16th Australasian Fluid Mechanics Conference, 2007, pp. 961–966. R. T. Aghaei, A. M. Tousi, and M. Soltani, “Design and CFD analysis of centrifugal compressor for a microgasturbine.,” no. April 2014, 2007. C. Jie and H. Guoping, “Redesign of an 11 cmdiameter micro diffuser,” Chinese J. Aeronaut., vol. 23, no. 3, pp. 298–305, 2010. R. A. Tough, A. M. Tousi, and J. Ghaffari, “Improving of the micro-turbine ’ s centrifugal impeller performance by changing the blade angles,” ICCES, vol. 14, no. 1, pp. 1–22, 2010. B. B. van der Merwe, “Design of a Centrifugal Compressor Impeller for Micro Gas Turbine Application,” Master’s Thesis, University of Stellenbosch, 2012. L. C. B. De Villiers, “Design of a centrifugal compressor for application in micro gas turbines,”
[39]
[40]
[41] [42]
[43]
[44]
[45]
[46]
[47]
[48]
[49]
[50]
[51]
[52]
[53]
[54]
[55]
[56]
[57]
[58]
Masters Thesis, University of Stellenbosch, 2014. A. E. Alex and N. M, “3D Modelling and Analysis of Micro Gas Turbine Compressor Blade,” Int. J. Mech. Eng. Robot. Res., vol. 3, no. 4, pp. 493–506, 2014. C. Burger, “Design Procedure of a Compact Aerodynamic Crossover Diffuser for Micro Gas Turbine Application,” Masters Thesis, University of Stellenbosch, 2015. S. Farokhi, Aircraft propulsion, Second Edi. JohnWiley & Sons Ltd, U.K., 2014. A. Lefebvre, D. Ballal, and D. Bahr, Gas turbine combustion: alternative fuels and emissions, Third Edit. CRC Press, Taylor & Francis Group, U.S., 2010. E.-S. Ahmed, F, Aircraft propulsion and gas turbine engines. CRC Press, Taylor & Francis Group, U.S., 2008. Ridzan.J.J.M, EldrainyY.A., and Jaafar.M.N.M, “Prediction of the flow inside a micro gas turbine combustor,” no. 25, pp. 50–63, 2008. J. Guidez, P. Roux, N. Poirson, E. Jourdanneau, M. Orain, and F. Grisch, “Investigation of combustion in miniaturised combustor for application to micro gas turbines,” Prog. Propuls. Phys., vol. 1, pp. 469–480, 2011. K. V Chaudhari, D. B. Kulshreshtha, and S. A. Channiwala, “Design and CFD Simulation of Annular Combustion Chamber with Kerosene as Fuel for 20 kW Gas Turbine Engine,” Int. J. Eng. Res. Appl., vol. 2, no. 6, pp. 1641–1645, 2012. G. C. Krieger, A. P. V. De Campos, F. L. S. Filho, and R. C. De Souza, “A swirler stabilized combustion chamber for a micro-gas turbine fuelled with natural gas,” J. Brazilian Soc. Mech. Sci. Eng., vol. 34, no. 4, pp. 441–449, 2012. J. Armstrong and D. Verstraete, “Semi-Constant Volume Combustion in a Micro Gas Turbine,” in Fourth Australasian Unmanned Systems Conference, 2014, pp. 1–7. X. Gao, F. Duan, S. C. Lim, and M. S. Yip, “Experimental and numerical study of flow structure and liner wall temperature in reverse flow combustor,” in Proceedings of the 15th International Heat Transfer Conference, 2014, pp. 1–14. T. Suchocki, P. Lampart, and P. Klonowicz, “Numerical investigation of a GTM-140 turbojet engine,” De Gruyter, vol. 5, pp. 478–484, 2015. T. Giampaolo, The Gas turbine handbook: Principles and Practice, Third Edit. CRC Press, Taylor & Francis Group, U.S., 2006. B. Wojciech and C. Michal, “Design-point, off-design meanline analysis and CFD computations of the axial turbine to micro gas turbine engine,” vol. 16, no. 4, 2009. G. J. T. Basson, “Design methodology of an axial flow turbine for a micro jet engine,” Masters Thesis, University of