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ORIGINAL ARTICLE

Design and analysis of annular combustionchamber of a low bypass turbofan engine in a jettrainer aircraft

C. Priyant Marka,n, A. Selwynb

aDepartment of Aerospace Engineering, Karunya University, Coimbatore, T.N., IndiabA.E.R.D.C., Hindustan Aeronautics Ltd., Bangalore, KA, India

Received 19 May 2014; accepted 4 March 2016Available online 26 May 2016

KEYWORDS

Aerodynamic design;Annular combustionchamber;CFD (computationalfluid dynamics) analy-sis;Gas turbine engine;Optimization;Real-time model

tional Laboratory foe (http://creativecom

16/j.jppr.2016.04.00

or.

iyant.mark.c@gmail

esponsibility of Nat

Abstract The design of an annular combustion chamber in a gas turbine engine is thebackbone of this paper. It is specifically designed for a low bypass turbofan engine in a jettrainer aircraft. The combustion chamber is positioned in between the compressor and turbine.It has to be designed based on the constant pressure, enthalpy addition process. The presentmethodology deals with the computation of the initial design parameters from benchmarking ofreal-time industry standards and arriving at optimized values. It is then studied for feasibilityand finalized. Then the various dimensions of the combustor are calculated based on differentempirical formulas. The air mass flow is then distributed across the zones of the combustor.The cooling requirement is met using the cooling holes. Finally the variations of parameters atdifferent points are calculated. The whole combustion chamber is modeled using Siemens NX8.0, a modeling software and presented. The model is then analyzed using various parametersat various stages and levels to determine the optimized design. The aerodynamic flowcharacteristics is simulated numerically by means of ANSYS 14.5 software suite. The air-fuelmixture, combustion-turbulence, thermal and cooling analysis is carried out. The analysis isperformed at various scenarios and compared. The results are then presented in image outputsand graphs.& 2016 National Laboratory for Aeronautics and Astronautics. Production and hosting by Elsevier B.V.

This is an open access article under the CC BY-NC-ND license

(http://creativecommons.org/licenses/by-nc-nd/4.0/).

r Aeronautics and Astronautics. Produmons.org/licenses/by-nc-nd/4.0/).

1

.com (C. Priyant Mark).

ional Laboratory for Aeronautics and

ction and hosting by Elsevier B.V. This is an open access article under the

Astronautics, China.

C. Priyant Mark, A. Selwyn98

1. Introduction

Gas turbine engine evolved as a critical part and the mostefficient propulsion unit for aircrafts. It is now used inalmost all of the passenger aircrafts worldwide withdifferent variations. Military aircraft made the debut inusing the turbojet engine. As the technology progressed,high performance military aircraft began using low bypassturbofan engines due to its advanced capabilities, efficiencyand reliability, even at supersonic speeds.Low bypass turbofans have a bypass ratio of around 1:1

or less [1]. A high specific thrust/low bypass ratio turbofannormally has a multi-stage fan, developing a relatively highpressure ratio and, thus, yielding a high (mixed or cold)exhaust velocity. The core airflow needs to be large enoughto give sufficient core power to drive the fan. A smaller coreflow/higher bypass ratio cycle (for the fan operation) can beachieved by raising the high pressure (HP) turbine rotorinlet temperature. The temperature rise of the airflow fromthe intake to the nozzle of the engine is also less, whichresults in a reduced fuel flow leading to a better specific fuelconsumption (SFC) for the same pressure ratio. Thus, a lowbypass turbo fan would add to the efficiency of the engine.Jet fighters as well as trainers are high performance

aircraft that use the most powerful engines for producingthrust. The process of upgrading military hardware hasinitiated the race to develop even more powerful engines.By increasing power, the engines require more fuel input,thereby resulting in fuel guzzling engines. This directlypoints to an inefficient engine in terms of fuel consumption.Fuel consumption efficiency is required even in militaryaircraft as it can aid in increasing the range. For improvingefficiency, the very fundamentals lie in the combustionchamber. An efficient combustion chamber is the answerfor better performance.The most commonly used type of combustor is the fully

