Yolanda R. Hicks, Christopher M. Heath, Robert C. Anderson, and Kathleen M. Tacina

Glenn Research Center, Cleveland, Ohio

Investigations of a Combustor Using a 9-Point 

Swirl-Venturi Fuel Injector: Recent 

Experimental Results

NASA/TM—2012-217245

April 2012

ISABE–2011–1106
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Yolanda R. Hicks, Christopher M. Heath, Robert C. Anderson, and Kathleen M. Tacina

Glenn Research Center, Cleveland, Ohio

Investigations of a Combustor Using a 9-Point 

Swirl-Venturi Fuel Injector: Recent 

Experimental Results

NASA/TM—2012-217245

April 2012

ISABE–2011–1106

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

20th International Symposium on Air Breathing Engines (ISABE 2011)

sponsored by the International Society for Air Breathing Engines

Göteborg, Sweden, September 12–16, 2011
Acknowledgments

This work was supported by the Subsonic Fixed Wing and Supersonics Projects in NASA’s Aeronautics Research Mission 

Directorate. The authors gratefully acknowledge Changlie Wey for his assistance and guidance. We also thank the CE-5 test crew, 

especially our technicians Jeff Hamman, Bill Rozman, and Harold Redloske.

Available from

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Trade names and trademarks are used in this report for identification 

only. Their usage does not constitute an official endorsement, 

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Space Administration.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217245 1

Investigations of a Combustor Using a 9-Point Swirl-Venturi 

Fuel Injector: Recent Experimental Results

Yolanda R. Hicks, Christopher M. Heath, Robert C. Anderson, and Kathleen M. Tacina

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

This paper explores recent results obtained during testing in an optically-accessible, JP8-fueled, flame 

tube combustor using baseline Lean Direct Injection (LDI) research hardware. The baseline LDI 

geometry has nine fuel/air mixers arranged in a 3 × 3 array. Results from this nine-element array include 

images of fuel and OH speciation via Planar Laser-Induced Fluorescence (PLIF), which describe fuel 

spray pattern and reaction zones. Preliminary combustion temperatures derived from Stokes/Anti-Stokes 

Spontaneous Raman Spectroscopy are also presented. Other results using chemiluminescence from major 

combustion radicals such as CH* and C2

* serve to identify the primary reaction zone, while OH PLIF 

shows the extent of reaction further downstream. Air and fuel velocities and fuel drop size results are also 

reported. 

Nomenclature

Symbols

Af, m2 combustor cross-sectional area

Ae, m2 effective mixer exit area

C2

*

, diatomic carbon radical

CH* CH radical

D32, μm Sauter mean diameter

f, mm focal length

L/D length-to-diameter ratio

OH* hydroxyl radical

NOx oxides of nitrogen

P, kPa pressure

P3, kPa inlet combustor pressure

T, K temperature

T3, K inlet combustor temperature

Tad, K adiabatic flame temperature

U, m/s axial velocity component

V, m/s transverse velocity component

W, m/s vertical velocity component

X axial flow direction

Y transverse flow direction

Z vertical flow direction

ΔP/P combustor pressure drop

Ф equivalence ratio

Abbreviations

AS Anti-Stokes

CFD Computational Fluid Dynamics 
NASA/TM—2012-217245 2

FWHM Full Width Half Maximum

(I)CCD (Intensified) Charge-Coupled Device

LDI Lean Direct Injection

LDV Laser Doppler Velocimetry

NASA National Aeronautics and Space Administration

Nd:YAG Neodymium: Yttrium Aluminum Garnet

PLIF Planar Laser-Induced Fluorescence

PDI Phase Doppler Interferometry

PIV Particle Image Velocimetry

RMS Root Mean Square

S Stokes

SRS Spontaneous Raman Scattering

UHCs Unburned Hydrocarbons

UV Ultraviolet

Introduction

The International Civil Aviation Organization estimates that aviation accounts for nearly two percent 

of the anthropogenic-induced carbon dioxide emissions global wide and contributes 3 percent to the 

potential warming effect (Ref. 1). Those numbers are expected to increase an additional 1 to 2 percent by 

2050 (Ref. 1). Near the primary flight corridors of the globe, the effects from aviation can become more 

severe. At ground level, NOx increase the production of ozone, leading to smog. Estimates are that 

commercial aircraft operations generate 0.4 percent of all national low-altitude NOx emissions (below 

913-m) (Ref. 2). From a survey released in 2005, the Federal Aviation Administration (FAA) estimates 

airports contribute between 0.7 to 6.1 percent to the NOx inventory of the surrounding locality (Ref. 3). In 

the troposphere, NOx have been found to increase ozone production. At stratospheric altitudes for 

supersonic cruise, NOx have been shown to play a reverse effect and decrease the protective ozone layer

(Ref. 4).

