1
DEGRUYTER
Francis Oppong*, Sybrand Johan van der Spuy, Theo von Backström.
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 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 performance of its individual components. The effectiveness 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 engine gas generator units affect the mutual performance of one another and the overall engine output. Thus, insufficient turbine work and low component pressure ratio 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 performance. The objectives of the study are to provide a fundamental overview of the methodologies used in the modelling of micro gas turbine performance. The literature 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 conventional gas turbine into smaller engines were described. The second part of the paper discussed the various methodologies that have been employed in investigating and improving MGTs performance output. The paper 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. Keywords: Micro gas turbine, performance, thermal
efficiency, output
1 Introduction The use of micro gas turbine (MGT) engines is becoming increasingly popular in the commercial avia-
Correspondent Author: Francis Oppong, University of Stellenbosch, South Africa, Department of Mechanical and Mechatronic Engineering. Email:
[email protected]
tion and hobby industries. They are employed in unmanned aerial vehicles (UAVs) used in missions related to national security, telecommunications, realtime 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. Some of the technical barriers that impede the engine 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 characteristics of the engine allied with relatively high operating temperatures, very low engine component pressure 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 elements. They evaluated the components characteristic effects on the engine performance output. Experimental and theoretical measurements of MGT thermo-mechanical parameters have been investigated by Bakalis and Stamatis [14]. In the investigation by Xie et al. [15], a three-dimensional fluid thermal structural coupled analysis for a radial inflow micro gas turbine is conducted. Tip clearance losses contributed to inhomogeneous flow in the impeller blades. The hot turbine
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heat flux heats the cold air in the compressor via conduction, hence reducing the mass flow rate through the engine, efficiency, and pressure rise. Combustion interrelated drawbacks such as effective design of combustion chambers to achieve flame stability and recirculation zones in the chamber also contributes to MGTs low performance outputs. The implementation of different and new designs of combustors has been investigated as well as the use of alternative 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 MW54 micro gas turbine. They sought to confirm that there are no hot spots in the liner walls and measure the maximum exhaust temperatures when the engine operates under variable conditions. From the above brief discussions, the geometry of MGTs and its accompanied deficiencies 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 numerous 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 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 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 back of the engine.
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 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 produce 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 propelling nozzle accelerates the exhaust gases from the turbine to create thrust for propulsion.
2. MGT Thermodynamic Working Cycle 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, openair cycle the engine working fluid exits/exhausts into the atmosphere after expansion in the turbine and the exhaust propelling nozzle.
Fig. 2: T-s diagram for Brayton open-air cycle.
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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
ηc is the compressor total-to-total efficiency Knowing the compressor parameters, the combustion outlet stagnation temperature is estimated as:
T T 04
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): .
T T 05
04
mc T T . mc a
pa
g
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 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:
c
p p
03
(1)
00
1
T c 1 T T 00
c
03
00
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 ṁg is the mass flow rate of gas
Fig. 3: p-v diagram for Brayton open-air cycle
03
(2)
where, πc is the compressor total-total pressure ratio 𝑝00 is the compressor inlet stagnation pressure 𝑝03 is the compressor downstream stagnation pressure T00 is the compressor upstream stagnation temperature T03 is the compressor outlet stagnation temperature
With known turbine parameters, the engine exhaust velocity is determined.
u 2c T T
e
pg
05
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 ma ue u0 (6) where, u0 is the freestream velocity Fnet is the net thrust The net power is calculated
P W W net
T
C
where, WT is the turbine work
(7)
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Wc is the compressor work Pnet is the net power
3. MGT Performance Modelling The compressor and turbine maps shown in Figs 4 and 5 are used to model the engine performance at the design and off-design points. The maps consist of efficiency 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]. Accordingly, the impact of a change in the engine’s nonlinear thermodynamic behaviour characteristics on the engine performance is examined at the off design state.
Fig. 4: Compressor map [22]
Fig. 5: Turbine map [22]
Ahmed [22] emphasise that the engine performance at off-design conditions should be satisfactory and safe irrespective 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, condition monitoring and degradation analysis [23]. The engine 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 mathematically formulated with linear and non-linear equations into computer programs for performance simulations.
