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Propulsion and Power Research

Propulsion and Power Research 2013;2(2):96–106

2212-540X & 2013 Nhttp://dx.doi.org/10.10

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ORIGINAL ARTICLE

Performance assessment of simple and modifiedcycle turboshaft gas turbines

Barinyima Nkoin, Pericles Pilidis, Theoklis Nikolaidis

Department of Power and Propulsion, School of Engineering, Cranfield University, Cranfield, Bedfordshire MK43 0AL, United Kingdom

Received 14 March 2013; accepted 15 April 2013Available online 4 June 2013

KEYWORDS

Gas turbines;Turboshaft;Technicalperformance;Intercooled;Recuperated;Low pressurecompressor (LPC)zero-staged;Simple cycle;Comparativeassessment

ational Laboratory f16/j.jppr.2013.04.00

hor: Tel.: +44 75742

nkoi@cranfield.ac.u

esponsibility of Natnautics, China.

Abstract This paper focuses on investigations encompassing comparative assessment of gasturbine cycle options. More specifically, investigation was carried out of technical performanceof turboshaft engine cycles based on existing simple cycle (SC) and its projected modifiedcycles for civil helicopter application. Technically, thermal efficiency, specific fuel consump-tion, and power output are of paramount importance to the overall performance of gas turbineengines. In course of carrying out this research, turbomatch software established at CranfieldUniversity based on gas turbine theory was applied to conduct simulation of a simple cycle(baseline) two-spool helicopter turboshaft engine model with free power turbine. Similarly,some modified gas turbine cycle configurations incorporating unconventional components,such as engine cycle with low pressure compressor (LPC) zero-staged, recuperated enginecycle, and intercooled/recuperated (ICR) engine cycle, were also simulated. In doing so, designpoint (DP) and off-design point (OD) performances of the engine models were established. Thepercentage changes in performance parameters of the modified cycle engines over the simplecycle were evaluated and it was found that to a large extent, the modified engine cycles withunconventional components exhibit better performances in terms of thermal efficiency andspecific fuel consumption than the traditional simple cycle engine. This research made use ofpublic domain open source references.& 2013 National Laboratory for Aeronautics and Astronautics. Production and hosting by Elsevier B.V.

All rights reserved.

or Aeronautics and Astronautics. Pro9

42014.

k (Barinyima Nkoi)

ional Laboratory for

1. Introduction

Gas turbine engines which provide power for rotary wingaircrafts, or helicopters, are referred to as turboshafts.Turboshaft engines may be single-spool or multi-spool.The single-spool configurations are employed basically to

duction and hosting by Elsevier B.V. All rights reserved.

Nomenclature

c specific heat (unit: kJ/kg)cp specific heat at constant pressure (unit: kJ/kg)h specific enthalpy (unit: kJ/kg)p total pressure (unit: N/m2)q heat flow (unit: kW)qin heat flow in (unit: kW)qout heat flow out (unit: kW)S entropy (unit: kJ/(kg �K))T total temperature (unit: K)Z surge margin parameterturbomatch gas turbine engine performance model code

Abbreviations

ACARE Advisory Council for Aeronautics Research in EuropeCN non-dimensional speedCU Cranfield UniversityCW compression work (unit: kJ)DP design pointEC EurocopterEW expansion work (unit: kJ)FPT free power turbineFPTW free power turbine workGT gas turbineHP high pressureHPC high pressure compressorHPCW high pressure compressor workHPT high pressure turbineHPTW high pressure turbine workICR intercooled/recuperatedISA international standard atmosphereISA Dev international standard atmosphere deviationLP low pressure

LPC low pressure compressorLPCW low pressure compressor workLPT low pressure turbineOD off-design pointOEM original equipment manufacturerOPR overall pressure ratioPR pressure ratioSC simple cycleTERA techno-economic and environmental risk analysisTET turbine entry temperature (unit: K)USAF United State Air ForceZS zero-stagedSFC specific fuel consumption (unit: kg/(MW � s))

Greek letters

ε heat exchanger effectivenessη efficiency (unit: %)ηth thermal efficiency (unit: %)ηc compressor isentropic efficiency (unit: %)ηt turbine isentropic efficiency (unit: %)π pressure ratio at design pointπchoke pressure ratio at choke conditionπsurge pressure ratio at surge condition

Subscripts

p at constant pressureth thermalin inletout outletregen regenerationinter intercooler1,2,3,4,5,6,7 engine components station numbers

Performance assessment of simple and modified cycle turboshaft gas turbines 97

minimize weight. This is particularly true for light weighthelicopters [1].

