Cowl and Cavity Effects on Mixing and Combustionin Scramjet Engines

Sang Hun Kang,∗ Yang Ji Lee,∗ and Soo Seok Yang†

Korea Aerospace Research Institute, Daejeon 305-333, Republic of Korea

and

Michael K. Smart‡ and Milinda V. Suraweera§

University of Queensland, Brisbane, Queensland 4072, Australia

DOI: 10.2514/1.48818

To investigate the supersonic combustion patterns in scramjet engines, a model scramjet engine was tested in the

T4 free-piston shock tunnel. The testmodel had a rectangular intake, which compressed the freestream flow through

a series of four shock waves upstream of the combustor entrance. A cavity flame holder was installed in the

supersonic combustor to improve ignition. The freestream test condition was fixed at Mach 7.6, at an altitude of

31 km. This experimental study investigated the effects of varying fuel equivalence ratios, the influence of the cavity

flame holder, and the effects of cowl shape. As a result, supersonic combustion was observed at equivalence ratios

between 0.11 and 0.18. Measurements indicated that the engine thermally choked at a fuel equivalence ratio of 0.40.

Furthermore, the cavity flame holder and theW-shaped cowl showed improved pressure distribution due to greater

reaction intensity. With the aid of numerical analysis, the cavity and theW-shaped cowl are shown to be effective in

fuel–air mixing.

Nomenclature�h = average heat transfer coefficient,_mf = fuel mass flow rate,Pf = measured pressure in plenum chamber,Po;f = final pressure in fuel reservoir,Po;i = initial pressure in fuel reservoir,�Pt�noz = nozzle supply pressure,P1 = freestream pressure,Rf = ideal gas constant of fuelTo;i = initial temperature in fuel reservoirtf = final timeti = initial timeVo = volume of fuel reservoir (1:66 � 10�3 m3)VAyH2 = variance of normalized fuel mass flux� = fuel calibration constant� = fuel equivalence ratio

I. Introduction

T HE scramjet engine is one of the most promising candidates forfuture transport systems because of its high specific impulse and

light weight [1]. Even though a combined-cycle rocket engine or gasturbine would be more practical for transportation [2,3], therealization of supersonic combustion is still a critical aspect ofhypersonic airbreathing engine development, and it has beenextensively studied at the component level [4–7].

Because of a strong ram effect, a scramjet engine does not requirean air compressor; therefore, its geometry is very simple. However,inflow perturbations due to the intake configuration, such as shock

waves and flow nonuniformity, are difficult to eliminate in theabsence of a compressor. Perturbations in the combustor inflow caninfluence the performance of fuel injectors and flame holders.Therefore, to develop a clear understanding of various phenomena ina scramjet engine, a free jet test setup is most suitable [8–10].

In this paper, we investigate the supersonic combustion char-acteristics of various component configurations using free jet tests ofthemodel scramjet engine in the T4 free-piston shock tunnel. Resultsof the numerical analysis are also presented to characterize the engineperformance.

A scramjet model was designed to be tested in the T4 free-pistonshock tunnel. The design flight Mach number was 7.6, and the flightaltitude was set to 31 km. A rectangular intake with a four-shock-wave system was employed for a high total pressure recovery androbust combustion. The intake ramp angles were determined usingLevenberg–Marqurdt’s optimization method and Korkegi’s criteria[11,12]. With the installation of the W-shaped cowl, intakestartability was also enhanced. In the combustor, a cavity was in-stalled for mixing enhancement and flame holding. The perfor-mance of the supersonic combustor was predicted by the Rayleighline theory [13] and the perfectly stirred reactor model [14,15]. Forthe design of the cavity flame holder, the flow residence time in thecavity was determined by the Davis and Bowersox relation [14].

II. Experimental Apparatus

A. T4 Shock Tunnel

T4 is a free-piston driven, reflected shock tunnel. It consists of anannular reservoir, a free piston, a compression tube, a shock tube, anozzle, a test section, and a dump tank. An unscoredBrightform steelprimary diaphragm of varying thickness separates the driver gas inthe compression tube from the test gas in the shock tube. A secondarymylar diaphragm (0.1 mm thick) separates the shock tube sectionfrom the test section. The specifications of the T4 shock tunnel arepresented in Table 1 [16,17].