To reduce the environmental effects of aviation and ensure that aircraft gas turbines meet future more 

stringent noise and emissions regulations, NASA is developing engine technologies that will decrease 

emissions while simultaneously increasing fuel efficiency. However, many technologies that increase 

engine efficiency also increase temperature and pressure at the inlet of the combustor, T3 and P3; for a 

given combustor technology, increasing T3 and P3 often increase NOx emissions. NASA has investigated 

several combustion concepts to reduce NOx emissions. This paper describes optical measurements from 

one low NOx combustor concept: lean direct injection.

As the name implies, in LDI the combustor operates fuel-lean without a rich front end: all of the 

combustor air except that used for liner cooling enters through the combustor dome (Ref. 5). Like other 

lean burn combustor concepts, LDI reduces NOx by minimizing flame temperature; in fuel lean 

combustion, NOx is an exponential function of temperature. To eliminate local “hot spots” that produce 

high levels of NOx, lean burn combustion concepts rely on the fuel and air being well-mixed before 

burning occurs. Thus, LDI requires rapid fuel vaporization and fuel-air mixing. LDI achieves this by 

using a multi-element concept in which several small fuel/air mixers replace a single conventional fuel/air 

mixer.

Improved understanding of LDI operation will permit guideline development for future LDI 

combustor refinement. In order to achieve this understanding, examination of the physical and chemical 

processes that occur is required. The required information can be provided either by computational fluid 

dynamics (CFD) or by detailed optical measurements. In any case, detailed optical measurements are 

required because CFD results also require validation.

Optical measurements that have been applied to flows in simulated gas turbine combustors include 

Planar Laser-Induced Fluorescence (PLIF), Planar Elastic Scattering, Spontaneous Raman Spectroscopy 

(SRS), Phase Doppler Interferometry (PDI), and Particle Image Velocimetry (PIV) (Refs. 6 to 10). 
NASA/TM—2012-217245 3

Results of those measurements include combustion species maps, velocities, fuel/air ratios, and turbine 

profile factors. These measurements are used to study fuel injection, mixedness, and combustion 

processes and are part of a database of measurements that will be used for validating computational 

combustion models (Refs. 11 to 15).

This paper presents additional measurements for a nine-element LDI array. Measurements include 

high speed imaging, flame spectra, velocity, and turbulence. The data were collected over a range of inlet 

conditions that simulate the conditions of next-generation aircraft combustors. Results that compare 

fueling all nine elements with fueling only the center element are presented here for the first time.

Experimental Setup

The LDI hardware was tested in a high temperature and high pressure flame tube combustor rig with 

partial optical access directly downstream of the fuel injection site. The experimental facility and 

instrumentation offered the ability to perform non-intrusive optical and laser-based diagnostics, to 

examine fluid and chemical processes during combustion. The test cell was capable of supplying nonvitiated inlet air to the combustor at temperatures between 505 and 866 K, pressures between 1034- and

1724-kPa with flow rates up to 4.54 kg/s. 

An isolated full-section view of the window assembly and downstream flame tube are displayed in 

Figure 1. The combustor housing is 74-cm in length and lined with an aluminum oxide ceramic to 

withstand average temperatures reaching 2033 K. UV-grade fused silica windows (38.1- by 50.8-mm) are 

mounted just downstream of the fuel injection site and provide optical access from three sides of the 

square combustion chamber. The combustion chamber cross-section measures 76.2- by 76.2-mm. The 

bottom of the chamber is reserved for a spark igniter and provides no physical or optical access. The 

grade of quartz used in the windows offers a light transmission of at least 80 percent within a spectral 

range from 180- to 3300-nm. Nitrogen film cooling is used internally to regulate window temperature and 

accounts for no more than 10 percent by mass of the total inlet air flow.