annular combustor, the others being tubular and tuboannu-lar combustor. Annular combustors [2] do away with theseparate combustion zones and simply have a continuousliner and casing in a ring (the annulus). There are manyadvantages to annular combustors, including more uniformcombustion, shorter size (therefore lighter), and less surfacearea. Additionally, annular combustors tend to have veryuniform exit temperatures. They also have the lowestpressure drop of the three designs (on the order of 5%).The annular design is also simpler, although testing gen-erally requires a full size test rig. Most modern engines useannular combustors; likewise, most combustor research anddevelopment focuses on improving this type. This paperdeals with designing an efficient annular combustionchamber for use in jet trainer aircrafts.Conrado [3] has discussed a design methodology which

follows a similar approach for designing a micro gas turbinecombustor. It also showcases an example and furtherautomating the same approach using a computer program

for ease of use. Silva [4] has discussed a consolidated designmethodology for an automotive turbocharger utilizing amicro gas turbine combustor. It gives a brief report alongwith heat-transfer analysis. Generally, the computationalfluid dynamics (CFD) analysis of a combustor is carriedout based on different combustion models [5]. Few modelssuch as Westbroor-Dryer one step model and Westbroor-Dryer two step model hold good for laminar combustionsimulation. Likewise K-epsilon model, K-omega model andK-omega shear stress transport (SST) model hold good forturbulent combustion simulation. The present paper discussesmainly about designing a gas turbine combustor at a scale ofa jet trainer aircraft engine using the most straightforwardand transparent approach. It also focuses on reducing thedevelopment time and gives ample support for refining thedesign at every phase. The paper also presents a computeraided design (CAD) model designed using the same princi-ples to show the practicality in using the design. For anaccurate CFD analysis result of a gas turbine combustionchamber, it needs to simulate combustion and turbulencesimultaneously. This paper gives a detailed CFD analysisreport of the designed combustor based on the combustion-turbulence interaction model.

2. Aerodynamic design

2.1. Preliminary design procedure

The procedure purposed by Melconian and Modak(1985) [6] to design a combustor is described in Figure 1.The equations utilized in the design procedure is presented,which is sufficient for the reader to understand the designmethodology idea.

2.2. Initial design parameters

The initial design parameters are mostly the compressorexit and turbine inlet constraints, which is usually absorbedfor any combustion chamber design. Others include custo-mer specifications, constants, experimental values andlimits. Table 1 shows the initial parameters used for thedesign, which were obtained from real-time data.

3. Dimensions

3.1. Casing area

Eq. (1) calculates the reference area [7].

Aref ¼ R

2_m3T0:5

3

P3

� �2 ΔP3�4

qref

ΔP3�4

P3

� ��1" #0:5

ð1Þ

Nomenclature

Dout outer diameter (unit: m)_m3 inlet air mass flow rate (unit: kg/s)_mDZ dilution zone air mass flow rate (unit: kg/s)_mDcool dome cooling air mass flow rate (unit: kg/s)_mPZ primary zone air mass flow rate (unit: kg/s)_mRZ recirculation zone air mass flow rate (unit: kg/s)_mSW swirler air mass flow rate (unit: kg/s)_mSZ secondary zone air mass flow rate (unit: kg/s)_man annulus air mass flow rate (unit: kg/s)_mcool cooling air mass flow rate (unit: kg/s)_mf fuel mass flow rate (unit: kg/s)ΔP3� 4P3

combustor pressure lossΔP3� 4qref

combustor pressure drop factor

ΔPLqref

liner pressure drop factor

ΔPSWqref

swirler pressure drop factor

ΔPdif

P3diffuser pressure loss

A3 inlet (compressor exit) area (unit: m2)AL liner area (unit: m2)AS snout area (unit: m2)ASW swirler flow area (unit: m2)Aan annulus area (unit: m2)Ao snout outer area (unit: m2)Aref reference area (unit: m2)Cds snout discharge coefficientD3 inlet diameter (unit: m)Dhub injector hub diameter (unit: m)DL liner diameter (unit: m)DSW swirler diameter (unit: m)Dint inner diameter (unit: m)Do snout outer diameter (unit: m)Dref reference diameter (unit: m)KSW swirler concordance factorLDZ dilution zone length (unit: m)LL liner length (unit: m)LPZ primary zone length (unit: m)LRZ recirculation zone length (unit: m)LSZ secondary zone length (unit: m)Ldif diffuser length (unit: m)Ldome dome length (unit: m)Mw molecular weight (unit: kg/mol)P3 inlet pressure (unit: Pa)R̂ arrhenius molar rate of creation/destruction