Free power turbine configuration with a single-spool or insome cases two-spool gas generator characterize turboshaftengines. Pressure ratio is usually in the range of 7:1 to 10:1and with turbine entry temperature (TET) of about 1250 K to1450 K requiring turbine blade cooling. For some mediumand large turboshaft engines the pressure ratio may be about17:1 and turbine entry temperature is about 1500 K, inwhich case, blade cooling is required. For specific power,this pressure ratio of 17:1 is about the optimum [1].

The turboshaft engine is designed so that the speed of thehelicopter rotor is independent of the rotating speed of thegas generator [2]. A schematic illustration of the turboshaftengine and its common configuration with output shaft andgear box assembled is as shown in Figure 1.

Using heat exchangers (both recuperators and inter-coolers) in an engine exhibits tremendous potential to cutfuel consumption and thereby reducing CO2 emissions.Stefan Donnerhack (technical expert at MTU Aero Enginesin Munich and head of the Intercooled Recuperative Aero-Engine (IRA) research team) once explained that the

recuperator utilises part of heat from the exhaust gas toraise the temperature of the air entering the combustor [4].This method makes achievement of the same turbine entrytemperature as in a conventional engine possible, butwith the advantage of utilizing lesser fuel [4]. Besides,gas turbine user requirements have, over the years, necessi-tated technological advancement in engine performance,and comprehensive researches are being conducted toachieve this [5].

Technically, improvement of thermal efficiency forindustrial and aero gas turbines is of paramount importanceto the overall performance of the engines. Increase inthermal efficiency depends on certain factors including:changes in some engine cycle parameters, such as overallpressure ratio (OPR), and TET. Cutting-edge technologyof engine components like methods of cooling, efficienciesof components, ducts pressure losses, and introduction ofdifferent overall thermodynamic cycle, for example,use of unconventional components like intercoolersand regenerators or recuperators [6,7]. More so, perfor-mance and economic viability of gas turbines are insepar-able [1].

Figure 1 Turboshaft gas turbine schematics [3].

Barinyima Nkoi et al.98

This paper considered turboshaft gas turbines in civilaviation. The investigation undertook a comparative assess-ment of simple and modified engine cycle options. Thecontribution of this work lies in the averment of thetechnical worth of modified engine cycles like low pressurecompressor (LPC) zero-staged, and intercooled/recuperated(ICR) cycles in helicopter turboshaft.

2. Theory and method

2.1. Design point (DP) performance

The design point of a gas turbine could be defined as thevery condition in the operating range of a gas turbine whenthe engine is running at the very mass flow, speed, andpressure ratio for which the components were designed [8].In establishing the design point of the engine, pressure ratioand TET that results in an overall highest thermal efficiencyare normally determined from preliminary cycle calculations.After this is done, other appropriate design parameters of thegas turbine system may be allotted. Then, detail design ofdifferent engine components can be done in order to providethe specified requirements of the complete system whenoperating at the DP. There are many requirements from a gasturbine engine. These may be referred to as design priorities,and always these requirements are in conflict. The design ofthe engine is greatly influenced by a set of these prioritiesdepending on the engine application [6,9].

2.2. Off-design (OD) performance

Besides the DP performance of the gas turbine, it ismandatory to ascertain its general performance over theentire operating range of power output and speed. This isknown as off-design (OD) performance [8]. Componentcharacteristics as indicated by component maps of com-pressor, turbine, and combustor, are very useful in ascer-taining off-design behaviour of the gas turbine system. Atsteady state operation of the engine, corresponding operat-ing points on the component maps are matched and can beplotted on the compressor characteristic diagram to form anequilibrium running line.