Figure 1.—Illustration of flame tube and window housings.
NASA/TM—2012-217245 4

The baseline 9-point LDI test hardware is described in Tacina et al. (Ref. 5) and consists of a 3 × 3 

injector array measuring 76.2- by 76.2-mm with injectors spaced 25.4-mm (center-to-center). Each 

injector element is comprised of a set of six helical, axial vanes angled at 60° that surround a simplex 

nozzle designed to produce a hollow cone fuel spray. The nozzle atomizes and injects liquid fuel (Jet-A, 

JP-8) at the throat of a converging-diverging venturi. The calculated swirl number for this injector 

architecture is 1.0. A schematic of a single injector element is displayed in the cut-away view of Figure 2. 

An end-view of the full 9-point array is shown in Figure 3.

The optical instrumentation was fixed to a set of stages to provide traversing ability along up to three 

orthogonal directions. In this work X- is defined as the axial stream-wise direction, with X = 0-mm being 

the exit (dump) plane of the 9-point LDI hardware. Y- is the transverse horizontal direction and Z- is the 

vertical direction, with (Y = 0-mm, Z = 0-mm) at the center of the center injector. Positive directions 

follow from the right-hand-rule, with positive X downstream and Z up.

Figure 2.—Schematic drawing that shows the 

relative spatial positioning of the air swirler, fuel 

nozzle, and venturi for each injector element.

Figure 3.—End view of the 9-point LDI test hardware.
NASA/TM—2012-217245 5

Diagnostic Setup

Imaging Techniques—PLIF, Luminescence, PIV

The experimental setups that fall under the category of imaging techniques were similar in that any 

light scattered or emitted was collected, using an imaging array (camera), from an angle perpendicular to 

the combustor flow direction, as illustrated in Figure 4. Various filters were applied so that only the 

wavelength bands of interest were allowed to pass to the detector. For PIV and PLIF, the two 2–D

techniques that use lasers, the laser beam was formed into a light sheet using a set of cylindrical lenses. 

The sheet was passed vertically through the test section and aligned parallel with the flow direction. For 

these methods, the laser sheet and collection optics were traversed together so as to maintain focus on the 

laser sheet. For chemiluminescence measurements, either species-specific, or high speed imaging, the 

camera was positioned so that the vertical center plane of the test rig was in focus. 

OH and fuel PLIF images, along with chemiluminescence images of C2, and CH, were obtained using 

the same receiving optics and camera. Collection optics included the appropriate filter (FWHM of 10-nm, 

typical) attached to a UV-grade, f = 105-mm, f/4.5, macro camera lens. The light was collected using a 

gated, 16-bit, 1k × 1k pixel array, intensified CCD (ICCD) camera having a Gen II Super-Blue-SlowGate intensifier. The gate time used was 50-ns. For PLIF, OH and fuel molecules were excited using a 

10-Hz, frequency-doubled Nd:YAG-pumped dye laser/frequency mixer system to achieve wavelengths 

around 282-nm. The laser was formed into a sheet using a pair of cylindrical lenses, to obtain a sheet 

approximately 300 μm thick. For PLIF, we used on-chip averages of 200 gates, and traversed across the 

flow along the Y axis in 1-mm increments. For chemiluminescence, we averaged 600 gates on-chip. The 

filters used to isolate the species were 313-nm (OH), 334-nm (fuel), 430-nm (CH), and 515-nm (C2).

Figure 4.—Illustration that shows a typical 2–D or 1–D imaging arrangement such as for PLIF or PIV.
NASA/TM—2012-217245 6

Flame imaging was achieved using a 12-bit, high speed CMOS camera (1k × 1k) focused on the 

vertical center plane. Light emitted from the volume and within the field-of-view of the collection optics 

was collected. In the work presented here, we used a frame rate of 10,000 frames per second, image 

exposure time of 40-μs, and image resolution of 576 × 944 pixels. The flame chemiluminescence was 

collected using an f = 150-mm, f/1.2 lens. The camera array was sensitive to visible light; the light 

collected was largely a combination of CH and C2 emissions.