An axisymmetric contoured nozzle, capable of producingMach 7.6 flows, is used in the present study. The nozzle exit diameteris 270mm, and its length is 1150mm.The nozzle consists of an initialconical section that can produce an expanded uniform source flowand a contoured section that can straighten the flow. Based on pitot-pressure survey results, the test core flow diameter is 200 mm at thenozzle exit plane and reduced to 140mm at a downstream distance of500 mm from the nozzle exit plane [18]. The test model is placed

Received 6 January 2010; accepted for publication 17 June 2011.Copyright ©2011 by theAmerican Institute ofAeronautics andAstronautics,Inc. All rights reserved. Copies of this paper may be made for personal orinternal use, on condition that the copier pay the $10.00 per-copy fee to theCopyright ClearanceCenter, Inc., 222RosewoodDrive,Danvers,MA01923;include the code 0748-4658/11 and $10.00 in correspondence with the CCC.

∗Senior Researcher, Aero Propulsion SystemDepartment.Member AIAA.†Head of Department, Aero Propulsion System Department.‡Professor, Centre for Hypersonics, Division of Mechanical Engineering.

Senior Member AIAA.§Postdoctoral Research Fellow, Centre for Hypersonics, Division of

Mechanical Engineering; currently Research Engineer, Gexcon. MemberAIAA.

JOURNAL OF PROPULSION AND POWER

Vol. 27, No. 6, November–December 2011

1169

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

within the core flow by being inserted 275 mm into the nozzle. Thecore flow diameter at the cowl capture plane, which is 270 mm awayfrom the nozzle exit, is 190 mm.

B. Model Installation

Figure 1 shows the configuration of the test model. The flowturning angles due to the intake ramps and cowls are 12, 15,�12, and�15�. As a result, the freestream Mach 7.6 flow is compressed andslowed down to aMach 2.0–2.3flow at the entrance of the supersoniccombustor. In the combustor, a 3-mm-deep and 9-mm-long cavity isplaced 75.5 mm downstream of the cowl’s leading edge. Gaseoushydrogen is injected at an angle of 45� to the local flow through a row

of four sonic injectors. The injector holes are 2 mm in diameter andare spaced laterally at 25 mm intervals. For the intake startability, theW-shaped cowl is used. By cutting out the cowl in a W shape, theinternal contraction ratio of the intake stays within the Kantrowitzlimit [19]. The details of this design are explained in Kang et al. [20].

Figure 2 shows the pressure transducers installed in the test modelwithin the test section. In this picture, the Mach 7.6 nozzle is tempo-rarily unequipped to provide a clearer view of the model. In the testmodel, static pressures are measured at 32 stations. KuliteTM andPCBTM piezoelectric pressure transducers are used to measure thepressure levels. Kulite XTEL-190M piezoelectric pressure trans-ducers have an excitation voltage of 10 V and pressure ranges of0–70, 0–170, and 0–700 kPa. High pressure levels, such as pitotpressures and plenum chamber pressures, are measured by PCBtype 111A26 piezoelectric pressure transducers. The transducers’sensing faces are thermally protected from the flow by 25 �mcellophane disks covering them. Using measurement uncertaintyanalysis, the total systematic uncertainties in the pressure measuredby the Kulite and PCB transducers were estimated to be �2:5 and�3:8%, respectively.

Gaseous hydrogen is injected into the scramjet engine combustorthrough a row of four holes. Figure 3 shows the layout and schematicof the fuel supply systemwithin the testmodel. Fuel is injected fromaroom-temperature reservoir through a fast-acting solenoid valve. Thereservoir is a coiled Ludwieg tube that keeps the temperature of thefuel constant at approximately 300 K during injection. The injectionflow is initiated at least 8 ms before the test flow arrival.

The fuel system was calibrated before testing, in order to deter-mine the mass flow rate of hydrogen as a function of the reservoirpressure. The calibration procedure for the shock tunnel fuel systemis described in Robinson et al. [17]. The instantaneous mass flow rateof the fuel is given by

_mf � 1�P���1�=2�o;i P���1�=2�f (1)

where � is the experimentally determined fuel calibration constant,which in turn is given by

Table 1 Specifications of T4 free-piston shock tunnel

Description Quantity

Piston mass 92 kgCompression tube 229 mm inside diameter � 26 m longShock tube 76 mm ID � 10 m longNozzles Machs 4, 6, 7, 7.6, 8, and 10Enthalpy range 2:5–15 MJ=kgPressure range 10–50 MPa

Fig. 1 Model scramjet engine configuration.