Stellenbosch, 2014. A. S. B. Ennil, S. Mahmouda, R. K. Al-Dadaha, and A. M. Al-Jubori, “Optimization of Small Scale Axial Air Turbine Using ANSYS CFX.,” in Proceedings of 22nd TheIIER International Conference, London, United Kingdom, 2015, pp. 16–19. P. Akbari, R. Nalim, and N. Müller, “Performance Enhancement of Microturbine Engines Topped With Wave Rotors,” J. Eng. Gas Turbines Power, vol. 128, no. 1, p. 190, 2006. F. Iancu, J. Piechna, E. Dempsey, and N. Müller, “Ultra-Micro Wave Rotor Investigations,” in The Fifth International Workshop on Micro and Nanotechnology, 2005, pp. 93–96. E. Dempsey, N. Müller, P. Akbari, and M. R. Nalim, “Four-Port Wave Rotors,” in 4th International Energy Conversion Engineering Conference and Exhibit ., 2006, pp. 1–14. J. Piechna and D. Dyntar, “Numerical investigation of the wave disk micro-engine concept,” Int. J. Gas Turbine, Propuls. Power Syst., vol. 2, no. 1, pp. 1–8, 2008.
12
[59]
[60]
[61]
[62]
[63] [64]
[65]
[66]
[67] [68]
[69]
[70]
[71]
[72]
M. HÎRCEAGĂ and F. IANCU, “Wave Rotors Technology and Applications,” in The 11th International Conference on Vibration Engineering, 2005, pp. 1–7. J. Wilson and E. Daniel Paxson, “Jet Engine Performance Enhancement Through Use of a WaveRotor Topping Cycle.,” 1993. P. H. Snyder and R. E. Fish, “Assesment of a Wave Rotor Topped Demonstrator Gas Turbine Concept,” in The International Gas Turbine and Aeroengine Congress & Exhibition, 1996, pp. 1–8. P. Akbari and N. Müller, “Performance Investigation of Small Gas Turbine Engines Topped With Wave Rotors,” 39th AIAA/ASME/SAE/ASEE Jt. Propuls. Conf. Exhib. 20-23, no. July, pp. 1–11, 2003. T. Trebunskikh, A. Ivanov, and G. Dumnov, “FloEFD simulation of micro-turbine engine,” 2012. J. Yin, M. S. Li, and W. Huang, “Performance Analysis and Diagnostics of a Small Gas Turbine,” in International Gas Turbine Congress,Tokyo., 2003, pp. 1–5. N. U. Rahman and J. F. Whidborne, “A numerical investigation into the effect of engine bleed on performance of a single-spool turbojet engine,” in Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, 2008, vol. 222, no. 7, pp. 939–949. M. Badami, P. Nuccio, and A. Signoretto, “Experimental and numerical analysis of a small-scale turbojet engine,” Energy Convers. Manag., vol. 76, pp. 225–233, 2013. Z. Leylek, “An investigation into performance modeling of a small gas turbine engine,” 2013. A. V Krivcov, L. S. Shabliy, and O. V Baturin, “GasDynamic Modeling of Gas Turbine Engine Components Collaborative Workflow,” pp. 445–449, 2014. F. Oppong, “Micro Gas Turbine Performance Evaluation,” Masters Thesis, University of Stellenbosch, 2016. A. M. Elzahaby, M. K. Khalil, and H. E. Khalil, “Theoretical and Experimental Analysis of a Micro Turbojet Engine’s Performance,” Int. J. Sci. Eng. Res., vol. 7, no. 1, pp. 404–410, 2016. R. Nalim and K. Pekkan, “Internal Combustion Wave Rotors for Gas Turbine Engine Enhancement,” in The International Gas Turbine Congress, 2003, pp. 1–6. R. Tuccillo and M. C. Cameretti, “Combustion and Combustors for MGT Applications,” NATO Res. Technol. Organ., pp. 1–56, 2005.
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Correspondent Author: Francis Oppong, University of Stellenbosch, South Africa, Department of Mechanical and Mechatronic Engineering. Email:
[email protected]
DEGRUYTER