The PIV data were acquired using a 15-Hz double-head, frequency-doubled, Nd:YAG laser and a 

single, interline transfer, CCD camera. For the un-fueled, air-only tests, 0.7-μm diameter alumina

particles were seeded into the air inlet plenum 10-L/D before the plenum exit. The inlet plenum feeds the 

flow to the flame tube test section. For combusting experiments, seed was not used. Instead, we collected 

the light scattered by fuel particles to measure the liquid phase velocity. We used a mechanical leaf 

shutter to block light from the luminescent flame.

Imaging Techniques—Raman Scattering

As with the 2–D imaging above, 1–D Spontaneous (Vibrational) Raman Scattering (SRS) was 

collected 90° from the incoming light and square to the test rig. We used a 30-Hz, frequency-doubled, 

Nd:YAG laser focused using a spherical lens into the test section. The beam waist at the focus was 

approximately 2-mm. Collection of the Raman-scattered light was achieved using an f = 60-mm, f/2.8, 

macro lens to 0.22NA, 37 fiber linear array. Each fiber is 200-mm in diameter. The light collected passed 

through a notch filter, and focused onto the slit of an f = 300-mm spectrometer. The slit was set full open, 

to 1-mm. The light was dispersed using a 300 groove/mm grating blazed at 500-nm. The light was 

collected using an intensified CCD camera (1k × 1k pixel array, 18-mm, Gen III, Super Blue intensifier). 

Each data file was comprised of 1005 laser shots, acquired using a combination of 30 shots on-chip with 

67 software accumulations. The combination of input and receiving optics used resulted in a probe 

volume approximately 2.6-mm high, 2-mm diameter, and positioned 4.4-mm above the rig centerline. The 

system was traversed horizontally in 5-mm increments both in X- and Y- directions. 

Phase Doppler Interferometry

Three component Phase Doppler Interferometry (PDI) measurements were performed using a setup 

similar to that displayed in Figure 5 where transmitter and receiver optics were arranged in a 30° off-axis 

forward scatter configuration, with both transmitter and receiver oriented 15° with respect to the 

combustor Y- planes. 

An argon-ion laser was used to generate a green (514.5-nm), blue (488-nm) and violet (476.5-nm) 

beam pair. Due to limited optical access, the three-color PDI system used a single optical head to transmit 

the beam pairs that intersected at the measurement volume. As shown in Figure 5, the beams did not 

provide orthogonal measurements, so a combined optical and geometric transform of the three measured 

velocity channels into the rectilinear coordinate system was required. 

Data were obtained in a point-wise manner. Focal lengths for both the transmitter and receiver were 

500-mm. The measurement volume measured 0.076-mm in diameter by 1.03-mm in length.
NASA/TM—2012-217245 7

Figure 5.—PDI setup. Top: transmitter-receiver orientation about test 

rig; Bottom: arrangement of beams transmitted through the 

transceiver and spatial relation to orthogonal components.

Results and Discussion

In Figure 6 and Figure 7, we present PIV results from the Y=0 plane that help describe the 

combustor/mixer system air flow. The inlet conditions were T3 = 828 K and P3 = 1034-kPa. The time 

between image pairs was 5-μs, with a pair sample rate of 15-Hz. Figure 6 shows a sequence of ten 

consecutive “instantaneous” 2–D velocity fields which allows one to see the variation in flow structure. 

Flow is left to right. Regions in red represent downstream flow and those in blue show upstream motion. 

Although no two instantaneous fields are alike, we observe similar structures within each. The regions 

between injectors exhibit high downstream velocity, whereas upstream, near z = 0, negative velocities 

occur. The average velocity and RMS velocity fields shown in Figure 7, which are the result of averaging 

500 instantaneous fields, serve to confirm those observations. The absolute magnitudes of the 2–D

velocities are plotted as gray-scale contours, and the direction of fluid motion is given using arrows. The 

region immediately downstream of the injectors shows upstream motion and the region between the 

injectors has high downstream velocity. The velocity plot also shows the upstream recirculating eddy 

regions that exist between the injectors where the upstream and downstream regions shear. The black 

contour shows the boundaries of zero velocity that separate the central recirculation zones from the bulk 

downstream motion. The RMS plot shows that the greatest variation in the air velocity occurs in the 

region between injectors. These features are in agreement with CFD calculations (Ref. 12).
NASA/TM—2012-217245 8

Figure 6.—Ten consecutive instantaneous PIV Axial-Vertical velocity Fields in the vertical center plane. (Y = 0). 