Fig. 2 Test model installation of the test section.

1170 KANG ETAL.

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

��RfTo;i

Vo�Po;f � Po;i�P���1�=2�o;i

Zf

i

P���1�=2�f (2)

Figure 4 shows a typical fuel mass flow rate change during testing.As demonstrated in the figure, the fuel mass flow rate remainsconstant within 3% of the maximum value during the test interval.

C. Test Conditions

The freestream conditions are fixed at a Mach number of 7.6 andan altitude of 31 km. A detailed list of the flow conditions of the

nominal freestream is shown in Table 2.With these parameters fixed,the effects of fuel equivalence ratios and varying component config-urations were investigated. The tests are summarized in Table 3.

III. Experimental Results

A. Characteristics of Nonreacting Flows

In Fig. 5, the normalized pressure distributions within the modelfor nonreacting flow cases are shown. Here, pressure is normalizedby the freestream value P1. When fuel is absent, pressure fluc-tuations are observed in the combustor due to shock and expansionwave reflections. However, the mean pressure level is largelyunchanged within the combustor. At the intake, there is a region ofincreased pressure near the deflection point at x� 420 mm. Thispressure rise is indicative of a separation bubble. During the designphase, intake compression angles were determined using Korkegi’scriteria for two-dimensional (2-D) flat-plate ramp cases [12].Nevertheless, according to Korkegi, if three-dimensional (3-D)effects from the sidewalls are strong enough, separation can occureven in stable regions [12]. Therefore, effects from the test modelsidewalls are believed to be significant in this case. However, the sizeof the separation bubble is too small to affect the flow along thesecond ramp and the combustor. Moreover, at the end of the secondramp, at x� 564 mm, the measured pressure level is the same as thetheoretical prediction. This indicates that the performance of theintake is acceptable, even if 3-D effects from the sidewalls aredistorting the inflow.

As is also evident from Fig. 5, the measured pressure levels in thecase of fuel injection into a nitrogen freestream are similar inmagnitude and trend to the case with no fuel, demonstrating that fuelinjection alone has little influence on the flow at an equivalence ratioof 0.12.

B. Characteristics of Reacting Flows

In Fig. 6, pressure distributions within the model for reacting flowcases are presented.When�� 0:11 and 0.18, themeasured pressurelevels of the fuel-into-air test start to rise above those for the fuel-into-nitrogen test at approximately 700 mm from the leading edge,indicating a combustion phenomenon. The increase in measuredpressure levels due to combustion is greater for �� 0:18, with amaximum normalized pressure level of P=P1 226; 785 mmdownstream of the leading edge. For the case in which �� 0:40,pressure levels start to rise upstream of the fuel injection point. Thepressure distribution suggests that the boundary layer has separatedand that there are subsonic regionswithin the combustion chamber. Ifthe combustion efficiency of the model is assumed to be 0.7,theoretical analysis based on the Rayleigh line flow [13] predicts thatthermal choking will occur when �> 0:374.

To further investigate choking, Fig. 7 shows time histories of themeasurements by the first pressure sensor (Cb1), located just

Fig. 3 Schematic of the fuel delivery system.

Time (mms)

Pre

ssur

e(M

Pa)

Mas

sf lo

wra

te(g

/s)

0 1 2 3 46

8

10

12

14

16

18

20

0

0.5

1

1.5

2

2.5

3(Pt)noz

Fuel flow

Test time

Fig. 4 Typical fuel mass flow rate change during testing.

Table 2 Detailed conditions of

nominal freestream flow

Parameter Value

Ps, MPa 11.8Ts, K 2595Hs, MJ=kg 3.0Ppitot, kPa 85.4Pe, kPa 1.1Te, K 243�e, kg=m

3 0.016Ue, m=s 2360Me 7.56� 1.39Unit Reynolds number 2:4 � 106

Table 3 Test run summary

Shot number Test gas Amount of fuel ing=s (equivalence ratio, �)

Cavity Cowl

1 9481 Air 0.0 (0.0) Y W cowl2 9486 Air 1.551 (0.11) Y W cowl3 9487 Air 2.464 (0.18) Y W cowl4 9483 Air 5.620 (0.40) Y W cowl5 9489 N2 1.603 (0.12) Y W cowl6 9493 Air 0.0 (0.0) N W cowl7 9494 Air 1.518 (0.11) N W cowl8 9508 Air 0.0 (0.0) Y Flat cowl9 9509 Air 1.523 (0.11) Y Flat cowl

Distance (mm)

P/P

∞

0 200 400 600 800 10000

50

100

150

200

Fuel off, AirFuel-N2 : φ=0.12Theoretical Prediction

Separation

Fig. 5 Normalized pressure distribution for nonreacting flows.