R3 inlet radius (unit: m)Ro snout outer radius (unit: m)T3 inlet temperature (unit: K)T4 exit temperature (unit: K)Tin zone inlet temperature (unit: K)Tmax max exit temperature (unit: K)Tout zone exit temperature (unit: K)kb backward rate constantkf forward rate constantnB number of swirler bladesqref reference dynamic pressure (unit: kg/(m � s2))ΔT temperature rise (unit: K)CAD computer aided designCFD computational fluid dynamicsCPF circumferential pattern factorFAR fuel air ratioHP high pressureR reactantRPF radial pattern factorSFC specific fuel consumptionA empirical constant, A¼4.0B empirical constant, B¼0.5C molar concentration (unit: (kg �mol)/m3)N number of reactionsP productPF pattern factorR universal gas constant (unit: (N �m)/(kg �K))R net rate of productionY mass fractioni; j speciesr reaction

Greek letters

βSW turning angle of the airflow (unit: 1)η0

rate exponent for reactantη00

rate exponent for productηcc combustor efficiency (unit: %)θRZ recirculation zone angle (unit: 1)κϵ large-eddy mixing time scaleν0

stoichiometric coefficient for reactantν00

stoichiometric coefficient for productΓ net effect of third bodies on the reaction rateρ density (unit: kg/m3)φ diffuser angle (unit: 1)

Design and analysis of annular combustion chamber of a low bypass turbofan engine in a jet trainer aircraft 99

Aref ¼ 0:08217 m2

3.2. Liner area

The combustor sectional area (AL) can be calculated byEq. (2) [6].

AL ¼ 0:66Aref ð2Þ

AL ¼ 0:05423 m2

3.3. Annulus area

The annulus area Aan, is the difference between Aref andAL and can be calculated from Eq. (3).

Aan ¼ Aref �AL ð3Þ

Figure 1 Preliminary design procedure.

Table 1 Initial design parameters.

Parameter Value Units

_m3 28.7103 kg/sT3 743.352 KP3 2083450 Pa_mf 0.25818 kg/s

Figure 2 Reference length Dref for annular configuration.

C. Priyant Mark, A. Selwyn100

Aan ¼ 0:02794 m2

3.4. Casing and liner diameter

Figure 2 presents the reference length Dref for annularcombustor configuration. The value of Dref is calculatedfrom Aref and DL is calculated from AL and it must bechosen such that it accommodates the aerodynamic con-siderations in every operating condition.

Dref ¼ 0:061 m

DL ¼ 0:04026 m

3.5. Pattern factor

The pattern factor or temperature traverse quality givesthe temperature distribution of the efflux gasses across theradial and circumferential direction at the exit of thecombustor. It is an important factor for the turbine inletblades. It also influences the liner length. It is defined inEq. (4).

Pattern factor¼ Tmax�T4

T4�T3ð4Þ

PF ¼ 0:25

3.6. Liner length

The liner length [7] provides the total length of the zones.It can be calculated from Eq. (5).

LL ¼�DL

0:05 ΔPLqref

ln 1�PFð Þ ð5Þ

LL ¼ 0:15719 m

3.7. Primary zone length

The length of the primary zone can be calculated fromEq. (6) [6].

LPZ ¼34DL ð6Þ

LPZ ¼ 0:03020 m

3.8. Secondary zone length

The length of the secondary zone can be calculated fromEq. (7) [6].

LSZ ¼ 12DL ð7Þ

LSZ ¼ 0:02013 m

3.9. Dilution zone length

The length of the dilution zone can be calculated fromEq. (8) [6].