T3 = 828 K, P3 = 1034-kPa. Flow is left to right.

Figure 7.—Average velocity field (left) and RMS of velocity (right) within the vertical center plane, (Y = 0) 

computed from 500, 2–D image pairs. T3 = 828 K, P3 = 1034-kPa. Flow left to right.

Once we added fuel and combustion occurred at this inlet temperature, our PDI measurements indicated 

that, on average, the fuel droplets were small, with maximum Sauter mean diameter, D32, around 40-μm 

(Figure 8). The majority of drops were less than 12-μm. Figure 9 shows a result during unseeded 

combustion acquired using the PDI system at inlet conditions T3 = 828 K, P3 = 1379-kPa and φ = 0.45. It is 

an end view at x = 3-mm, with flow coming out of the page that plots the azimuthal velocities for three drop 

size bins: drops less than 10-μm, greater that 10-μm and less than 20-μm, and droplets greater than 20-μm. 

From this perspective, one can see that there is rotation approximately about the centerline, which has little 

to no V-W velocity. The V-W components increase in magnitude with greater distance from the centerline, 

predominately radially outward. Near the center, most of the vectors are overlapped and present mostly as 
NASA/TM—2012-217245 9

axial velocity. Once out near the diffuser circumference, we see some differentiation in velocity, with the 

larger (heavier) drops lagging in direction (assuming those drops smaller than 10-μm follow the flow). 

Although the different directions the drops take become distinguishable, there is no obvious pattern that 

would allow us to say that any one size class moves in a particular manner. The separation does, however, 

show that the larger drops probably assume more of a ballistic trajectory, with minimal influence from the 

local air field.

Figure 8.—Plot showing sauter mean diameter within the Y=0 plane. The 

scatter points represent valid droplet count during 10-sec acquisition time. 

Inlet conditions: T3 = 828 K, P3 = 1034-kPa, φ = 0.45. Flow is left to right.

Figure 9.—End view showing azimuthal velocity vectors differentiated by fuel droplet size for 

three size bins. Inlet conditions: T3 = 828 K, P3 = 1379-kPa, φ = 0.45. Flow is out of page.
NASA/TM—2012-217245 10

Figure 10.—PIV from fuel drops at Y = 0. T3 = 617 K, 

P3 = 1034-kPa. Left: φ = 0.45. Right: φloc = 1.45. 

Flow is left to right.

The PIV measurements acquired during combustion result from droplet light scatter only; as with the 

PDI measurements, the flow is unseeded. In the results presented in Figure 10, the combustor pressure is 

1034-kPa, and the inlet temperature is 617 K. We used a lower inlet temperature because at 828 K, we did 

not have many drops from which to scatter. With lower inlet temperatures, there was less fuel 

vaporization and we were able to acquire images with good contrast and dynamic range for PIV 

processing.

The two images presented in Figure 10 represent two flow configurations. The left image was 

acquired with all nine injectors fueled equally, with φloc = φtot = 0.45 (left); the right image was acquired 

with only the center injector fueled, with φloc = 1.45 and φtot = 0.16. They are plotted on the same scale. 

For this injector, all the inlet air passes through the dome, so all nine swirlers act upon the air. A lot 

more fuel passes through the center injector when it operates alone compared to when all nine injectors 

are fueled; this must be the case in order to maintain stability without the flame blowing out. We find that 

we can measure farther downstream in the one-injector case partly because there is more fuel in the 

region to begin with and also because at this low inlet temperature, less fuel has vaporized. Based on the 

direction of the fluid within the fuel spray cone around the central injector, there is no change in the fuel 

spray angle; it appears that the presence of adjacent air swirlers helps to confine the fuel spray to the 

region immediately downstream from its corresponding air swirler element. Expected larger fuel droplets 

unable to follow the flow also act to confine the fuel mass closer to the exit plane.