KANG ETAL. 1171

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

upstream of the fuel injection point in the combustor. The pressuredistribution for the case inwhich�� 0:40 shows severe fluctuationsthat are not seen at lower fuel equivalence ratios. This suggests theoccurrence of the inlet buzz phenomenon, in which the inlet startsand unstarts rapidly. This is evidenced by a considerable transientpressure rise emanating in the aft section of the combustor, movingupstream toward the combustor entrance during testing. The pressureincrease is immediately followed by a rapid reduction in pressurewithin the combustor, indicating flow spillage.

C. Effects of Component Variations

For the investigation of cavity and cowl shape effects, the modelconfiguration was changed, as shown in Fig. 8. The normalizedpressure distributions for different model configurations are shownin Fig. 9. It can be seen that the configuration with the cavity and theW-shaped cowl shows higher pressures than the configurationswithout them. Figure 10 compares the increases in pressure due tocombustion for various configurations. In the case involving fuelinjection and the cavity, the pressure levels start to surpass thosewithout fuel injection at the position of pressure sensor Cb4(x� 705 mm). However, in the case without the cavity, pressurelevels starts to rise at approximately 725 mm from the leading edge.Ben-Yaker and Hanson confirmed that the cavity generates flowoscillations that enhance mixing in the supersonic shear layer,creating a recirculation zone that acts as a continuous ignitionsource [4].

In the case with the flat cowl, the starting point of the pressure risedue to combustion is almost the same as that in the case with the W-shaped cowl. However, with the flat cowl, the slope of the pressurerise is smaller. Figure 10 shows that the case with the flat cowlexhibits a lower combustion pressure. As a result, the maximumnormalized pressure is 162, whereas it is 187 with the W-shaped

Distance (mm)

P/P

∞

0 200 400 600 800 10000

100

200

300

400

Fuel OffFuel-N2 : φ=0.12Fuel-Air : φ=0.11Fuel-Air : φ=0.18Fuel-Air : φ=0.40

Thermal Choking

SupersonicCombustion

Fuel Injection

Fig. 6 Normalized pressure distribution for reacting flows.

Time (µms)

Pre

ssur

e(k

Pa)

0 5000 10000 150000

100

200

300

400

500

600 Fuel-Air : φ=0.11Fuel-Air : φ=0.18Fuel-Air : φ=0.40

Fig. 7 Time history of the first pressure sensor measurement in the

combustor.

a) Cavity and W cowl

b) No cavity and W cowl

c) Cavity and flat cowlFig. 8 Variations of the test model configuration.

Distance (mm)

P/P

∞

0 200 400 600 800 10000

50

100

150

200Cavity and W cowl

No Cavity and W cowl

Cavity and Flat cowl

Fig. 9 Effects of component variations on the normalized pressure

distribution for reacting flows (�� 0:11).

1172 KANG ETAL.

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

cowl. Therefore, it can be established that the cavity and theW-shaped cowl have better combustion enhancing effects.

IV. Numerical Analysis

A. Numerical Method

Numerical simulations were carried out to analyze the detailedflow and combustion characteristics of the model. Computationswere performed using the commercial code FluentTM, along with thecoupled-implicit solver. For turbulent characteristics, the shear-stresstransport (SST) k-!model was used [21]. The SST model is knownfor its good prediction of mixing layers and jet flows [22,23].

The wall boundary conditions of scramjet engine combustors arehard to describe using numerical analysis. Computations by Mitaniand Kouchi [24], done with simple adiabatic or isothermal wallconditions, showed combustion characteristics that differed fromexperimental data. Complicated wall boundary conditions had to beused to describe the experimental conditions. Star et al. tried variousother wall boundary conditions [25]. In their study, it was concludedthat simple isothermal wall conditions are not suited for the accurateprediction of the wall pressure distribution. On the other hand, esti-mation of the heat flux into the wall agreed better with experimentaldata.

