dual inlet ducted ramjet combustor

Tiré à Part

Experimental and numerical study of the turbulent flowinside a dual inlet research ducted rocket combustor

A. Ristori, G. Heid, A. Cochet, G. Lavergne

XIV Symposium ISOABEFlorence (Italy), September 05-10, 1999

TP 2000-14

Ce Tiré à part fait référence au Document d’Accompagnement de Publication DEFA0004

Experimental and numerical study of the turbulent flowinside a dual inlet research ducted rocket combustor

Étude expérimentale et numérique de l'écoulement turbulent

dans une chambre de combustion de Recherche du Statofusée à deux entrées d'air

par

A. Ristori, G. Heid, A. Cochet, G. Lavergne

XIV Symposium ISOABEFlorence (Italy), September 05-10, 1999

RésuméRésumé : Un programme de recherche sur les statoréacteurs a été initié à l'Onera en 1995 avecle soutien de la DGA/SPNuc dans le but d 'améliorer la méthodologie de conception et de mise aupoint des foyers de statoréacteurs par l'utilisation de codes CFD validés.Cet article présente les deux modèles de Statoréacteurs de Recherche conçus spécialement poursimuler expérimentalement les moteurs de type statofusée : le premier est dédié aux écoulementsnon réactifs et à pression atmosphérique, et le second aux expérimentations en écoulement réactifdans des conditions plus réalistes. Les visualisations hydrauliques et les résultats des mesures devitesse obtenues par anémométrie Laser 2D sur le modèle transparent sont présentés. Les résultatsdes simulations numériques de l'écoulement, menées à l'aide du code Navier-Sokes 3D turbulent etréactif développé à l' Onera et dénommé MSD, sont également présentés.Les résultats des simulations dans les cas en non réactif sont comparés aux mesures LDV ; unaccord plutôt satisfaisant entre les mesures et les calculs est obtenu. De plus, les résultats dessimulations numériques menées dans des conditions plus réalistes (cas d'écoulements réactifs) sontprésentés et analysés.

1American Institute of Aeronautics and Astronautics

EXPERIMENTAL AND NUMERICAL STUDY OF TURBULENT FLOW INSIDE

A DUAL INLET RESEARCH DUCTED ROCKET COMBUSTOR

A. Ristori , G. Heid, A. Cochet, G. Lavergne

Office National d’Etudes et de Recherches AérospatialesPalaiseau, France

Abstract

A research ramjet program has beeninitiated at ONERA with the support ofDGA/SPNuc in 1995 with the aim to improvemethodology for ramjet combustion chamberdesign and tuning by using validated CFD codes.

This paper presents the two dual inletramjet research models specifically designed tosimulate experimentally solid-propellant ductedrocket (SDR) motors: the first one is dedicated tocold flow experiments at atmospheric pressure andthe second one to hot flow experiments under morerealistic conditions. Hydraulic visualizations andaerodynamic 2D LDV results obtained on thetransparent model are presented. Computations ofthe flow with the 3D reacting turbulent Navier-Stokes code developed at ONERA and namedMSD are also presented.

The calculated results in non reacting flowcases are compared with LDV data; a rather goodagreement between measurements and calculationsare obtained. Moreover, calculated results in morerealistic conditions (reacting flow cases) arepresented and analyzed.

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Introduction

A three-dimensional ramjet combustorgeometry has been defined (combustion model) inorder to have operating conditions of the combustor(pressure, velocity, temperature) comparable to realmotors ; this combustor (Fig. 1) could be usedeither as a Solid Ducted Rocket (SDR) or LiquidFueled Ramjet (LFRJ) motor.

Validation studies of numerical tools usedfor SDR or LFRJ motors 1-6 have been made alwaysin non reacting flow cases. This suggests that anexperimental study of both SDR and LFRJpropulsion system in non reacting and reacting flowcases would be worthwhile to validate numericaltools in a large range of ramjet operation.