We analyzed the flame luminescence images obtained with the ICCD and the high speed camera and 

present the results in Figure 11 to Figure 14. Figure 11 is a matrix showing the average 

chemiluminescence signal for CH* for three inlet temperatures (row-wise) and two fuel staging conditions 

for a combustor pressure of 1034-kPa. The left and center columns have all injectors equally fueled with 

equivalence ratios of 0.45 and 0.35; the last column has results with only the center injector fueled. The 

overall equivalence ratio for the center-only cases is 0.15 or 0.16. Flow is from left to right. The images 

were obtained by integrating, on-chip, 600 50-ns wide gates (as for PLIF), for a total exposure of 30-μs 

over the 1 minute data acquisition time. The images shown are auto-scaled to highlight their structural 

features.
NASA/TM—2012-217245 11

Figure 11.—On-chip integrated averages of CH* at various inlet conditions. Rows (top to bottom): T3 = 828, 728, 

617 K. Columns (L-M): φ = 0.45, 0.35 Right column, center-only fueled. Flow is left to right.
NASA/TM—2012-217245 12

The extent of signal downstream follows from the amount of fuel that flows from a given injector. 

For example, in considering the center injector row for the 828 K inlet temperature, the CH* signal 

extends from X = 10-mm (φloc = 0.35) to X = 16-mm (φloc = 0.45) to X = 20-mm (φloc = 1.32). For centeronly fueling, as well, we note a delay in the onset of chemical activity, i.e., the flame front sits farther 

downstream from the dome than with the corresponding all-fueled conditions. Regardless of fueling, we 

also observe an upstream shift in the location of peak signal intensity with increase in inlet temperature.

Figure 12 shows statistical results of processing 9001 consecutive flame luminosity images acquired 

at 10-kHz, and is typical of all high speed image sets. (Figure 13 shows a sequence of 10 such 

instantaneous images.) The inlet conditions were T3 = 828 K, P3 = 1034-kPa, and φ = 0.35. The average 

of those images is representative of the steady-state system (and identical in form to those obtained via 

the ICCD), while the RMS (standard deviation) of those images tells us there is a large variation about the 

mean, particularly in the region immediately downstream of the fuel injection sites. The third image 

shows the signal range (signal maximum – signal minimum) and like the RMS, indicates a large 

variation, but it also serves to point out that variation extends well downstream, where the RMS is low.

We also processed the images in a manner typical for PIV processing. For these images, we are 

tracking the bulk flame intensity (as represented by C2

* and CH* within the flame front) changes over 

time. The time between images is 100-μs, a relatively long time compared to the 5-μs time between PIV 

image pairs. However, in observing the play-back of the high speed movies, and in considering that we 

can track by eye, “movement” from frame to frame (as in the single frames of Figure 13), the time 

between frames is not too long. It is reasonable that PIV-type processing may provide insights into LDI 

flame structure, especially since C2

* and CH* are indicators of the primary reaction zone. It is also 

important to keep in mind that each individual image from the high speed camera captures the light 

emitted throughout the volume of the optically-accessible region and does not isolate individual planes as 

with PLIF and PIV. 

The results are shown in Figure 14, with the combustor operating at 1034-kPa. In the top row are 

images acquired with an inlet temperature of 828 K and in the bottom row, 617 K. The left column has all 

injectors evenly-fueled with equivalence ratio of 0.35 for the top image and 0.45 for the bottom image. In 

the right column, only the center injector is fueled, with φloc = 1.32 for the top and φloc = 1.45 at the 

bottom. The contour denotes the primary flame zone and the arrows can be interpreted as showing the 

average direction of motion within the flame zone.

Figure 12.—Statistics derived from processing the signal intensities of 9001 consecutive flame images. Inlet 

conditions: T3 = 828 K, P3 = 1034-kPa, φ = 0.35, flow left to right.
NASA/TM—2012-217245 13

Figure 13.—Ten Consecutive Instantaneous Frames from High Speed Movies, framed at 10 kHz. Inlet conditions: 

T3 = 672 K, P3 = 1034 kPa, φ = 0.45. Flow is left to right.

In considering the images of Figure 14 we note the following observations. The high temperature case 

with all nine injectors fueled is the odd case in terms of the flame visualization observed. It is the only 

case that does not indicate recirculating eddies within the central core located directly downstream of the 

fuel injection sites. However, earlier work using phase Doppler interferometry (Ref. 10) indicates a 

weakened recirculation zone during combustion. The lower temperature, all-fueled case also supports that 

work because we observe a much shorter recirculation region (Figure 14, lower left). In comparing the 

center-fueled injector cases to those with all nine injectors, the former show relatively strong recirculation 

near the dome and a much greater extent of coherent structure much farther downstream, at x = 35-mm. 