4. MGT 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], [26]. Micro turbomachinery inherently have low Reynolds numbers due to the small size of the engine, therefore resulting in relative viscous forces dominating the inertia forces in the engine [27]. 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], [14]. The engine’s low Reynolds number can lead to flow separation and transition in the compressor [28], and consequently 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] assessed the effects of Reynolds number on an MGT’s miniaturised compressor. It was found that the compressor efficiency and pressure ratio decreases with decreasing Reynolds number. This is due to an increased entropy change, hence increasing the viscous losses. The engine internal heat transfer influences the engine 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], 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 [14]. A detailed analysis of internal
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heat transfer of small aero engine using 3D NavierStoke solver has been provided by Verstraete et al. [32]. 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. [33] and Verstraete et al. [34] 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 a large temperature difference between 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 nonadiabatic flow characteristics to evaluate the flow phenomenon and to quantify the physical mechanisms leading to heat transfer in MGT compressors. The results 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 that heat addition is a major contributing factor of MGT compressor performance 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 proposed 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 temperature (TIT) of the engine and consequently improves the engine performance output. Conventional larger gas turbines cool the turbine blades using bleed air from the compressor [36]. Smaller gas turbines cannot 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 cooling options [37]. Research thus far shows that MGTs turbine inlet temperatures are restricted between 600800°C [12], [38]. 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 to relatively short compressor and turbine blade
heights [29]. The higher the tip clearance the lower the component efficiencies [39] and the higher performance deterioration and fuel consumption of gas turbines [11]. An MGT’s 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 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 therefore high engine losses [26]. Additionally, excessive frictional losses in the engine bearings adversely affect the engine cycle efficiency.
5. MGT Performance Studies Different research projects have been undertaken on micro gas turbines with the purpose of improving their performance. This includes MGT compressor and turbine stage design improvements (new designs and flow physics analyses), new combustor designs, and combustion analysis. The intent of these projects 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 below.
5.1 Compressor Stage Investigations Fig. 6 shows the centrifugal compressor stage of a micro gas turbine engine.
Fig. 6: Compressor stage [41]
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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 pressure. 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 computational 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 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 reduce flow losses. The optimisation, 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 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 [47] discussed the re-design of an 11 cm diameter 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%. 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° produced 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 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 engine vaned diffuser. He aimed to maximise the compressor stage pressure recovery in order to increase the engine's total-tostatic 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 [50] 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 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 alloy materials to design different compressor stages. Vibrational and stress analysis were investigated for the various designs. The analysis showed that titanium alloy offers the best design in terms of safety factor. They stipulated that the centrifugal stress distribution is concentrated at the blade hub. Recently, Burger [41] designed 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 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. [29] implementing a steady and unsteady Reynolds Average Navier-Stoke 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 while 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.
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Diener [52] presented the design of a mixed flow impeller for the compressor stage of a 600N small turbojet engine. A 1-dimensional and multi-point optimization was performed to assess the aerodynamic performance 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 numerical evaluations. According to Kock [53] the optimised diffuser showed static-to-static pressure recovery increase of 21.5% whereas the total-to-total isentropic efficiency and pressure ratio yielded increase of 8.75% and 7.7%.
5.2 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 [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 employed in recent combustor design and analysis processes for reliable modelling of the internal flow behaviour.
Fig. 7: Combustion chamber
Micro gas turbines to date use a 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 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 common challenges faced in the design and performance of this combustors [55]. These parameters ascertain the engine operational range, durability, cost, maintainability and emissions characteristics [22]. Ridzan.J.J.M et al. [56] 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. [57] 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 a combustion efficiency of 80%. Chaudhari et al. [58] 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. [59] 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
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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. [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 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 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 coefficients 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 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 orifice distribution and combustor 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] 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 a 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 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. Basson [63] deal with the investigation of swirl effects on the mass flow distribution and flow structures in a miniature combustion chamber. She assumed constant pressure drop along the liner walls in the computational 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 experimental data.
5.3 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 8 shows the axial turbine stage of a small jet engine.
Fig. 8: Turbine stage [11]
The turbine stage consists of a 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 high-temperature 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 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 increase 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
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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 efficiency compared to the baseline engine. Basson [66] 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 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 total-to-total temperature ratios due to non-uniform distribution of flow conditions at the immediate turbine stage outlet; however, he suggested that extra measuring instrumentation should be integrated into the measuring of the turbine inlet and outlet temperatures. Lakshmy [8] numerically designed and analysed a 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. [68] using ANSYS CFX. The turbine blade was optimised with the multiobjective 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. Yedla et al. [69] have considered the design of an axial turbine blade for a micro gas turbine engine employing 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 turbine 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.
5.4 Wave-Rotor Approach A wave rotor engine uses pressure waves to transfer energy between the engine’s working fluids [71], therefore, increasing the pressure and temperature of the low
energy fluid of the engine. Fig 9 depicts the thermodynamic cycle, thus the T-s diagram of a wave rotor engine and engine without wave disk. As depicted in Fig 9, the wave rotor technique allows pressure gain in the combustion chamber, therefore 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].