LDZ ¼DL 3:83�11:83PF þ 13:4PF2� � ð8Þ

LDZ ¼ 0:06933 m

4. Air flow distribution

For conventional design [2], about half of primary zoneair mass flow rate would be admitted through the swirler

Table 2 Air mass flow rate distribution.

Air mass flow rates Symbol Percentage/%

Inlet _m3 100Recirculation zone/snout _mRZ 20Swirler _mSW 12Dome cooling _mDcool 8Annulus _man 80Primary zone _mPZ 20Secondary zone _mSZ 10Dilution zone _mDZ 10Cooling air _mcool 40

Figure 3 Airflow apportioning.

Table 3 Fuel/air ratio and equivalence ratio for the zones.

Parameter Recirculationzone

Primaryzone

Secondaryzone

Dilutionzone

Fuel/air ratio 0.0450 0.0150 0.0118 0.0090Equivalenceratio

0.6701 0.2234 0.1763 0.1340

Figure 4 Front end geometry of combustor.

Design and analysis of annular combustion chamber of a low bypass turbofan engine in a jet trainer aircraft 101

and as dome cooling. The mass flow rate _mRZ correspondsto the sum of the air admitted in primary zone through theswirler and the air admitted through dome cooling slots.The swirler mass flow rate ( _mSW ), for flame ignition andstability, should have an equivalence ratio above 1. The restof the airflow goes through the annulus ( _mAN). Then it isdistributed to the primary, secondary and dilution zones asper the requirement. For cooling, based on the formula [5]:Cooling air%¼ 0:1T3�30, 40% of the total mass flow rateis taken and distributed along the zones based on thetemperature. Table 2 and Figure 3 show the air mass flowrate distribution in the combustor.

The primary/recirculation zone equivalence ratio shouldnever be richer than 1.5 in order to minimize smoke, carbonmonoxide (CO) and unburned hydrocarbons (UHC) in theexhaust gases. To prevent nitrous oxides (NOx) and otherpollutants due to thermal dissociation, the equivalence ratiois capped at a maximum value of 0.6. The equivalence ratioof secondary zone shall not be higher than 0.8. The fuel/airratio and the equivalence ratio [8] for the respective zoneswas calculated and was found to be within the above limits.It is presented in Table 3.

5. Diffuser dimensions

The diffuser, swirler and the recirculation zone geometryis presented in Figure 4.

5.1. Snout outer area

The snout outer area Ao is calculated assuming the airvelocity in this sectional area is equal to Aan air velocity,then use Eq. (9) [3].

Ao ¼_m3

_manAan ð9Þ

Ao ¼ 0:03492 m2

5.2. Snout outer diameter

The snout outer diameter is obtained from Ao with thecalculations similar to liner diameter.

Do ¼ 0:02593 m

5.3. Diffuser angle

The diffuser angle φ can be obtained from Eq. (10) [7].

φ¼ tan �1ΔPdif

P3A3

2P32

502:4 1� A3Ao

� �2_m3

2T3

264

3751=1:22

ð10Þ

φ¼ 60:141

5.4. Diffuser length

The diffuser length can be obtained with the help ofEq. (11) [3], where Ro and R3 are Do/2 and D3/2,respectively.

Ldif ¼Ro�R3ð Þtan φ

ð11Þ

Table 4 The holes for the zones.

C. Priyant Mark, A. Selwyn102

Ldif ¼ 0:00170 m

Zones Main holes Cooling holes

No. ofholes

Holediameter/m

No. ofholes

Holediameter/m

Primary zone 40 0.01516 600 0.00395Secondaryzone

20 0.01502 480 0.00240

Dilution zone 20 0.01502 600 0.00329

6. Swirler dimensions

6.1. Snout area

The snout area can be calculated from Eq. (12) [6].

AS ¼ Ao_mRZ

_m3

1Cds

ð12Þ

AS ¼ 0:01075 m2

6.2. Snout diameter

The snout diameter is calculated from AS with thecalculations similar to liner diameter.