This is commensurate with the flame zone being longer because of the fuel-rich front end. In contrast, 

the two evenly fueled cases do not show coherent structure beyond x = 20-mm at low temperature and 

x = 15-mm at high temperature.

It is possible that heat release effectively decreases the swirl number; the decrease in swirl number 

would lead to decreased flame stability. As the fuel burns, the overall density of the fuel/air mixture 

decreases. This density decrease leads to an increase in axial velocity because the mass flow rate remains 

the same. However, it can be argued that the radial velocity remains approximately the same because the 

radial velocity is imparted to the flow by swirlers that are upstream of the combustion zone. Experiments 

with swirling jets show that the recirculation zone disappears for low swirl numbers (Ref. 16).
NASA/TM—2012-217245 14

Figure 14.—Average Flame/Flow structure derived using PIV-type processing of 9000 successive 

high-speed image pairs. Top row: T3 = 828 K. Bottom row: T3 = 617 K. Left column: Nine injectors 

fueled equally. Right column: Center injector fueled only.
NASA/TM—2012-217245 15

Figure 15 serves to help characterize combustion at a single inlet condition using complementary 

diagnostics techniques: CH* chemiluminescence, OH and fuel PLIF, and spontaneous Raman 

spectroscopy. T3 = 828 K, P3 = 1034-kPa, and φ = 0.35. Key features are that the injector produces a 

highly symmetric flow field. The fuel PLIF (bottom left) shows that the fuel, as represented by 

naphthalene-based molecules, is consumed quickly. CH* (upper left) supports this; CH 

breakdown/recombination begins immediately at the dump plane, and the primary reaction zone is located 

there. The plot of OH PLIF (upper right) shows axial, horizontal, and vertical slices from a 3–D data 

block, highlighting the symmetry, but also showing (along with the CH*

) that little or no interaction 

occurs between LDI elements. Finally, the SRS (lower right) data provides us with reasonable 

combustion temperatures 4-mm above the 9-point LDI centerline, and shows that the high temperature 

regions lie immediately downstream of the injector elements. The SRS data also suggest little interaction, 

and its form is consistent with 1-point CFD results (Ref. 13).

Figure 15.—Results using (top row) CH* and OH PLIF and (bottom row) Fuel PLIF and 

SRS at T3 = 828 K, P3 = 1034-kPa, and φ = 0.35.
NASA/TM—2012-217245 16

Summary

In this paper, we have presented recent measurements taken in an optically-accessible, JP8-fueled, 

flame tube combustor using baseline nine-element lean direct injection (LDI) research hardware. 

Extensive measurements were presented for the cases where all elements are evenly fueled and 

measurements where only the center element is fueled. PIV and chemiluminescence measurements show 

that the burning zone extends farther downstream for the center-only cases than it does for the 

corresponding equally-fueled cases; this is consistent with the local equivalence ratio being much higher 

for the center-only cases than it was for the evenly-fueled cases. Results from the high speed 

chemiluminescence measurements suggest that combustion reduces the size of the recirculation zone. 

PLIF, chemiluminescence, and SRS results suggest that the burning results for each element are local and 

that the interaction between elements is limited.

References

1. U.S. Government Accountability Office. Report to Congressional Committees. “Aviation and 

Climate Change: Aircraft Emissions Expected to Grow, but Technological and Operations 

Improvements and Government Policies Can Help Control Emissions,” GAO-09-554, 2009.

2. U.S. Congress. Partnership for Air Transportation Noise and Emissions Reduction. “Report to the 

United States Congress: Aviation and the Environment,” 2009. http://web.mit.edu/aeroastro/partner/ 

reports/congrept_aviation_envirn.pdf; Accessed: 4/21/2011.

3. U.S. Federal Aviation Administration, “Aviation & Emissions: A Primer,” Office of Environment 

and Energy, Washington, DC 2005.