Various performance plots of power output, specific fuelconsumption (SFC), thrust, specific thrust or power, etccould be made once the operating conditions of an enginehave been determined. It is important to note that off-designperformance is very much affected by factors such asambient conditions of temperature and pressure, altitudes,flight speed (for aero engines), etc. The off-design perfor-mance analysis is normally achieved by the use of computermodel simulations of engines [10].

2.3. Turbomatch

Engine components operating point matching to establishOD performance is normally a tedious and time consumingtask since it is an iterative process. Computer basedsimulation is normally employed to accomplish the task.

Figure 2 Schematics of simple cycle two-spool turboshaft enginewith free power turbine.

Figure 3 T-S diagram of actual simple cycle of the two-spool enginewith free power turbine.

Performance assessment of simple and modified cycle turboshaft gas turbines 99

Turbomatch is an in-house gas turbine engine performancesoftware developed and established at Cranfield University.It is employed to simulate the DP and OD performances ofa broad range of aero and industrial gas turbines. Simplesingle shaft engines, complex multi-spool engines, as wellas novel cycle engine configurations can be modelledadequately using the scheme.

In the scheme, different engine components (intake,compressor, combustor, turbine, nozzle, etc) are representedby bricks (building blocks of the programme). These bricksare pre-programmed routines deployed to simulate, on amodular basis, the performance of the various enginecomponents they represent. The cycle thermal efficiency,specific fuel consumption, power, or thrust of the engine,etc are essential performance output parameters that areobtained as desired results of the simulation. Besides theseoverall cycle results, individual component performancecharacteristics, and the working-fluid properties at variousstations within the engine are also outputted [8,11,12].

2.4. Simple cycle two-spool turboshaft engine withfree power turbine

In this paper, two-spool turboshaft engine with a freepower turbine (FPT) was considered in which a lowpressure compressor (LPC) and a high pressure compressor(HPC) are driven by the high pressure turbine (HPT). Theschematic of this engine is shown in Figure 2.

The T-S diagram of the simple cycle is as shown inFigure 3 considering isentropic efficiencies of compressorsand turbines. With the notations of Figure 2 and Figure 3,and applying steady flow energy equation, heat flow intothe cycle in the combustion chamber (process 3–4) per unitair mass flow is given by Eq. (1).

qin ¼ h4−h3 ¼ cpðT4−T3Þ ð1Þ

Heat rejected at constant pressure (process 6-1) in theexhaust is given by Eq. (2).

qout ¼ h6−h1 ¼ cpðT6−T1Þ ð2ÞEquation (3) gives the total compressor work (CW)

(process 1-2-3) per unit air mass flow, where process 1-2occur in the LPC and process 2-3 occur in the HPC.

CW¼ LPCWþ HPCW¼ ðh2−h1Þ þ ðh3−h2Þ

CW¼ cp½ðT2−T1Þ þ ðT3−T2Þ� ð3Þ

High pressure turbine work (HPTW) (process 4–5) perunit air mass flow is defined by Eq. (4).

HPTW¼ h4−h5 ¼ cpðT4−T5Þ ð4Þ

Free power turbine work (FPTW) (process 5–6) given byEq. (5)

FPTW¼ h5−h6 ¼ cpðT5−T6Þ ð5Þ

This implies that total expansion work (EW) is obtainedas stated in Eq. (6)

EW¼HPTWþ FPTW

EW¼ cp ðT4−T5Þ þ ðT5−T6Þ½ � ð6Þ

The thermal efficiency is calculated using Eq. (7) below.

ηth ¼useful workheat input

¼ EW−CWcpðT4−T3Þ

ð7Þ

2.5. Modifications to the simple cycle

To increase the efficiencies of the simple-cycle, uncon-ventional components are added to the cycle. Thesecomponents include such like intercoolers, regenerators(recuperators), or reheaters. However, the initial and main-tenance costs of the cycle may increase due to theseadditional components. The improvements in cycle

Barinyima Nkoi et al.100

performance brought about by these components can onlybe justified if the decrease in fuel costs offsets the increasein other costs. There is the general urge to reducefuel consumption in gas turbine operation [13]. This isachieved by the introduction of these modifications to thesimple cycle. The descriptions of these modifications areoutlined below.

