Christopher M. Heath and Robert C. Anderson

Glenn Research Center, Cleveland, Ohio

Randy J. Locke

Arctic Slope Regional Corporation, Cleveland, Ohio

Yolanda R. Hicks

Glenn Research Center, Cleveland, Ohio

Optical Characterization of a Multipoint Lean Direct 

Injector for Gas Turbine Combustors: Velocity and 

Fuel Drop Size Measurements

NASA/TM—2010-216365

October 2010
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Christopher M. Heath and Robert C. Anderson

Glenn Research Center, Cleveland, Ohio

Randy J. Locke

Arctic Slope Regional Corporation, Cleveland, Ohio

Yolanda R. Hicks

Glenn Research Center, Cleveland, Ohio

Optical Characterization of a Multipoint Lean Direct 

Injector for Gas Turbine Combustors: Velocity and 

Fuel Drop Size Measurements

NASA/TM—2010-216365

October 2010

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Turbo Expo 2010

sponsored by the American Society of Mechanical Engineers (ASME)

Glasgow, Scotland, United Kingdom, June 14–18, 2010
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, Harold Redloske, and Tom Ralys.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

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Available electronically at http://gltrs.grc.nasa.gov

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—2010-216365 1 

Optical Characterization of a Multipoint Lean Direct Injector for Gas 

Turbine Combustors: Velocity and Fuel Drop Size Measurements 

Christopher M. Heath and Robert C. Anderson 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Randy J. Locke 

Arctic Slope Regional Corporation 

Cleveland, Ohio 44135 

Yolanda R. Hicks 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Performance of a multipoint, lean direct injection (MP–LDI) 

strategy for low emission aero-propulsion systems has been 

tested in a Jet-A fueled, lean flame tube combustion rig. 

Operating conditions for the series of tests included inlet air 

temperatures between 672 and 828 K, pressures between 

1034 and 1379 kPa and total equivalence ratios between 0.41 

and 0.45, resulting in equilibrium flame temperatures 

approaching 1800 K. Ranges of operation were selected to 

represent the spectrum of subsonic and supersonic flight 

conditions projected for the next-generation of commercial 

aircraft. 

This document reports laser-based measurements of in situ

fuel velocities and fuel drop sizes for the NASA 9-point LDI 

hardware arranged in a 3×3 square grid configuration. Data 

obtained represent a region of the flame tube combustor with 

optical access that extends 38.1-mm downstream of the fuel 

injection site. All data were obtained within reacting flows, 

without particle seeding. Two diagnostic methods were employed to evaluate the resulting flow path. Three-component 

velocity fields have been captured using phase Doppler interferometry (PDI), and two-component velocity distributions 

using planar particle image velocimetry (PIV). Data from 

these techniques have also offered insight into fuel drop size 

and distribution, fuel injector spray angle and pattern, turbulence intensity, degree of vaporization, and extent of reaction. 

This research serves to characterize operation of the baseline NASA 9-point LDI strategy for potential use in future 

gas-turbine combustor applications. An additional motive is 

the compilation of a comprehensive database to facilitate 

understanding of combustor fuel injector aerodynamics and 

fuel vaporization processes, which in turn may be used to 

validate computational fluid dynamics codes, such as the 

National Combustor Code (NCC), among others. 

Nomenclature 

MP–LDI Multi-point Lean Direct Injection 

LDI Lean Direct Injection 

CAEP Committee on Aviation Environmental Protection 

ICAO International Civil Aviation Organization 

NCC National Combustion Code 

PDI Phase Doppler Interferometry 

PIV Particle Image Velocimetry 

LTO Landing-Take-Off Cycle 

ICCD Intensified Charge-Coupled Device 

Nd:YAG Neodymium-Doped Yttrium Aluminum Garnet 

UHC Unburned Hydrocarbon 

HC Hydrocarbon 

FWHM Fixed Width at Half Maximum 

CO Carbon Monoxide 

NOx Oxides of Nitrogen 

RMS Root mean square 

Φ [-] Equivalence Ratio, fuel-to-air basis 

X [mm] Axial flow direction 

Y [mm] Horizontal flow direction 

Z [mm] Vertical flow direction 

U [m/s] Axial velocity component 

V [m/s] Horizontal velocity component 

W [m/s] Vertical velocity component 

D10 [μm] Linear drop diameter 

Subscripts: 