In the present study, a steady-state convective wall boundarycondition is used for the combustor wall. Because of the high-speed

flow in the combustor, the flowfield can be expected to reach a steadystate in the short test times of a shock tunnel [8]. For the convectivewall boundary condition, the local Nusselt number for the turbulentflow is calculated as follows [26]:

Nux � St Rex Pr� 0:0296Re4=5x Pr1=3; 0:6< Pr < 60 (3)

For the given Reynolds number, Prandtl number, and geometry,

the calculated average heat transfer coefficient �h is calculated to be1368:5 W=m2K.

For the numerical analysis of the combustor, 3-D calculations arecarried out. A nine-species 18-step detailed chemical kinetics modelis employed for the simulation of hydrogen combustion. Because offlow symmetry, only half the combustor along the symmetry plane issolved for. In the grid resolution study, results with 1,760,000 gridpoints, 980,000 grid points, and 520,000 grid points all predicted thesame pressure level. Thus, for reliable computational fluid dynamics(CFD) results, 980,000 grid points are chosen for the combustoranalysis. To provide accurate inflow conditions, the computationaldomain of the combustor includes a part of the second ramp of theintake. At the inlet of the computational domain, only two obliqueshockwaves from the leading edge and the second ramp influence theinflow. A complicated shock and expansion wave reflection region isincluded in the computational domain. By the 3-D calculation of theintake compression ramps, the boundary-layer thickness is con-firmed to be smaller than 0.3mmat the computational domain inlet ofthe combustor.

For the intake analysis, a 2-D calculation is performed due to thesimplicity of the geometry. Sixty-thousand grid points are distributedin the computational domain. Here, an adiabatic wall boundarycondition is used. The computational domains of the intake and thecombustor are shown in Fig. 11.

B. Results of Numerical Analysis

1. Comparison of Computational Fluid Dynamics Results

with Experimental Data

In Fig. 12, numerical results for unfueled and fueled cases, with airas the test gas, are comparedwith corresponding experimental resultsfor the setup with no cavity and theW-shaped cowl. As evident in thefigure, there is good overall agreement between the numerical andexperimental test results. The slight pressure differences near thecombustor exit might be due to accumulated error in the estimationsof the pressure and the boundary layer from the inlet to the exit. Inparticular, the growing boundary layer could affect the overallpressure level by changing theflowpath area. In the combustor, shockwaves generated by the cowl are continuously reflected by thecombustor walls and propagated downstream. In this process,complex shock wave/boundary-layer interactions occur along theflowpath. However, Knight et al. have pointed out that Reynolds-

Distance (mm)

P/P

∞

0 200 400 600 800 10000

50

100

150

200Fuel OffFuel-Air : φ=0.11

Sensor Cb4

a) Cavity and W cowl

Distance (mm)

Distance (mm)

P/P

∞

0 200 400 600 800 10000

50

100

150

200

Fuel OffFuel-Air : φ=0.11

Sensor Cb4

b) No cavity and W cowl

P/P

∞

0 200 400 600 800 10000

50

100

150

200

Fuel OffFuel-Air : φ=0.11

Sensor Cb4

c) Cavity and flat cowl Fig. 10 Normalized pressure distribution changes due to combustion

for different configurations.

Fig. 11 Computational domain.

KANG ETAL. 1173

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

averaged Navier–Stokes turbulence models are not accurate enoughto predict shock wave/boundary-layer interactions precisely [27].Nevertheless, overall flow patterns and combustion characteristicscan be observed with such turbulence models [22].

2. Unfueled Air Run

For the unfueled air run, the static temperature contour plots of thecombustor are shown in Fig. 13. As can be seen here, in the no-cavity

case, the high-temperature regions are confined to the local thinboundary layer at the combustor wall. However, when a cavity ispresent, a high static temperature region is formed inside the cavity.Since this high-temperature region is located 6mmdownstream fromthe fuel injector, the cavity can play the role of a flame holder.

In Fig. 14, static pressure contour plots of the combustor with anonreacting flow are shown. As seen in the figure, the static pressuresin the configuration with the W-shaped cowl show transversedirectional fluctuations. Such pressure distributions could causeenhancement of mixing and reactions.

Fig. 13 Static temperature contours of the combustor for nonreacting

flow cases for different configurations.

Fig. 14 Static pressure contours of the combustor for nonreacting flow

cases for different configurations.

Fig. 15 Static temperature contours of the combustor for reacting flow

cases with different configurations.