2.5.1. Recuperated cycleThe turboshaft engine in Section 2.4 with a recuperator is

represented on the T-S diagram as shown in Figure 4. Therecuperator or regenerator is a heat exchanger connectedbetween the turbine exhaust and the compressor exit. Thethermal efficiency of the cycle increases due to recuperationbecause the portion of heat in the exhaust gases that issupposedly wasted by flaring is now utilised to preheat theair at the exit to the compressor. This, in effect, reduces theheat gain from burning fuel, and hence, decreases fuelconsumption for same power output. However, if thecompressor outlet temperature is equal to or higher thanthe turbine exhaust temperature, the use of a regenerator isnot recommended. Else, there will be a reversal of heat flowto the exhaust gases, causing the efficiency to decrease.Very high pressure ratios in gas-turbine engines could causethis adverse situation.Referring to the cycle in Figure 4, T6 is the maximum

temperature that can occur within the recuperator which isthe temperature of the exhaust gases entering the recup-erator and leaving the turbine. Air in the regenerator(recuperator) can only be preheated to a temperature belowT6, and air will normally exit the regenerator at T7, a lowertemperature [14].The heat input per unit air mass flow here is given by Eq. (8).

qin ¼ h4−h7 ¼ cpðT4−T7Þ ð8Þ

Figure 4 T-S diagram of actual cycle with a regenerator.

Equation (9) calculates the thermal efficiency in this casemaking reference to Eq. (3) and Eq. (6) [14].

ηth ¼useful workheat input

¼ EW−CWcpðT4−T7Þ

ð9Þ

2.5.2. Intercooled/recuperated cycleIncorporating an intercooler between the LPC and HPC

of the recuperated engine in Section 2.5.1 such that airleaving the LPC is cooled before entering the HPC,resulting in an intercooled/recuperated cycle. Intercoolingreduces the total compressor work, thereby, increasinguseful work output, turbine work remaining the same [6].Also, intercooling will increase the specific work of thecycle, increase heat input from combustor, and thus fuelconsumption will rise [15]. However, this intercooler effectof reducing thermal efficiency is compensated by recupera-tion effect in ICR. The T-S diagram for the cycle with bothintercooling and recuperation is as shown in Figure 5,where process 2–3 is intercooling.

Using the station numbering and notations in the T-Sdiagram of Figure 5, intercooler effectiveness is given byEq. (10) below [16].

εinter ¼T2−T3T2−T1

ð10Þ

3. Simulation results

Performance analysis of a gas turbine engine starts withpreliminary design of the engine model. The nature ofapplication, the shaft power requirement, and of course, theclass of helicopter, are important factors to be consideredwhen making choice of engine model sizing.

Figure 5 T-S diagram of actual intercooled/recuperated cycle.

Performance assessment of simple and modified cycle turboshaft gas turbines 101

3.1. Choice of baseline engine core

Considering the envisaged application scenario for heli-copter on mission offshore oil/gas rig, the class of helicop-ter that would be involved, coupled with availability ofdata, the choice of engine core for the preliminary enginemodel design of this project was made. This core is inspiredby Makila 2A core, a Turbomeca turboshaft engine. Thisengine was specially developed to power 11-ton twins, likeEurocopter EC225 and EC725 [17]. The Makila 2A is atwo-spool turboshaft engine with a free power turbinewhich is capable of delivering 1567 kW (2101 shp) attake-off ISA sea level static (SLS). Its power output shaftspeed is 23,000 rad/min; air mass flow is 5.7 kg/s, andpressure ratio of 11:1 [17–19].

Table 1 Design parameters of the simple cycle engine.