in inlet 

T total 
NASA/TM—2010-216365 2 

Introduction 

Aviation emissions are becoming an internationally recognized environmental concern as air traffic continues to rise in 

an effort to keep pace with global economic, military, and 

private demands. In response, laws regulating pollutant levels 

are growing progressively stringent as governing organizations continue to tighten existing emission certification standards and adopt new policies for managing the 

next-generation of aero-propulsion systems. The Committee 

on Aviation Environmental Protection (CAEP), a subset of 

the International Civil Aviation Organization (ICAO), is 

responsible for reviewing engine performance trends and 

recommending operating practices and standards to further 

mitigate the impact of civil aviation on global atmosphere and 

climate. Current ICAO regulations limit ground level emissions of hydrocarbons (HC), carbon monoxide (CO), and 

oxides of nitrogen (NOx) based on landing and take-off cycle 

(LTO) operation. Future laws are expected to more closely 

regulate exhaust products at altitude or impose emission 

related levies in the form of landing taxes, with collected 

funds intended for financing research to reduce the environmental impact of civil air transport (Refs. 1 and 2). 

Advanced combustion and fuel injection methods will be 

necessary to overcome the challenges posed by impending 

emission laws. Aircraft will need to operate within the constraints of increasing regulations while satisfying the potentially complex mission objectives expected of next-generation 

vehicles. Multi-point lean direct injection (MP–LDI) is a 

NASA developed low-emission fuel injection technology that 

appears to assist in meeting the dual requirements of low-NOx

production and high performance. The strategy was developed to facilitate rapid, stable, and uniform combustion under 

lean operation, resulting in lower combustor peak flame temperatures and reducing primary emission contributions from 

the thermal NOx mechanism. MP–LDI schemes appear to 

maintain CO and unburned hydrocarbon emission levels 

relative to limits established in the 1996 ICAO standard 

(Refs. 3 and 4). 

The MP–LDI concept incorporates a large number of small 

scale fuel injectors in place of a conventional, single injector. 

The reduced scale promotes rapid mixing of fuel and oxidizer, 

and the conditions at the flame front approach those in a premixed, prevaporized flame. Multiple injectors can be integrated into modules to allow for a variety of fuel staging 

possibilities appropriate for an LTO engine cycle. MP–LDI 

was chosen as our fuel injector/mixer test case for validating 

the National Combustion Code (NCC) because the concept 

provides low emissions, allows for different fuel staging 

strategies, and has realistic, complex fluid interactions. 

A large number of variables play a role in developing and 

optimizing low emission fuel injection strategies. Design 

choices include the number and size of injector elements, 

whether to use radial or axial air swirlers, swirler angle and 

direction, and type of fuel nozzle. Tacina et al. (Ref. 4) 

considered these factors in comparing combustion performance, based on NOx and CO emissions. 3×3, 5×5, 6×6, and 

7×7 element arrays were studied. NOx emissions were low in 

all cases and found to decrease with swirl number and combustor operability range. 

The MP–LDI concept is currently under evaluation and 

testing to support future gas-turbine combustor applications. 

In this study, the multipoint injection strategy has been tested 

in a square array configuration comprised of nine LDI elements arranged in a 3×3 grid. This configuration has been 

designated as the baseline configuration. 

Several useful parameters for defining fuel injection operation and emission levels include liquid and gas phase velocities, turbulence intensity, and fuel vaporization. Optical and 

laser-based diagnostics have been employed to capture this 

information for the 9-point hardware at test points simulating 

subsonic and supersonic flight conditions. This document 

focuses on reporting and comparing results from two diagnostic techniques with emphasis on (1) developing a database for 

verifying computational combustion flow models and 

(2) assessing and refining diagnostic techniques specifically 

for application to our facility which operates at realistic aircraft combustor conditions. 

We describe here our efforts to use two complementary, 

laser-based optical techniques to acquire and compare 

velocity distributions and RMS measurements. Phase Doppler 

interferometry (PDI) provided the means to capture 

three-component velocity fields, turbulence intensity and 

injector full cone spray angle. Using planar particle image 

velocimetry (PIV) we were able to acquire two-component 

fuel spray velocity distributions as well as injector spray 

angle. The results from both methods were very similar in all 

cases where direct comparison was possible. 

In this work all experimental data were acquired using 

non-preheated Jet-A fuel supplied to the combustor through 

the 9-point lean direct injection (LDI) hardware. 