D ¼ 0:00798 m

o

θRZ ¼ cos �1 �DL DL�2DSWð Þ�ðDL�4LRZ ÞffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiDL

2�4DLDSW þ 4DSW2�8DLLRZ þ 16LRZ2

p2DL

2�4DLDSW þ 4DSW2�8DLLRZ þ 16LRZ2

" #ð16Þ

6.3. Swirler flow area

The swirler flow area can be calculated from Eq. (13) [3].

ASW ¼ffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi

Aref2

ΔPSWqref

KSW

_m3_mSW

� �2þ Aref

AL

� �2 cos 2 βSW

vuuut ð13Þ

ASW ¼ 0:00464 m2

6.4. Swirler diameter

The diameter of swirler is calculated using Eq. (14).

DSW ¼ffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiASW

nBþ π

4Dhub

2� �� �

4π

sð14Þ

DSW ¼ 0:02290 m

7. Recirculation zone dimensions

7.1. Recirculation zone length

The length of the recirculation zone approximates twoswirler diameters and can be obtained from Eq. (15) [3].

LRZ ¼ 2DSW ð15Þ

LRZ ¼ 0:04579 m

7.2. Recirculation zone angle

The recirculation zone angle can be obtained from Eq.(16) [3].

θRZ ¼ 14:41

7.3. Dome length

The dome length can be calculated from Eq. (17) [3].

Ldome ¼DL�DSW

2 tan θRZð17Þ

Ldome ¼ 0:03448 m

8. Holes

Firstly, it is necessary to verify if there is enough air toenter through every combustion chamber zone holes. Then,it is determined how much air enters through the primaryzone, the secondary zone and the dilution zone from thedifference between the total amount of air to enter eachzone and the amount of air that enters through other means(for example, through the swirler or slots). After thedetermination of the mass flow rate in each zone and themain hole type – plain and cooling hole type – ‘stackedring’, then it is possible obtain the hole area for each zone.This is an iterative process that follows a structuredsequence [3,9]. Table 4 shows the calculated holes.

Figure 5 Adiabatic temperature rise curves for kerosene (JP-5) fuel.

Figure 6 The gas temperature profile.

Figure 7 The whole combustor.

Figure 8 The combustor with the casing removed.

Figure 9 The analyzable model.

Table 5 CFD modeling data.

Parameter Value

Discretization Finite volume methodDomain Combustion-eddy dissipationMeshing method Octree/advancing frontTotal elements 2906742Total nodes 592157

Design and analysis of annular combustion chamber of a low bypass turbofan engine in a jet trainer aircraft 103

9. Gas temperature profile

9.1. The adiabatic flame temperature

The gas temperature profile is predicted theoreticallyusing numerical calculations to obtain design point values.The profile is more associated with the core temperature ofthe gas mixture due to the flame concentration at the corewhere the combustion is at the maximum. It is connectedwith the adiabatic flame temperature as this temperaturedetermines the core temperature of the gas mixture insidethe liner. This is the temperature that the flame would attainif the net energy liberated by the chemical reaction thatconverts the fresh mixture into combustion products werefully utilized in heating those products. In practice, heat islost from the flame by radiation and convection, so theadiabatic flame temperature is rarely achieved. It meansthat, the theoretical calculations give only an approximatevalue of the gas temperature. Nevertheless, it plays animportant role in the determination of combustion efficiencyand in heat-transfer calculations.

9.2. Calculation

For the average gas temperature calculations [3] insidethe liner at different zones, only the core air mass flow rateis considered. The core air mass flow excludes cooling airand other combusted byproducts.

The combustor is divided into four zones: recirculationzone, primary zone, secondary zone and dilution zone. Foreach zone, the local temperature is assumed to vary linearlybetween the zone inlet temperature (Tin) and zone outlet

Figure 10 The velocity vector in the model.

Figure 11 The velocity streamline in the model.

Figure 12 The total pressure contour in the model.

C. Priyant Mark, A. Selwyn104

temperature (Tout). For every zone, the outlet temperature iscalculated by Eq. (18).

Tout ¼ T3 þ ηccΔT ð18Þ

For recirculation zone, Tin is assumed to be equal to T3and the inlet temperature for every other zone is the outlettemperature of the preceding zone. ΔT is the temperaturerise from T3 to adiabatic flame temperature which is

Figure 13 The total temperature contour in the model.