Design parameter Value

Power rating (free power turbine) 1567 kWCompressor isentropic efficiencies 0.79Surge margin parameter 0.85Overall pressure ratio 11.25LP compressor pressure ratio 2.5HP compressor pressure ratio 4.5Turbine entry temperature 1500 KInlet air mass flow 5.7 kgCombustor efficiency 0.99HP turbine isentropic efficiency 0.88Power turbine isentropic efficiency 0.89

3.2. Simple cycle engine design point performancesimulation

It should be understood that the design point ofthe inspiring engine core is proprietary information of theoriginal equipment manufacturer (OEM), and as such, thedesign point is reasonably chosen by the engineeringjudgment. This is because some key defining parametersof the DP like the TET etc are not usually disclosed by theOEM [20]. Based on this fact, and due to availability ofdata, take-off condition at international standard atmo-sphere, sea level static was chosen as the design point withthe adoption of USAF standard for pressure recovery atintake. In order to carry out computation of the engineperformance, the engine components were modeled inturbomatch bricks as shown in Figure 6 following theschematics of Figure 2. The design parameters of the engineare presented in Table 1 below.

The turbine entry temperature and component efficiencieswere assumed. Both compressors are driven by the HPturbine. A 10% air mass flow bleed for cooling of HPturbine blades, and a combustion chamber pressure loss of5% of the HPC delivery pressure were allowed. The surge

Figure 6 Simple cycle two-spool turboshaft engin

margin parameter in Table 1 is the factor that limits thecompressor operating point from coinciding on the surgeline in order to prevent the compressor from surge. Inturbomatch code, surge margin parameter lies between 1.00and zero. Depending on pressure ratios, in turbomatch code,as surge margin tends to the value of 1.00, operating pointapproaches surge condition and as it tends to zero, operat-ing point approaches choke condition. The surge margin of0.85 allotted in this design is the default value in turbo-match and means that the operating point is kept literally ata safe distance from the surge line. Surge margin parameterZ is defined by Eq. (11).

Z ¼ ðπ−πchokeÞðπsurge−πchokeÞ

ð11Þ

where, π¼pressure ratio at design point, πsurge¼pressureratio at surge condition, and πchoke¼pressure ratio at chokecondition [11].

The simulation was run in turbomatch using fixedgeometry mode, and the output file gave the performanceresults as shown in Table 3. The quantified T-S diagram ofthe simple cycle analysis is presented in Figure 7, notingthat temperature variation is the paramount parameter.

e with free power turbine in turbomatch bricks.

Barinyima Nkoi et al.102

3.3. Recuperated turboshaft engine DP simulation

Maintaining the baseline engine core arrangement, arecuperator (a set of heat exchanger) was coupled betweenthe outlet of the free power turbine and HP compressor. TheTET, OPR, inlet mass flow, and component efficienciespresented in Table 1 above were maintained, while addi-tional design parameters for the recuperated engine are asshown in Table 2.

Table 2 Additional design parameters of the recuperated cycleengine.

Design parameter Value

Recuperator effectiveness 65%Cold side pressure loss 1%Hot side pressure loss 3%

Table 3 Summary of DP performance results of the turboshaft engine

Performance parameter Val

Sim

Power turbine rating/kW 156Inlet mass flow/(kg/s)Exhaust mass flow/(kg/s)Fuel flow/(kg/s)Exhaust gas temperature/K 90Overall pressure ratio 1Thermal efficiencyLPC power/kW 62HPC power/kW 153HPT power/kW 216Percentage variation in HPT power of advanced cycles over that of SC

Figure 7 T-S diagram of DP analysis of the simple cycle.

A bleed of about 7% of inlet air flow for the HP turbineinlet blades cooling, and a mass flow leakage of 0.02 kg forthe recuperator, were allowed. Maintaining take-off at ISASLS as design point, the engine was simulated and outputresults generated. The DP performance results as obtainedin the simulation output file are shown in Table 3 and theDP cycle analysis is indicated in the T-S diagram ofFigure 8.