The first part of the program is focused oncold flow studies: visualizations (water flow) of themixing inside the SDR combustor (a specifictransparent model at scale 1.6 with respect to thecombustion model has been built) as a function offuel-air-ratio and LDV measurements (air flow)inside this model have been carried out.____________Copyright 1999 The American Institute of Aeronautics andAstronautics Inc. All rights reserved

The second part of the program will befocused on Combustion studies: visualizations,LDV, LIF and gas analysis measurements will bemade in the SDR combustor (combustion model)under realistic conditions.

Today, cold flow studies have started(limited to cold flow experiments with low airvelocities). Moreover, first numerical resultsobtained with a one-step fast chemical reactionscheme are presented for the SDR propulsionsystem (combustion model) ; those results needs tobe confirmed by experiments.

This paper presents the two ramjetresearch models specifically designed to simulateexperimentally SDR motors :

- the first one is dedicated to cold flowexperiments at atmospheric pressure (transparentmodel in plexiglass ). A cold flow hydraulic test righas been used to perform flow visualization of themixing process between air and gas generatorproducts inside the 3D test combustor. Theflowfields inside this dual inlet combustor werestudied experimentally using colorimetrictechniques and image processing for flowvisualizations of the mixing process. Then, a coldflow aerodynamic test rig has been used to perform2D LDV measurements inside the test combustor.The experimental data obtained from this study arehelpful in understanding the mixing processes andflow structures inside such a combustor, and alsofor partial validation of the numerical code MSD.Cold flow experimental results obtained with thetransparent model (visualizations of mixing, 2D-LDV measurements) and comparisons withnumerical results are presented.

PROBE

WINDOWS

IGNITER

Fig. 1: 3-D view of the combustion SDR model.


- the second one is dedicated to hot flowexperiments under more realistic conditions(combustion model equipped with quartz windows).At this time, numerical results under realisticconditions on this model have been performed.

Next steps of this study will beaerodynamic and combustion testing with localcharacterization of the flowfields (LDV,visualizations, LIF, gas analysis) on the SDRcombustor model. The transparent and combustionset-ups will also be used for the Liquid FueledRamjet (LFRJ) study; in this case, the aim will bethe two phase flow characterization for thevalidation of two phase flow models in CFD codes.

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Research Ducted Rocket Combustor

The SDR propulsion system consideredhere consists of a main combustor with two lateralair inlets and a fuel injection into the head end ofthe combustor, as shown in Fig. 1.

In real ducted rocket, the fuel is a solidpropellant located into an auxiliary combustorwhich is used as a gas generator; the partiallyburned products are ejected through holes into thehead end of the main combustor and continue toburn with the incoming air. In the ducted rocketdesigned for experiments, gas generator productsare replaced by air for cold flow experiments andpropane for hot flow experiments.

The combustion ramjet design (Fig. 2) isrepresentative of a real engine with the followingcharacteristics :

* The dome plate is flat and dome heightis adjustable from 30 mm to 100 mm (basicconfiguration of dome height = 50 mm) ;

* Area ratio A4 / A2 = 2 ;* Area ratio Ainj / A4 = 0.09 for semi-propellant products as fuel ;* Mach number M4 = 0.35 ;* Mach number M6 = 1.55.

Dimensions of the ramjet combustionmodel have been determined by using those enginecharacteristics and test rig performances. Testconditions for different flight case are given intable I for stoechiometric air/fuel mixtures:

Table IFLIGHT

CONDITION

Tair

(K)m°air

(kg/s)

m°C3H8

( g/s )

Pduct

( bar )

Tduct th

( K )

Low altitude 520 2,9 186 7,1 2418

Middle altitude 600 1,9 122 4,7 2440

High altitude 750 0,9 58 2,2 2472

In order to facilitate optical access insidethe combustor, the geometry has been designedwith a square section for the air inlets and the duct.