Facility 

The baseline hardware was tested in a high temperature 

and high pressure flametube combustor rig with partial optical access directly downstream of the fuel injection site. The 

experimental facility and diagnostic instrumentation offered 

the combined unique ability to perform non-intrusive optical 

and laser based techniques demonstrating fluid and chemical 

processes during combustion. The test cell in this study was 

capable of supplying non-vitiated inlet air to the combustor at 

temperatures between 450 and 1100 °F (505 to 866 K), pressures between 150 and 250 psia (1034 to 1724 kPa) with flow 

rates up to 10-lbm/s (4.54 kg/s). Total equivalence ratios supporting stable lean operation of the flametube were generally 

between 0.33 and 0.55 depending on the simulated engine 

conditions. The entire rig was operated remotely and sustained 

by an adjacent control room and laser facility. 
NASA/TM—2010-216365 3 

An isolated full-section view of the window assembly and 

aft flametube are displayed in Figure 1. The combustor housing measures 74-cm in length and is lined with an aluminum 

oxide ceramic to withstand average temperatures reaching 

3200 °F (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 has a cross 

section measuring 76.2 by 76.2-mm. The bottom of the chamber is reserved for a spark igniter system and provides no 

physical or optical access. The grade of quartz used in the 

windows offers a minimum light transmission of 80 percent 

within a spectral range from 180 to 3300-nm, making it ideal 

for supporting optical measurements (Ref. 5). 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. The external housing and downstream flange 

section are water cooled to reduce thermal and material 

stresses at high temperature and pressure conditions. 

Injector Architecture 

The baseline 9-point test hardware is described in Reference 4 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 60o surrounding a simplex-type nozzle 

that atomizes and injects liquid fuel (Jet-A) at the throat of a 

converging-diverging venturi. The calculated swirl number is 

1.0. A schematic of a single injector element is displayed in the 

full section view of Figure 2. The array is shown in Figure 3. 

Diagnostic Setup and Procedures 

Two optical diagnostic methods were used to interrogate 

the 9-point LDI combusting flow and acquire a number of 

critical flowfield parameters. Phase Doppler interferometry 

(PDI) was used to characterize the vaporization and mixing of 

the 9-point LDI by providing mean diameter statistics, fuel 

injector spray angle and pattern, three-dimensional liquid 

velocity field characteristics, and turbulence levels. Particle 

image velocimetry (PIV) was used to obtain planar twocomponent fuel velocity distributions further demonstrating 

directional flow properties and emphasizing interaction zones 

between injectors. Both PDI and PIV are well-established 

techniques that have commercially-available systems. 

For both techniques, the optical instrumentation was fixed to 

a set of orthogonal stages to provide traversing ability in either 

one or three-dimensions (depending on technique) in order to 

map out the accessible region-of-interest within the combustor.
NASA/TM—2010-216365 4 

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. 

In this work X- is defined to be the stream-wise direction, 

with X = 0-mm being the exit (dump) plane of the 9-point 

LDI hardware. Y- is the horizontal direction and Z- is the 

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

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

Particle Image Velocimetry 

The experimental setup for PIV was comparable to the 

Figure 4 illustration. PIV data were obtained using a dual 

head, frequency-doubled Nd:YAG laser operating at 15 Hz. 

The laser beams were expanded into sheets using a cylindrical 

lens and transmitted vertically into the combustor rig, with 

the sheet axis aligned with the flow direction. Images were 

acquired through a side window using the light scattered from 

fuel droplets to capture liquid phase velocity. We used an 

f = 200-mm, f/4 lens to collect the light. This collection 

scheme imaged a region that spanned roughly 17-mm axially 

by 23-mm vertically, to isolate the center row of injectors. An 

interline transfer CCD camera captured the scattered light 

from adjacent laser pulses onto two individual frames. 

Because the second frame of standard interline transfer cameras has relatively long exposure times, the camera lens was 
NASA/TM—2010-216365 5 

outfitted with a mechanical leaf shutter capable of operating at 

15-Hz. All synchronization was handled using a LaVision, Inc. 

programmable timing unit, PTU-8. Cross-correlation methods 

were then applied to create a displacement vector field for 

each image pair. LaVision’s DaVis software version 7.2 was 

used to collect and process the images and produce axialvertical 2–D velocity fields. To map the optically-accessible 

volume, we traversed across the flow in the Y-direction in 

1-mm increments. 