3.4. DP simulation of engine with LP compressorzero-staged

A compression stage was added at the inlet of the LPcompressor of the simple cycle-baseline engine therebymodifying it to a different engine configuration. Thisadditional stage (zero-stage) increased the overall pressureratio by 1.13, and is intended to increase air mass flowthrough the engine.

s.

ue at DP of simulated turboshaft engines

ple cycle (SC) (baseline) Recuperated LPC zero-staged ICR

7 1567 1567 15675.7 5.7 5.7 5.75.83 5.81 5.83 5.800.130 0.107 0.127 0.1048 916 887 9311.25:1 11.25:1 12.71:1 11.25:10.281 0.339 0.287 0.3495 625 727 6258 1538 1606 11663 2163 2333 17910 0 7.86 −17.2

Figure 8 DP analysis of the recuperated cycle on T-S diagram.

Performance assessment of simple and modified cycle turboshaft gas turbines 103

The combination of all the compressors is still driven bythe HP turbine. A bleed of about 8% of inlet air mass flowwas permitted to cool the HP turbine inlet blades. With thecore arrangement of the baseline engine and data of Table 1maintained, this zero-staged configuration was modeled inturbomatch bricks, and simulated with take-off at ISA SLSas design point. The performance results are as indicated inTable 3 and the quantified T-S diagram is shown inFigure 9.

3.5. Intercooled/recuperated turboshaft engine DPsimulation

The recuperated engine cycle described in Section 3.3was modified by incorporating an intercooler between theLP and HP compressors. This is aimed at lowering thetemperature of the air mass entering the HP compressor inorder to reduce the work done by the HP compressor, andconsequently, the overall compression work is reduced, andas such, cycle efficiency will improve. The parameters ofthe recuperator cycle were maintained, except that a bleedof 10% of the inlet air mass flow was allowed for HPturbine inlet blade cooling. The air leaving the LP com-pressor was cooled to a temperature of 300 1C which resultsfrom an intercooler effectiveness of 88%, and a pressureloss of 4% of LP compressor delivery pressure.

Maintaining take-off at ISA SLS as design point, theintercooled/recuperated engine was simulated, and the DPperformance result is as indicated in Table 3. Its T-Sdiagram is presented in Figure 10.

The variation in compressor work due to zero-stagingLPC and intercooled/recuperation is indicated in the valuesof LPC power and HPC power in Table 3. The powerproduced by the HPT changes accordingly to the sum of thecompressor powers being the only turbine that drives all thecompressors. At DP, the FPT output was assumed fixed forall four cycles and so is not affected by the variation in

Figure 9 T-S diagram of DP analysis of the LPC zero-staged cycle.

compressor work because it is not driving any of thecompressors. The rational here was to compare perfor-mances of cycles with the same nominal power rating interms of thermal efficiency and fuel consumption. Recup-eration alone has no noticeable effect on compressor work.However, all the advanced configurations reduced the fuelflow which eventually reduced heat energy input incombustor and hence increased thermal efficiency.

3.6. Off-design performance of the engines

As described in Section 2.2, gas turbine engines normallydo not operate at the DP in practice due to the effects ofconditions such as changing ambient temperature, altitude,turbine entry temperature, among others. By the use ofcomponent maps in turbomatch codes, the off-designperformance of all the engine versions were simulated andthe variation of some key output parameters were as plottedin Figures 11–14. The discussion of these results ispresented in Section 4.

The compressor maps are the most important componentmaps when it pertains analysing the off-design perfor-mance of an engine [12]. Figures 15–18 show the scaled

Figure 11 Variation of SFC with ambient temperature at ISA SLS.

Figure 10 T-S diagram of DP analysis of ICR cycle.

Figure 14 Effect of ambient temperature on thermal efficiency.

Figure 15 LP compressor map of efficiency versus pressure ratio.

Figure 16 LP compressor map of pressure ratio against correctedmass flow.

Figure 17 HP compressor map of efficiency versus pressure ratio.

Figure 12 Variation of thermal efficiency with TET at ISA SLS.

Figure 13 Effect of altitude on thermal efficiency.

Figure 18 HP compressor map of corrected mass flow versuspressure ratio.