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Visualizations in the transparent SDR model

This model built in plexiglass is onlyavailable for atmospheric tests with water or airflow. For this reason, there is no nozzle at the outletand the length of the model has been limitedcompared to the combustion model. Fig. 3 show apicture of the transparent SDR model installed onthe hydraulic test rig. The scale factor between thetransparent model and the combustion model shownin is equal to 1.6 (all dimensions are multiplied bythis factor in order to facilitate measurements on theplexiglass model). Dimensions are for air inlets80x80 mm2, for duct 160x160 mm2 and for fuelholes φ=17.6 mm for hydraulic tests andφ=23.6.mm for aerodynamic tests (to keepmomentum quantity or differential injectionpressure when propane is replaced by air foraerodynamic tests).

* 2 Air inlets : section 50x50 mm2

* Distance between air inlet axis : 250 mm* Air injection angles : θ = 45°* Dome height : 50 mm* 2 Fuel inlet holes : φ = 11 mm* Distance between fuel holes : 50 mm* Duct: section 100x100 mm2

* Nozzle: section 55.8x100 mm2

Fig. 2 : Schematic of the ramjet combustion model.


For these tests, a water flow simulate airflow and a tracer dye (fluoresceine) mixed withwater simulate fuel flow. With the objective tovisualize fuel and air mixing processes, a laser lightsheet can move across side and lengthwise sectionsof the combustor. A CCD video camera (SonyXC75) is used to record images on a S-VHSvideocassette recorder ; the intensity delivered byeach pixel of the camera is directly proportional tothe local concentration of fuel.

Experiments were conducted for differentfuel injection conditions. Visualizations of themixing process, for the lengthwise section ZE = 40mm, show (Fig. 4) that the increase of theequivalence ratio improve the fuel jet penetrationinto the ram-air streams. However, if the fuel jetmomentum becomes too high ( ϕ = 1.7 andϕ.=.2.3.), a large fraction of the fuel passes throughthe ram-air stream without mixing with the air ; as aconsequence, we could observe a decrease of theefficiency or the blow-out of the combustor.

Visualization pictures of Fig. 5 are givenfor an equivalence ratio of 0.5. The fuel jetpenetration in the rich dome region is shown for thetwo lengthwise following sections : the first one islocated in the axis of the duct at ZE = 0 mm, and,the second one in the axis of one injector atZE.=.40 mm.

Air fuel mixing process is shown for foursections from the head end of the combustor todownstream air inlets respectively atXE.=.15.mm.(Section 1) ; 48 mm .(Section 2) ;71.mm .(Section 3) ; 134 mm .(Section 4) ; 188 mm(Section 5); 210 mm .(Section 6) ; 261 mm(Section 7) and 337 mm (Section 8).

Sideways to air inlets, the formation offour contra-rotative secondary flows createsconditions for fuel and air mixing in the chamber.The intensity level of those recirculation structuresis very important to create an efficient mixing. Themore rich region is quickly located near the axis ofthe duct (see section 8 of Fig. 5); this is favorable toincrease the flame surface and consequently thecombustion efficiency.

Fig.4: Influence of equivalence ratio on fuel jet penetration in the SDR model.

Fig.3: Transparent SDR model installed on thehydraulic test rig.

ϕ = 1.7

ϕ = 2.3

ϕ = 0.5

ϕ = 1.1

MIN MAX


Fig.5: Visualization pictures for ϕ = 0.5.

Numerical results obtained for low altitudeflight conditions in reacting flow with anequivalence ratio of 0.5 are qualitatively inagreement with hydraulic experiments concerningmixture ratio distribution in the duct (Fig. 6). Thehigher mixture ratio zone is situated more close tothe center of the duct when we approach to theexhaust plane; however, hydraulic visualizationsshow a faster mixing in the duct than computations(Fig. 7) probably because higher residence time inwater.

Fig. 6: Experimental (left) and numerical (right)mixture ratio distribution at Section 7.

Fig. 7: Experimental (top) and numerical(bottom) mixture ratio distribution at ZE=0mm.