Figure 14 The temperature contour at the outlet.

Design and analysis of annular combustion chamber of a low bypass turbofan engine in a jet trainer aircraft 105

calculated from the chart [7] on Figure 5 for the fuel JP-5using the corresponding FAR values of each zone.

9.3. Theoretical results

The calculated zone outlet temperatures [4] for each zoneis presented as graph in Figure 6.

10. Modeling

10.1. CAD model

The design was modeled using Siemens NX 8.0. Theviews are presented in Figures 7 and 8.

10.2. Analyzable model

To analyze the created model and obtain quicker results,due to computing limitations, the model was simplified intoa 201 cut section for a single burner. The view is presentedin Figure 9.

11. Aerodynamic analysis

The computational aerodynamic analysis is carried out tovalidate the theoretical results and to obtain a detailedpreview of the outcome of the design in real-time workingconditions. It was done using the commercially availableCFD code ANSYS 14.5 CFX to get a quick report of thecomputed data. The analysis was performed using thedesign parameters from Table 1 as inlet and turbine inletdata as outlet conditions. The initial setup data is given inTable 5. The eddy dissipation combustion model whichuses Eq. (19) and (20) [10], in combination with the finiterate chemistry model which uses Eq. (21) [10] was used inthe analysis, which allows accurate simulation of the heatrelease and the distribution of the main chemical species.This is a combustion-turbulence interaction model, whichsignificantly improves accuracy of analysis results.Figures 10–14 give the results.

Ri;r ¼ ν0i;rMw;iAρ

ϵ

κ

min

RYR

ν0R;rMw;R

!ð19Þ

Figure 15 The gas temperature comparison.

Figure 16 The radial and circumferential pattern factor.

Figure 17 The pressure loss.

C. Priyant Mark, A. Selwyn106

Ri;r ¼ ν0i;rMw;iABρ

ϵ

κ

ΣPYP

ΣNj ν

″j;rMw;j

ð20Þ

R̂i;r ¼ Γ ν00i;r�ν

0i;r

� �kf ;r ∏

N

j ¼ 1Cj;r

� �η0j;r �kb;r ∏N

j ¼ 1Cj;r

� �η00j;r !

ð21Þ

The comparison between the theoretically predicted gastemperature and the computed gas temperature is given inFigure 15.

The comparison [11] of radial and circumferential patternfactor is given in Figure 16.

The comparison of target and achieved pressure lossacross the combustor is given in Figure 17.

12. Results and discussion

The complete annular combustor design using just theinitial design parameters has been clearly discussed in thispaper. This is a more refined design methodology whichcan be used for the preliminary design. The transparent anddetailed approach is focused on reducing design time andcomplexity. This gives an overall advantage in total designtime and prototype building. Using the methodology, apractical design is presented. It follows the optimum valuesfor a preliminary setup. The obtained values are used formodeling and further simplified for analysis. The analysiswas also carried out with higher accuracy using thecombustion-turbulence interaction model and the resultsshow that the optimum (higher) gas exit temperature wasobtained for the present design. This gives a very promisingoutcome with respect to the exit temperature. The patternfactor was obtained as desired. The SFC was also reducedby regulating the temperature rise across the combustor.The required efficiency and pressure loss was achieved by athin margin. The designed combustor was found to beshorter than the other combustors of its class, which gives itan edge over others in the space constraint category. Basedon theoretical calculations and obtained results, the designpoint combustor exit temperature (enthalpy addition) wasachieved within 96% efficiency. Thus, the design is capableof reaching higher exit temperatures.

13. Conclusions

The design was successfully calculated and modeled. Therequired simpler model for analysis was also created. Thenthe model was aerodynamically analyzed at design pointand the geometry was optimized based on the results. Thishas delivered one of the most efficient combustion chamberdesign that can be used in the Jet Trainer Aircraft.

Acknowledgments

The authors are grateful to HAL, Bangalore for thecontribution of resources to complete the design andanalysis successfully.

References

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