Barinyima Nkoi et al.104

compressor maps applied in turbomatch to implement theoff-design performances of the engine cycles, where CNrefers to non-dimensional speed.

4. Results discussion

From the simulation results shown in Table 3 and plots ofFigure 11 to Figure 14, it could be observed that therecuperated, LP compressor zero-staged, and the inter-cooled/recuperated engines all have increased thermalefficiency, compared to the simple cycle–baseline engine

Figure 19 Percentage changes in performance parameters of mod-ified cycles over simple cycle.

Performance assessment of simple and modified cycle turboshaft gas turbines 105

at both DP and off-design point. The percentage increasesin thermal efficiency of these unconventional cycle enginesover the traditional simple cycle engine especially at DP areshown in Figure 19. More so, specific fuel consumption isimproved in the modified cycle engines than thesimple cycle.

The negative sign on the SFC in Figure 19 indicatespercentage reduction in SFC of modified cycles over simplecycle–baseline engine. The advantage of improved perfor-mance of the recuperated cycle is as a result of reduction inthe heat flow rate required from fuel combustion. This isoccasioned by increased temperature of the air stream at theinlet to the combustor by the recuperator, thereby reducingfuel flow. This feature of the recuperator coupled withfurther reduction in compressor work occasioned by theincorporation of the intercooler makes the ICR cycleregister far better performance. The cycle with LP com-pressor zero-staged showed a better performance than thesimple cycle because of the increase in air mass flowthrough the engine by the additional stage of compressionin front of the LP compressor.

However, it is important to note that though the thermalefficiency is improved by using the modified cycles, theincorporation of intercoolers, recuperators, and additionalcompression stage, would render the engine more complex.This would raise the capital and maintenance cost actually,but cost of fuel would reduce due to reduction in fuelconsumption.

5. Conclusions

Appropriate software based on gas turbine theory wasapplied to carry out design simulation of a simple cycle(baseline) two-spool turboshaft engine and some modifiedcycle configurations. Such modified cycles include LPCzero-staged, recuperated, and ICR engine cycles. In doingso, design and off-design point performances of the engine

models were established. The results of performance para-meters for the simulated baseline engine compares favor-ably with those obtained from public domain sourcereference.

It is observed that to a large extent, the modified enginecycles with unconventional components exhibit betterperformances in terms of thermal efficiency and specificfuel consumption, than the traditional simple cycle engines.The percentage increases in thermal efficiencies of enginecycle with LP compressor zero-staged, recuperated enginecycle, and intercooled/recuperated engine cycle, over sim-ple cycle at DP are 2.1%, 20.6%, and 24.2% respectively,whereas percentage reduction in specific fuel consumptionin these cycles over simple cycle at DP are 2.6%, 17.3%,and 21.1% respectively.

These results compare favorably with values obtained inpublic domain literature. For instance, it was reported thatthe 1.4 MW intercooled/recuperated Heron-1 turboshaft gasturbine manufactured by EECT of the Netherland, exhibitsthermal efficiency of 42.9%, while a simple cycle gasturbine of same power range has thermal efficiency of about26%–34%. This represents a thermal efficiency increase ofabout 26.2% at the minimum [21]. More so, zero-stagingthe LP compressor of a helicopter turboshaft engine of theRTM322 core was reported to have yielded 5.21% increasein thermal efficiency over that of the baseline engine [22].

The foregoing performance results of the engine cyclescall for further verification by way of techno-economic andenvironmental risk analysis (TERA), which of course isunderway by our team at CU.

Acknowledgments

The authors would like to thank Dr. V. A. Pachidis of theDepartment of Power and Propulsion of Cranfield Univer-sity, for his valuable contribution.

References

[1] P.P. Walsh, P. Fletcher, Gas Turbine Performance, Seconded., Blackwell Publishing, Malden, 2004, p. 38.

[2] M. Bellis, Turboshaft engines, Different types of engines,URL: ⟨inventors.about.com/od/jstartinventions/ss/jet_engine_4.htm⟩ [cited 28 June 2012].