1 2 3 4 5 7 86X (mm)

Section 8

Section 5

MIN MAX

1 2 3 4 5 7 8X (mm)

6

Section 2

Section 1

Section 3

Section 4

Section 6

Section 7

MIN

Numerical results (Reacting Flow)

Hydraulic visualization

MIN MAX


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LDV Measurements and Calculationsin the SDR transparent model

LDV measurements in aerodynamic test facility

In order to measure detailed gas-phasevelocity and turbulent fluctuations profiles duringcold flow tests, a four-beam, two-color (blue andgreen) Laser Doppler Velocimeter is used (Fig.8).A fiberoptic probe is used to collect laser beamsand create a probe volume, the forward light is thencollected by a multimode fiberoptic probe with twophotomultipliers (one for each component of thevelocity) at the end.

Flow solver of Navier-Stokes equations

The flow solver named MSD anddeveloped at ONERA 8 is an approximate Riemannsolver based on a 3-D finite volume technique usingstructured meshes. The numerical conservativeupwind implicit scheme features the shock-capturing Roe’s method. The fluxes decompositionis based on the flux difference splitting method.Favre-average conservation equations of the floware continuity equation, continuity speciesequations, momentum equations and energyequation. In order to close those equations, theReynolds stresses are formulated through

turbulence modeling; the classical K-epsilon modelis used in this study.

In order to complete the formulation, themean reaction rate ω.

α must be determined. A fastchemistry modelization is assumed to describe thediffusion flames encounted in the SDR combustor.The reaction is considered as a spontaneouslymixing of the fuel and oxidizer in a homogeneousreactor. The one-step scheme considered forpropane-air combustion gives the stoechiometriccoefficient and the Shvab-Zeldovich variable of thereaction. The fuel reaction rate ω.

F calculation isbased on the CRAMER 9 model.

Results and Discussions

Due to the presence of two symmetricalplanes [xy] and [xz] in its geometry (Fig. 9), only aquarter of the combustor has been taken intoaccount for computations. The mesh (a quarter ofthe combustor) is composed of 38625 nodes in theduct and 20250 nodes in the air inlet. Thecombustor operates at atmospheric pressure andtemperature (293 K) ; the mean air mass flow is66.4 kg/m2/s and inlet air boundary conditions aswell as fuel boundary conditions were imposed inorder to be in agreement with experiments. Themean fuel flow is 41.1 kg/m2/s, the resulting Fuel-Air-Ratio is equal to 4.23 %.

Fig.9: Mesh of transparent SDR model.

The schematic ofFlow pattern observedin the SDR combustoris given on Fig. 10.The recirculation zoneat the dome region iscaused by the jetimpingement of the tworam air streams and theexistence of freevolume at the head endof the combustor. Dueto the plane symmetry,there are two counterrotating vortices in thedome region.

Fig. 8: Laser Doppler Velocimeter system

X

Y

Z

XE (mm) = 250 390 645

YE (mm) = 80

YE (mm) = 0

YE (mm) = - 80

X

Y

Z

ZE (mm) = 0 40 80


Fig.10: Schematic of Flow pattern calculated in the SDR combustor.

In cross sections of ram-air outlets, fourcontra-rotating vortices are created (bottom part ofFig. 10) ; those vortices provide an efficientmechanism for mixing the ram air with fuel gases asshown previously on hydraulic visualizationspictures.

In order to analyze precisely the flowpattern in the SDR combustor, the LDVmeasurements were conducted on two differentlengthwise sections and three different crosssections in the chamber. Concerning lengthwiseplanes, one is the mid plane located at ZE=0 mmand the other is parallel to the mid plane andlocated at ZE=30 mm (Fig. 9) ; 585 positions wereselected for measurements on each half plane.Concerning cross section planes, there are locateddownstream ram-air outlets at XE=250 ; 390 and645 mm (Fig. 9) ; 224 positions were selected formeasurements on each half plane.