Distance (mm)

Pre

ssur

e(k

Pa)

0 200 400 600 800-50

0

50

100

150

200

250CFD, intake (2-D)CFD, combustor (3-D)Experiment

a) Unfueled run

Distance (mm)

Pre

ssur

e(k

Pa)

0 200 400 600 800-50

0

50

100

150

200

250CFD, intake (2-D)CFD, combustor (3-D)Experiment

b) Fueled run Fig. 12 Numerical analysis and experimental results comparison forthe configuration with no cavity, W cowl.

1174 KANG ETAL.

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

3. Fueled Air Run

For the fueled air run, the static temperature contour plots of thecombustor are shown in Fig. 15. In the configuration with the cavityand theW-shaped cowl, the overall temperaturewithin the combustoris higher than in configurations without these two elements. Further-more, a high static temperature region appears close to the fuelinjectors in this case. Therefore, the ignition length is demonstrablyshorter in the casewith the cavity and theW-shaped cowl. In Fig. 16,static pressure contour plots of the combustor with the reacting floware shown. Here, pressure fluctuations in the transverse direction arestill valid, even with fuel injection and the reacting flow. Suchpressure fluctuations can increase the transverse directional velocity

of the inflow and enhance the fuel–air mixing. Figure 17 showscontour slices of the Mach number in the combustor for differentconfigurations. In all these cases, there are no sonic lines across thecombustor, indicating that combustion occurs in the supersonicregime. Figure 18 shows the OH mass fraction contour slices in thecombustor. As depicted in the plots, the reactions are initiated justdownstream of the fuel injectors and, even with high axial flowvelocity, spread in the radial direction. Furthermore, the OH massfractions in the combustor with the cavity and the W-shaped cowlshow much higher values than those without them. The OH massfraction is at its highest value inside the cavity itself, indicating thatthe cavity has enhanced the reaction of hydrogen.

4. Numerical Experiment for Mixing

In the previous sections, numerical results indicated the existenceof a high-temperature region within the cavity, which can act as anigniter. Obviously, this enhances combustion and reduces theignition delay time. Another role for the cavity could be as a mixer ofthe fuel and the oxidizer. Mixing could also be enhanced by thetransverse directional pressure nonuniformity due to the W-shapedcowl. Experimental results notwithstanding, it is difficult to deter-mine how much mixing has been enhanced by the presence of thecavity and the W-shaped cowl. For this purpose, an additionalnumerical experiment was performed. Using the same inflow andfuel conditions as before, the cavity and W-shaped cowl effects onmixing were reproduced. To observe the mixing phenomenon alone,combustion is artificially turned off. Mixing is monitored by thevariance VAyH2 , defined as

VAyH2� 1

��u�2A

Z��uyH2

� �uyH2�2 dA

� 1

��u�2A

Z ��uyH2

� 1

A

Z�uyH2

dA

�2

dA (4)

Zero variance denotes a perfectly uniform mixture, and as thevariance value increases, the mixture becomes less uniform. InFig. 19, variance values for different configurations are displayed.When the first configuration (no cavity, flat cowl) is compared with

Fig. 16 Static pressure contours of the combustor for reacting flowcases with different configurations.

Fig. 17 Mach number contours of the combustor for reacting flow

cases with different configurations.

Fig. 18 OH mass fraction contours of the combustor for reacting flow

cases with different configurations.

KANG ETAL. 1175

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

the second (cavity,flat cowl), it can be seen that both variances start atthe samevalue and decrease.However, the casewith the cavity showsa more severe decrease in the vicinity of the cavity than the casewithout the cavity. Consequently, the cavity is shown to be effectivein fuel–air mixing. Comparing the first configuration (no cavity, flatcowl) with the third (no cavity, W-shaped cowl), VAyH2 of the third

configuration shows a lower value than the first from the verybeginning to the end of the combustor. The cowl is located upstreamof the fuel injection point. This allows the inflow, perturbed by theW-shaped cowl, to enhance the fuel–air mixing from the upstreamregion. Because of the combined effects of the cavity and the W-shaped cowl, the fourth configuration (cavity,W-shaped cowl) showsthe lowest variance value in the figure.

V. Conclusions

The effects of different cowl and cavity configurations on fuel–airmixing and combustion in scramjet engines were investigated usingexperimental and numerical approaches. The model scramjet enginewas tested with the T4 free-piston shock tunnel. The test modelconsisted of a four-shock-wave system intake, aW-shaped cowl, anda supersonic combustor with a cavity flame holder. The test wasconducted under Mach 7.6 conditions at a 31 km altitude.