[3] Rotorcraft flying handbook, U.S. Department of Transporta-tion, Federal Aviation Administration, Chapter 5–2, 2000.

[4] D. Dilba, Pioneering new core engine technologies, MTUAero Engines, MTU Aero Engines Holding AG., URL:⟨www.mtu.de/en/takeoff/report/archive_old/2_2010/210_newac/index.html⟩ [cited 6 September 2012].

[5] M.D. Paramour, M.J. Sapsard, Future technology andrequirements for helicopter engines, in: W.A. Stewart (Ed.),NATO-AGARD Conference on Helicopter Propulsion Sys-tems, Toulouse, France, 11–14 May 1981, AGARD-CP-302.

[6] R.T.C. Harman, Gas Turbine Engineering: Application,Cycles and Characteristics, The MacMillan Press Ltd, London,1981, pp. 34–52.

Barinyima Nkoi et al.106

[7] R.K. Bhargava, M. Bianchi, S. Campanari, A. De Pascale,G. Negri Di Montenegro, A. Peretto, A parametric thermo-dynamic evaluation of high performance gas turbine basedpower cycles, Journal of Engineering for Gas Turbines andPower 132 (2) (2010).

[8] J. Kurzke, A program to calculate design and off-designperformance of gas turbines, GasTurb 8.0 User's Manual,URL: ⟨www.gasturb.de⟩ [cited 06 January 2012].

[9] Y.G. Li, P. Pilidis, M.A. Newby, An adaptation approach forgas turbine design-point performance simulation, Journal ofEngineering for Gas Turbines and Power 128 (4) (2006)789–795.

[10] Q.Z. Al-Hamdan, M.S.Y. Ebaid, Modeling and simulation ofa gas turbine engine for power generation, Journal ofEngineering for Gas Turbines and Power 128 (2) (2006)302–311.

[11] J. Palmer, The turbomatch scheme for aero/industrial gasturbine engine design point/off design performance calcula-tion, Cranfield University, UK, 1999 (unpublished).

[12] W.L. MacMillan, Development of a modular-type computerprogram for the calculation of gas turbine off-design perfor-mance, Ph.D. Dissertation, School of Engineering, CranfieldUniversity, Cranfield, Bedford, 1974.

[13] D. Lane, Brayton cycle: the ideal cycle for gas-turbineengines in relation to power plants, URL: ⟨web.me.unr.edu/me372/Spring2001/Brayton%20Cycle.pdf⟩ [cited 19 June2012].

[14] H. Cohen, G.F.C. Rogers, H.I.H. Saravanamuttoo, P.V.Straznicky, Gas Turbine Theory, Sixth ed., Pearson Educa-tion Limited, 2009, pp. 1–3.

[15] T.K. Ibrahim, M.M. Rahman, A.N. Abd Alla, Study on theeffective parameter of gas turbine model with intercooledcompression process, Scientific Research and Essays 5 (23)(2010) 3760–3770.

[16] P.A. Bromnick, R.J. Pearson, D.E. Winterbone, Intercoolermodel for unsteady flows in engine manifolds, in: Proceed-ings of the Institution of Mechanical Engineers, Part D:Journal of Automobile Engineering, 1998.

[17] Safran Turbomeca Makila, ⟨www.turbomeca.com⟩ [cited 05January 2012].

[18] Aero-engines: turboshaft engines, Jane's All the World'sAircraft, Jane's Information Group, London, 2011.

[19] Turbomeca Makila, Aero-Engines 29 (2011) 154–156.[20] E. Andreadis, Upgrade evaluation of a military turbofan

engine, Dissertation, Power and Propulsion Department,School of Engineering, Cranfield University, Cranfield, Bed-ford, 2009.

[21] R. Farmer, Intercooled recuperated 1400-kW genset has an8000-Btu heat rate, Gas Turbine World 32 (2) (2002).

[22] J. Vickers, Growth potential of a helicopter turboshaft,Dissertation, Power and Propulsion Department, School ofEngineering, Cranfield University, Cranfield, Bedford, 1995.