Phase Doppler Interferometry 

PDI was performed using a setup similar to that displayed 

in Figure 5 which employed transceiver and receiver optics in 

a 30° off-axis forward scatter arrangement. Each transceiver 

beam was separately fine-adjusted to compensate for refraction through the combustor window and ensure that all beams 

overlapped within the rig. 

PDI measurements were obtained using a three-component 

fiber-based system consisting of a 112-mm diameter FiberFlow probe (Model 5402–01), BSAFlow software and 

processing unit developed by Dantec Dynamics. An argonion laser was used to generate the green (514.5-nm), blue 

(488-nm) and violet (476.5-nm) beam pairs for the transceiver. 

Due to limited optical access, the 3-color PDI system used a 

single optical head to transmit three beam pairs that intersected at the measurement volume. As shown in Figure 6(a), 

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. Also, since this transformation was manually 

calculated and applied, we do not present RMS measurements 

for the channels; rather we calculate and report the local turbulence intensities, which can be obtained from the raw mean and 

RMS values.

Data were obtained in a point-wise manner using two different optical configurations to more fully access the flowfield. These arrangements and the combustor regions 

accessed are shown in Figure 6(b). Focal lenses for the transceiver and receiver were 310-mm or 500-mm depending on 

optical access and probe orientation. The 310-mm focal 

length lens produced a measurement volume of 0.047-mm in 

diameter by 0.4-mm in length while the dimensions increased 

to 0.076-mm in diameter by 1.03-mm in length for the 

500-mm focal length lens. 

For the upstream setup, the transmitter and receiver were 

angled upstream with their optic axes in a horizontal plane 

(Z = constant) (see Fig. 5). This allowed us to probe the region 

immediately downstream of the center injector element, from 

the dump plane to approximately X = 10-mm. 

The downstream setup was similar to Figure 5, but with the 

transmitter and receiver angled downward. Their optical axes 

were in a vertical plane (X = constant). This permitted access 

to the lower central region of the combustor, from approximately X = 9-mm to 15-mm. 
NASA/TM—2010-216365 6 

(a) Probe details. (b) Approximate combustor regions probed. 

Figure 6.—PDI setup showing (a) arrangement of beams transmitted through the transceiver and their spatial relation to orthogonal components; (b) illustration that highlights the PDI beam insertion orientations and regions probed for the upstream 

(orange) and downstream (green) setups. 

Results and Discussion 

All data were collected during combustion, with calculated 

adiabatic flame temperatures between 1700 to 1833 K. Five 

inlet conditions were considered, including high (H), medium 

(M) and low (L) inlet temperatures and two pressures, designated (1) and (2). Table 1 provides the combustor test matrix 

including air flow rates and equivalence ratios. For all points, 

the air inlet pressure drop was 4 percent. The last three columns indicate the diagnostic techniques that were used to 

characterize operation at each condition. For the work presented, each diagnostic technique was used on a different test 

day and no techniques were performed simultaneously. 

All experimental data have been processed and are presented in formats that attempt to thoroughly describe the fluid 

behavior of each test point in a clear, visual manner. Data are 

presented as either scatter plots, 1–D line plots or 2–D field 

plots. Two-dimensional plots are either side view plots at the 

Y = 0 plane with the downstream flow direction left to right, 

or end view plots (X = constant) with downstream flow out of 

the page. 

The results presented reflect an analysis of the liquid fuel 

spray without regard for the atomizing air. To provide some 

insight into the air flow structure in the absence of fuel, 

Figure 7 depicts a computational result that shows the predicted air flow features present downstream of the 9-point 

LDI (Ref. 6). The key feature to point out is the air-swirlinduced recirculation zone which forms immediately downstream of each injector element. Recirculation zones are 

necessary to support and stabilize the flame(s) in aircraft 

combustors. Figure 7 also shows the zero-velocity contours 

within which these recirculation zones exist. Given this 

inherent structure, we will expect to see reverse flow of the 

fuel droplets within this region. The expected reverse flow 

was confirmed with our PIV measurements (Ref. 7). 