The measured velocity data at the differentcross and lengthwise sections defined is comparedwith the computational results.

The comparison between data andcalculated velocities in air inlets shows a goodestimation of profiles and acceleration of fluid flowin the inner bend. 2D LDV measured velocity andturbulence profiles for ram-air inlets and fuel inletshave been entered as boundary conditions in theCFD code for calculations in this configuration.

Fig.11: Velocity field measured and calculated for three sections (XE=250 mm to 645 mm).

Fig. 12: Comparison of measured and calculatedaxial velocity and fluctuations distributions.

Fuel injector

ZE = 40 mm

ZE = 0 mm

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A general good agreement between experimentaldata (on left parts of Fig. 11) and calculatedvelocities (on right parts of Fig. 11) are obtained inthe duct for the sections XE = 250 mm ; 390 mmand 645.m.

The measured turbulence intensitydistribution ((u’2)½ , <u’v’>) are shown in thebottom part of Fig. 11 at the section XE = 250 mm.We can notice that the higher turbulence intensityregions are corresponding to the higher meanvelocity gradient regions (20 mm< |YE| < 40 mm)

Fig. 12 shows a comparison of themeasured and calculated axial velocity andfluctuations distributions for ZE=0mm andZE=30mm . Fig. 13 and Fig. 14 shows moreprecisely the good agreement in experimental andnumerical profiles shapes, respectively for thelengthwise section ZE = 40 mm and ZE = 0 mm,even if axial velocity U and fluctuations (u’2)½

levels calculated are almost underestimated.

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Fig. 14: Calculated and measured profiles(lengthwise Section ZE = 0 mm)

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Fig. 13: Calculated and measured profiles(lengthwise Section ZE = 40 mm)

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LDV Experiments


Mean velocity is about 60 m/s at theoutlets of ram-air and about 30.m/s in the duct. Inreal case (combustion model), mean ram-airvelocity is 150 m/s and axial duct velocity is250.m/s. Thus, those cold flow aerodynamicmeasurements could be used only for a partialvalidation of the CFD code. In fact, compressibilityeffects are not taken into account in this case with aMach Number less than 0.1 in the duct against 0.35in real case.

Measured and calculated velocity andturbulent distributions are mostly in agreement.

Next steps will be the use of ASMturbulence model with probably an enhancement ofpredictions as a result.

Results obtained partially validates, forthis kind of applications, the theoretical andnumerical models developed in the MSD code.

Next steps will be the use of ASMturbulence model with probably an enhancement ofpredictions as a result.

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Calculations in the SDR combustion model inNon Reacting and Reacting Flow Cases

Calculations of the non-reacting andreacting flows in the SDR combustion model havebeen made with a mesh (based on a quarter of thecombustor) composed of 48260 nodes in the ductand 9800 nodes in the air inlet (Fig. 15)

In reacting flow case, instationnary resultsare obtained with a fluctuation of the mass flow rateat the exhaust of about 10 % at a frequency equal to430 Hz, which is characteristic of the longitudinalmode of the ramjet (Fig. 16).

By meaning values on several periods, weobtain the mean combustion efficiency which is

equal to 0.98 at the exhaust plane of the combustorin this case.

Fig. 16: Mass flow rate fluctuation in the SDRcombustor for reacting flow case.

Fig. 17 represents the fuel tracer massfraction and static temperature fields at differentcross sections along the duct. We can notice thatfuel tracer concentration is higher at the center ofthe duct ; as shown previously on hydraulicvisualizations (Fig. 5). The consequence of thisresult is that the temperature is also higher at thecenter of the duct.

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Summary and conclusions

In this study, we have analyzed the threedimensional flowfields and fuel air mixingdistributions inside a Research Ducted RocketCombustor. Hydraulic air fuel mixing visualizationsand Aerodynamic LDV measurements have beenexperimented on a transparent model of thecombustor in order to validate qualitatively andquantitatively the 3D, turbulent, reacting CFD codedeveloped at ONERA and named MSD. Numericalresults obtained on the transparent SDR model arequantitatively in agreement with LDV experiments ;other numerical results obtained on the combustionSDR model in realistic conditions are qualitativelyin agreement with hydraulic visualizations.