The experimental results showed active combustion for low(�� 0:11) and middle (�� 0:18) equivalence ratio test cases.However, in cases of high fuel equivalence ratios, thermal chokingand inlet unstarts were observed. Furthermore, the presence of thecavity and the W-shaped cowl resulted in greater combustion-induced pressure increases.

Numerically, the cavity in the combustor was predicted to generatea hot static temperature region that acted as an ignition source,improving the mixing characteristics. With the W-shaped cowl, thestatic pressure showed transverse directional fluctuations andresulted in improved mixing. Via the combined effects of the cavityand the W-shaped cowl, earlier ignition and more active combustionwere observed.

On the whole, the cavity and the W-shaped cowl generated flowperturbations to enhance fuel–airmixing and combustion effectively.However, in real-world applications, flow perturbations result in anincreased drag in most cases [6,28]. Sometimes, the advantages ofenhanced combustion can be smaller than the drawbacks of increaseddrag. Therefore, additional research should be conducted before theimplementation of such devices in real systems.

References

[1] Fry, R. S., “A Century of Ramjet Propulsion Technology Evolution,”Journal of Propulsion and Power, Vol. 20, No. 1, 2004, pp. 27–58.doi:10.2514/1.9178

[2] Kanda, T., Tani, K., and Kudo, K., “Conceptual Study of a Rocket-Ramjet Combined-Cycle Engine for and Aerospace Plane,” Journal of

Propulsion and Power, Vol. 23, No. 2, 2007, pp. 301–309.doi:10.2514/1.22899

[3] Quinn, J. E., “ISTAR: Project Status and Ground Test Engine Design,”AIAA Paper 2003-5235, July 2003.

[4] Ben-Yaker, A., and Hanson, R. K., “Cavity Flameholders for Ignitionand Flame Stabilization in Scramjets: Review andExperimental Study,”AIAA Paper 1998-3122, July 1998.

[5] Micka, D. J., and Driscoll, J. F., “Combustion Characteristics of a Dual-Mode Scramjet Combustor with Cavity Flameholder,” Proceedings ofthe Combustion Institute, Vol. 32, 2009, pp. 2397–2404.doi:10.1016/j.proci.2008.06.192

[6] Doster, J. C., King, P. I., Gruber, M. R., Carter, C. D., Michael, D. R.,and Hsu, K. Y., “In-StreamHypermixer Fueling Pylons in Supersonic,”Journal of Propulsion and Power, Vol. 25, No. 4, 2009, pp. 885–901.doi:10.2514/1.40179

[7] Smart, M. K., “Experimental Testing of a Hypersonic Inlet withRectangular-to-Elliptical Shape Transition,” Journal of Propulsion andPower, Vol. 17, No. 2, 2001, pp. 276–283.doi:10.2514/2.5774

[8] Stalker, R. J., Paull, A., Mee, D. J., Morgan, R. G., and Jacobs, P. A.,“Scramjets and Shock Tunnels: TheQueensland Experience,”Progressin Aerospace Sciences, Vol. 41, 2005, pp. 471–513.doi:10.1016/j.paerosci.2005.08.002

[9] Kanda, T., Tani, K., Kobayashi, T., Saito, T., and Sunami, T., “Mach 8Testing of a Scramjet Engine with Ramp Compression,” Journal of

Propulsion and Power, Vol. 18, No. 2, 2002, pp. 417–423.doi:10.2514/2.5950

[10] Tomioka, S., Hiraiwa, T., Kobayashi, K., and Izumikawa, M.,“Vitiation Effects on Scramjet Engine Performance in Mach 6 FlightConditions,” Journal of Propulsion and Power, Vol. 23, No. 4, 2007,pp. 789–796.doi:10.2514/1.28149

[11] Ozisik, M. N., Inverse Heat Transfer, 1st ed., Taylor and Francis,New York, 2000, pp. 37–58.

[12] Korkegi, R. H., “Comparison of Shock-Induced Two- and Three-Dimensional Incipient Turbulent Separation,” AIAA Journal, Vol. 13,1975, pp. 534–535.doi:10.2514/3.49750

[13] John, J. E., Gas Dynamics, 2nd ed., Allyn and Bacon, Boston, MA,1984, pp. 204–217.