In other related work with a single LDI, Cai et al. (Ref. 8) 

studied both reacting and non-reacting cases at ambient 

TABLE 1.—INLET CONDITION TEST MATRIX 

Test 

point 

Tin

(K) 

Pin

(kPa) 

Wair

(kg/sec) 

T PIV PDI 

H1 828 1034 0.45 0.41 --- X 

H2 828 1379 0.61 0.45 --- X 

M1 728 1034 0.50 0.45 X X 

M2 728 1379 0.68 0.45 --- X 

L1 672 1034 0.52 0.45 --- X 

Figure 7.—Air-only computational 

result showing the zero-velocity 

isocontours of the 9-point LDI 

hardware (from Ref. 6). 

temperature (~298 K) and pressure. Using PDI, they found that 

strong swirl affects mean diameter profiles, providing more 

uniform sizes and that heat released during combustion 

increases the mean and turbulent velocities relative to the nonreacting case. They concluded that the interaction between 

liquid and gas phases caused a decrease in the recirculation 

zone during combustion. The single LDI used in the Cai work 

has the same physical geometry as each element in the 9-point 

configuration used in this work, so Cai provides a useful 

 
NASA/TM—2010-216365 7 

comparison. However, the work by Cai used a square combustion chamber measuring 38-mm per side. To be exactly comparable with our work, the chamber should be 25-mm per side. 

To date, there is no study that compares the single element 

with this baseline 9-point LDI that duplicates both inlet and 

boundary conditions. 

Gross Features (PDI) 

Figure 8 contains side views within the vertical center 

plane of the mean axial and vertical velocities (vectors) and 

the local turbulence intensity contours immediately downstream of the center injector for the 728 and 828 K inlet conditions. Local turbulence intensity was calculated using the 

ratio of the RMS of the turbulent velocity fluctuations to the 

local Reynolds averaged mean velocity. Streamlines were 

added to the plots at ±1⁄2 of the diffuser exit radius to represent 

the outer boundaries of the hollow cone spray. We assumed 

and verified by measurement count number that the average 

bulk mass of fuel spray traveled through the paths swept out 

by the streamlines. The streamlines were then used to estimate spray angle for the center injector at each test condition. 

More accurate spray angle measurements could be made if 

the average radial distance to the centroid of the injected cone 

spray was precisely known. All images in Figure 8 show 

similar trends: (1) a spray angle roughly 80° full cone with 

greater tangential velocity near the dump plane, which suggests the main flow from the nozzle is a hollow cone spray; 

(2) the region near the injector centerline (within the cone) 

has lower velocity; (3) the highest axial velocities occur at or 

near the region where the flow from the upper and lower 

injectors mix with that of the center injector. The region 

where the jets mix is also the area of highest turbulence intensity. The second highest levels of turbulence occur near the 

injector centerline, within the first few millimeters from the 

dump plane. The lowest levels of turbulence occur around the 

boundaries of the spray cone. 

Figure 9 shows PDI results obtained at point M1, which 

plots in the top row the V–W vectors and U-contours for 

X = 0-mm (left), 3-mm (center), and 6-mm (right column). 

The bottom row shows the turbulence intensity contours 

along with the droplet count (scatter plot) that corresponds 

with each axial position. The circles represent the projected 

diffuser exit diameters at the dump plane. 

From the axial velocity contours, one can visualize the 

region of core recirculation, whose average fuel drop velocity 

is positive. The V–W vectors reflect the tangential swirl imparted to the fuel from both the fuel nozzle and the swirling 

air. The turbulence contours from these perspective views 

highlight the higher turbulence intensity along the centerline 

at the injector exit plane and show that levels dissipate from 

that central core as the flow progresses downstream. The 

scatter plot represents the number of drops measured at each 

location. 

Figure 10(a) shows a representative contour of the mean 

V velocity. Figure 10(b) shows a representative contour of 

the mean W velocity. The Z-Y plots show results from inlet 

condition M1, the Z-X plots show M2 results. Given our 

limited optical access, these plots are useful to highlight the 

symmetrical flow patterns from the 9-point LDI hardware. 

Point M1 Point M2 Point H1 Point H2

Figure 8.—Side views at y = 0 that show the average U–W fuel velocity vectors and local turbulence intensity 

contours for the medium and high inlet temperature conditions. Flow is left to right. 
NASA/TM—2010-216365 8 

X = 0-mm X = 3-mm X = 6-mm 

Figure 9.—Results from inlet condition M1 at X = 0, 3, and 6-mm from injector dump plane. Top row: 3–D velocity field 

with axial velocities represented by contour. Bottom row: Turbulence intensity contours and scatter plot representing 

number of drops measured. Positive X flow direction is out of the page. 