Next steps of this study will beaerodynamic and combustion measurements (LDV,visualizations, LIF, gas analysis) on the combustionmodel.

The transparent and combustion modelswill also be used for Liquid Fueled Ramjet (LFRJ)studies. Then, the aim will be the two phase flowcharacterization for the validation of two phaseflow models in MSD code.

Fig. 15: Mesh of the SDR combustor model

0.000 0.005 0.010 0.015 0.020 0.025TIME (s)

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E (k

g/s)

EXIT MASS FLOW RATE


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Acknowledgments

We would like to thank DGA/SPNuc forits support to this research program.

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References

[1] C.L. Chuang, D.L. Cherng, W.H. Hsieh,G.S. Settles, and K.K. Kuo, "Study of FlowfieldStructure in a Simulated Solid-Propellant DuctedRocket Motor" , 27th Aerospace Sciences Meeting,AIAA Paper-89-0011, 1989[2] Y.C. Chao, W.F. Chou, S.S. Liu,"Computation of Turbulent Reacting Flow in aSolid-Propellant Ducted Rocket", Journal ofPropulsion and Power, vol. 11, N° 3, May- June1995.[3] T.M. Liou, Y.H. Hwang, and Y.H. Hung,"Computational Study of Flow Field in Side InletRamjet Combustors", AIAA/ASME/SAE/ASEE24th Joint Propulsion Conference, AIAA Paper-88-3010, 1988.[4] T.M. Liou and Y.H. Hwang, "Calculation ofFlowfields in Side-Inlet Ramjet Combustors with anAlgebraic Reynolds Stress Model", Journal ofPropulsion, vol. 5, N° 6, Nov.-Dec. 1989.[5] P.K. Wu, M.H. Chen, T.H. Chen,"Flowfields in A Side-Inlet Ducted Ramrocketwith/without Swirler",31st AIAA/ASME/SAE/ASEE Joint PropulsionConference and Exhibit, AIAA Paper-95-2478,1995. [6] F.D. Stull, R.R. Craig, G.D. Streby, and S.P.Vanka, "Investigation of a Dual Inlet Side DumpCombustor Using Liquid Fuel Injection", Journal ofPropulsion, vol. 1, N° 1, Jan.-Feb. 1985.[7] AGARD Advisory Report 323 - WorkingGroup 22 on "Experimental and Analytic Methodsfor the Determination of Connected-Pipe Ramjetand Ducted Rocket Internal Performance", AGARDAR-323 , 1994. [8] D. Dutoya, M.P. Errera, P.J. Michard, A.Ristori, "Présentation d’un code de calculd’écoulements compressibles 3-D dans des canauxet des cavités de forme complexe", AGARDConference Proceedings CP-510 , 1992 pp 29-1 ,29-21 . Paper presented at the Propulsion andEnergetics Panel 77th Symposium held in SanAntonio (USA) 27-31 May 1991.[9] F. Dupoirieux, "Calcul numériqued’écoulements turbulents réactifs et comparaisonavec des résultats expérimentaux" , La RechercheAérospatiale n° 1986-6, Nov-Dec, pp 443-453.

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Nomenclature

A section [m2]m. mass flow rate [kg/s]M Mach NumberP pressure (Pa)Re Reynolds NumberT temperature [K]V velocity [m/s]XE distance [m]YE distance [m]ZE distance [m]Y mass fraction

αs Stochiometric coefficient of the reactionϕ equivalence ratioρ density [Kg/m3]ω. rate of production of species (s-1)

Subscriptsair air inlet stationC3H8 fuel inlet stationduct end of duct station (before nozzle)th theoreticalF fuel species

dual inlet ducted ramjet combustor

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