[14] Davis, D. L., and Bowersox, R.W., “Stirred Reactor Analysis of CavityFlame Holders for Scramjets,” AIAA Paper 1997-3274, July 1997.

[15] Davis, D. L., and Bowersox, R. W., “Computational Fluid DynamicsAnalysis of Cavity Flame Holders for Scramjets,” AIAA Paper 1997-3270, July 1997.

[16] Portwood, T. W., “Enhancement of Hydrocarbon SupersonicCombustion by Radical Farming and Oxygen Enrichment,” M.S.Thesis, Univ. of Queensland, Brisbane, QLD, Australia, 2006.

[17] Robinson, M. J., Rowan, S. A., and Odam, J., “T4 Free Piston ShockTunnel Operator’s Manual,” Univ. of Queensland, Department ofMechanical Engineering Rept. 2003-1, Brisbane, QLD, Australia,2003.

[18] Suraweera,M. V., “Mach 7.6 Nozzle Pitot Survey in T4 Shock Tunnel,”Univ. of Queensland, Dept. of Mechanical Engineering Rept. 2006-1,Brisbane, QLD, Australia, 2006.

[19] Kantrowitz, A., and Donaldson, C., “Preliminary Investigation ofSupersonic Diffusers,” NACAWRL-713, 1945.

[20] Kang, S. H., Lee, Y. J., and Yang, S. S., “Design of Model ScramjetEngine for Shock Tunnel Test,” 9th Asian International Conference onFluid Machinery, KFMA Paper 9-106, Seoul, Oct. 2007.

[21] Menter, F. R., “Two-Equation Eddy-Viscosity Turbulence Models forEngineering Applications,” AIAA Journal, Vol. 32, No. 8, 1994,pp. 1598–1605.doi:10.2514/3.12149

[22] Choi, J. Y., Ma, F., and Yang, V., “Combustion Oscillations in aScramjet Engine Combustor with Transverse Fuel Injection,”Proceedings of theCombustion Institute, Vol. 30, 2005, pp. 2851–2858.doi:10.1016/j.proci.2004.08.250

[23] Bardina, J. E., Huang, P. G., and Coakly, T. J., “Turbulence ModelingValidation,” AIAA Paper 1997-2121, June 1997.

[24] Mitani, T., and Kouchi, T., “Flame Structures and CombustionEfficiency Computed for a Mach 6 Scramjet Engine,” Combustion andFlame, Vol. 142, No. 3, 2005, pp. 187–196.doi:10.1016/j.combustflame.2004.10.004

[25] Star, J. B., Edwards, J. R., Smart, M. K., and Baurle, R. A., “NumericalSimulation of Scramjet Combustion in a Shock Tunnel,” AIAAPaper 2005-0428, Jan. 2005.

[26] Incropera, F. P., and DeWitt, D. P., Fundamentals of Heat and Mass

Transfer, 4th ed., Wiley, New York, 1996, p. 355.

Distance (mm)

VA

yH2

x10

4

650 675 700 725 750-1

0

1

2

3

4

5

6

7

8

No Cavity, Flat Cowl

Cavity, Flat Cowl

Nocavity, W Cowl

Cavity, W Cowl

FuelInjection

Cavity

Fig. 19 Variance of hydrogen mixing for different configurations.

1176 KANG ETAL.

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18

[27] Knight, D., Yan, H., Panaras, A., and Zheltovodov, A., “Advances inCFD Prediction of Shock Wave Turbulent Boundary LayerInteractions,” Progress in Aerospace Sciences, Vol. 39, 2003,pp. 121–184.doi:10.1016/S0376-0421(02)00069-6

[28] Sunami, T., Itoh, K., Komuro, T., and Sato, K., “Effects of Streamwise

Vortices on Scramjet Combustion at Mach 8-15 Flight Enthalpies—AnExperimental Study in HIEST,” 17th International Symposium on AirBreathing Engines, AIAA Paper 2005-1028, Sept. 2005.

J. SeitzmanAssociate Editor

KANG ETAL. 1177

Dow

nloa

ded

by I

ND

IAN

IN

STIT

UT

E O

F T

EC

HN

OL

OG

Y -

KA

NPU

R o

n M

ay 2

6, 2

014

| http

://ar

c.ai

aa.o

rg |

DO

I: 1

0.25

14/1

.488

18