Fig. 9: Wave rotor thermodynamic cycle [74]
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 allows higher temperatures in the combustion chamber, thence, increasing the combustor downstream pressure [75]. The wave rotor is embedded between the compressor 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.
Fig. 10: MGT topped with wave rotor [71]
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
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a considerable reduction of 22% in specific fuel consumption (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 temperatures compared to the baseline engine. Iancu et al. [72] described the uses and advantages of a wave rotor in MGTs. According to them, a waverotor 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 [75] presented the numerical investigation of a two-dimensional model wave disk micro-engine taking into account unsteady, centrifugal and Coriolis forces existing in such a configuration. The main goal was to determine the engine gas compression and torque generation. They concur that the main challenge faced by an axial wave disk configuration is scavenging in the channels; the use of a radial wave rotor prevents the scavenging phenomenon.
5.5 Overall System Analysis Due to the complex and difficult geometry and flow physics of the processes occurring in small turbojet engines, various methods of system evaluation have been 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 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.
Fig. 9: CFD analysis of micro gas turbine internal flow structure [79]
Yin et al. [23] evaluated the performance of an APU small gas turbine engine using experimental investigations and 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 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 reported 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 engine. The mass, momentum and energy equations were used to examine the nature of the 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. [81] 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 [82] 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 Numeca CompAero and 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 [82] 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. The author concluded that the numerical results correlate well with the experimental data.
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Krivcov et al. [83] studied the performance of a small gas turbine engine using ANSYS CFX and thermodynamic model ASTRA software. They employed a gasdynamic modelling analysis. The investigation did not account for the presence of leaks and heat transfer 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. Roberto Capata [5] Oppong [25] 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 a thrust augmentation test by increasing and reducing the engine turbine and nozzle exit areas. A 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, generating high exhaust gas 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 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 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’s performance. The authors confirmed that the theoretical calculations and the experimental test yielded good results when compared to the GasTurb simulations.
6. Conclusions 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 application. This paper focuses on a review of treatises of MGT performance studies. The following conclusions are drawn: Gas turbines performance is modelled and predicted at the off-design conditions or state using com-
ponent maps. Off-design offers better performance diagnostic and prediction of the non-linear behaviour of the engine. Component matching performance modelling 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, simplicity, 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 established for the scaled MGT engine and the larger gas turbine 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 performance output of the engine. Research shows that Reynolds number, heat transfer and thermal losses effects dominate in the gas generator units and significantly 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 performance are few; therefore, this subject area of MGTs performance evaluation still needs to be further investigated. 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 turbine to the cold compressor results in heat addition in the cold compressor via conduction. This leads to large temperature difference between the turbine and compressor, subsequently resulting in reduction of the system mass flow rate, efficiency, and pressure rise. Additionally, 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 recuperator 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.
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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 separation in the curved passages hence the volumetric flow rate of the fluid have to be increased in order to decrease this effect [45]. Redesign or reconfiguration of the compressor stage of the engine can substantially improve the compressor efficiency. Mixed flow compressor stage designs and other design configurations have been adopted for the stage to improve the overall performance of the engine. Efficiencies of between 7585% [1], [6], [41], [45], [50] and beyond can be achieved by redesigning the compressor stage. Additionally, redesigning and optimization of the compressor stage improves the stage pressure ratio and the fuel consumption of the engine. While, redesign of the compressor stage improves the compressor efficiency and pressure ratio, it is appropriate to consider the turbine stage parameters in the design process for smooth running of the engine. Thus, increasing the compressor stage parameters without considering the turbine stage could result in compressor-turbine mismatching. Consequently, the turbine supplies insufficient work to the compressor and could lead to excessive fuel consumption 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 turbine design engineers. The material used to manufacture MGT turbine blades is Inconel IN713LC, a nickelbased super-alloy. The yield strength of the material deteriorates under high temperatures. Therefore, the turbine blade inlet temperature is limited between maximum 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 engine output. In addition, redesigning the turbine stage with high load and flow coefficients enhances the turbine power and the overall power and thrust of the engine [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 complexities related to the flow dynamics. Numerical analyses and computational modelling is suitable for determine the combustion chamber material, liner wall size and improve the efficiency of micro combustion chamber 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 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 [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 turbine inlet and exhaust to assess the engine performance remains 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 [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 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 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 require to drive the compressor [80]. This results in higher turbine inlet and exhaust temperatures due to an increase in fuel consumption.
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