4. Johnson, H.S., Kinnison, D.E., Wuebbles, D.J., “Nitrogen Oxides from High Altitude Aircraft: An 

Update of Potential Effects on Ozone,” J. of Geophysical Research, Vol. 94, No. D13, 1989, p. 

16,351-16,363.

5. Tacina, R., Lee, P., and Wey, C., “A Lean-Direct-Injection Combustor Using a 9 Point SwirlVenturi Fuel Injector,” ISABE-2005-1106, 2005.

6. Locke, R.J, Hicks, Y.R, and Anderson, R.C., “Combustion Imaging Using Fluorescence and Elastic 

Scattering,” Optical Metrology for Fluids, Combustion, and Solids, edited by C. Mercer, Kluwer 

Academic Publishers, Norwell, MA, 2003, pp. 175-208.

7. Hicks, Y.R., Locke, R.J., and Anderson, R.C., “Optical Measurement and Visualization in HighPressure, High-Temperature, Aviation Gas Turbine Combustors,” Optical Diagnostics for Industrial 

Applications, edited by Neil A. Halliwell, editor, Proceedings of SPIE Vol. 4076, 2000, pp. 66–77.

Also published as NASA/TM—2000-210377, 2000.

8. Cai, J., Jeng, S.-M., and Tacina, R., “The Structure of a Swirl-Stabilized Reacting Spray Issued 

From an Axial Swirler,” AIAA-2005-1424, 2005.

9. Hicks,Y.R., Anderson, R.C., and Locke, R.J., “Optical Measurements in a Combustor Using a 9-

Point Swirl-Venturi Fuel Injector,” International Society for Air Breathing Engines, ISABE-2007-

1280, 2007.

10. Heath, C.M., Locke, R.J., Anderson, R.C., and Hicks, Y.R., “Optical Characterization of a 

Multipoint Lean Direct Injector for Gas Turbine Combustors: Velocity and Fuel Drop Size 

Measurements,” IGTI, GT2010-22960, 2010.

11. Chen, K.-H., Norris, A.T., Quealy, A., and Liu, N.-S., “Benchmark Test Cases for the National 

Combustion Code,” AIAA-1998-3855, 1998.

12. Davoudzadeh, F., Liu, N.-S., Moder, J.P., “Investigations of Swirling Air Flows Generated by Axial 

Swirlers in a Flame Tube,” IGTI, GT-2006-91300, 2006.

13. Iannetti, A.C., and Moder, J.P., “Comparing Spray Characteristics from Reynolds Averaged NavierStokes (RANS) National Combustion Code (NCC) Calculations Against Experimental Data for a 

Turbulent Reacting Flow,” AIAA-2010-578, 2010. Also published as NASA/TM—2010-216735, 

2010.
NASA/TM—2012-217245 17

14. Iannetti, A., Tacina, R., Jeng, S.-M., and Cai, J., “Towards Accurate Prediction of Turbulent, ThreeDimensional, Recirculating Flows With the NCC,” AIAA-2001-0809, 2001. Also published as 

NASA/TM—2001-210761.

15. Stubbs, R.M., and Liu, N.-S., “Preview of the National Combustion Code,” AIAA-97-3114, 1997.

16. Beer, J.M., Chigier, N.A., Combustion Aerodynamics, John Wiley and Sons, Inc., New York, 1972.
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Investigations of a Combustor Using a 9-Point Swirl-Venturi Fuel Injector: Recent 

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14. ABSTRACT

This paper explores recent results obtained during testing in an optically-accessible, JP8-fueled, flame tube combustor using baseline Lean 

Direct Injection (LDI) research hardware. The baseline LDI geometry has nine fuel/air mixers arranged in a 3 × 3 array. Results from this 

nine-element array include images of fuel and OH speciation via Planar Laser-Induced Fluorescence (PLIF), which describe fuel spray 

pattern and reaction zones. Preliminary combustion temperatures derived from Stokes/Anti-Stokes Spontaneous Raman Spectroscopy are 

also presented. Other results using chemiluminescence from major combustion radicals such as CH* and C2* serve to identify the primary 

reaction zone, while OH PLIF shows the extent of reaction further downstream. Air and fuel velocities and fuel drop size results are also 

reported. 

15. SUBJECT TERMS

Laser diagnostics; Fuel injection; Combustion 

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