(a) (b) 

Figure 10.—Examples that highlight symmetry. (a) Contour plot of mean V components from PDI measurements. (b) Contour 

plot of mean W components from PDI measurements. The Z-Y plots show end views (flow out of page) at the X = 3-mm plane 

at inlet condition M1. The Z-X plots show the side view at Y = 0, with flow from left to right, at inlet condition M2. 

Y (mm)

Z (mm)

-5 0 5 10 15 -15

-10

-5

0

5

10

15

U-Mean (m/s)

70.0

65.0

60.0

55.0

50.0

45.0

40.0

35.0

30.0

25.0

20.0

15.0

50 m/s

Y (mm)

Z (mm)

-5 0 5 10 15 -15

-10

-5

0

5

10

15

U-Mean (m/s)

70.0

65.0

60.0

55.0

50.0

45.0

40.0

35.0

30.0

25.0

20.0

15.0

50 m/s

Y (mm)

Z (mm)

-5 0 5 10 15 -15

-10

-5

0

5

10

15

U-Mean (m/s)

70.0

65.0

60.0

55.0

50.0

45.0

40.0

35.0

30.0

25.0

20.0

15.0

50 m/s

Y (mm)

Z (mm)

-5 0 5 10 15 -15

-10

-5

0

5

10

15

Turbulence

0.7

0.6

0.6

0.5

0.5

0.4

0.3

0.3

0.3

0.2

0.1

0.1

0.1

Y (mm)

Z (mm)

-5 0 5 10 15 -15

-10

-5

0

5

10

15

Turbulence

0.7

0.6

0.6

0.5

0.5

0.4

0.3

0.3

0.3

0.2

0.1

0.1

0.1

Y (mm)

Z (mm)

-5 0 5 10 15 -15

-10

-5

0

5

10

15

Turbulence

0.7

0.6

0.6

0.5

0.5

0.4

0.3

0.3

0.3

0.2

0.1

0.1

0.1
NASA/TM—2010-216365 9 

Drop Size Differentiated Flows (PDI) 

Figures 11 through 15 show results when considering fuel 

drop size. Because the average drop size range was small, we 

have three size bins. These three D10 size ranges are: D10

smaller than 10-μm, between 10- to 20-μm and greater than 

20-μm. 

Figure 11 shows the mean axial velocity at inlet condition 

M1 for the three drop size ranges along a vertical rake within 

the Y = 0-mm plane immediately downstream of the center 

injector. As expected, the largest drops are typically the slowest, and the smallest drops move fastest. Only the drops larger 

than 10-μm have negative average axial velocity, observed at 

the 3-mm downstream location, though all drops immediately 

downstream have relatively low axial velocity. The highest 

velocities occur in the region between adjacent injectors, 

around Z±12-mm. By X = 9-mm, only small drops remain. 

 Figure 12 presents V–W vectors resultant from the H2 

inlet condition at X = 3-mm. From this perspective, one can 

see that the center of rotation is near the centerline, which has 

little to no V–W velocity. The V–W components increase 

with greater distance from the centerline, predominately 

radially outward. The tangential influence of the incoming 

swirling air is evident in that we begin to see separation in the 

direction the drops move, depending on their size. 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. 

Distributions of the axial velocity component versus linear 

droplet diameter, D10, for the 728 and 828 K cases are displayed in Figure 13. Data contained in these images have 

been processed to remove outlier drops with unusually high 

velocity components well beyond average. In all cases, the 

region probed was within the vertical Y = 0-mm plane immediately downstream of the center injector, from the Z = 0-mm 

position to below the centerline at Z = –15-mm. Low counts 

at the X = 0-mm plane are attributed to high optical thickness 

at the direct exit of the injection system. 

The fuel injector pressure drop for the low combustor pressure cases (left column) is approximately 172 kPa (25 psid) 

below that of the higher pressure cases (right column). For 

each set, a similar trend may be noted. Plots for all test conditions show a general decrease in fuel drop size and count with 

downstream progression as fuel particles are being consumed 

by chemical reaction and vaporization processes. 

The high inlet temperature conditions present smaller fuel 

drops on average, compared to the lower inlet temperature 

cases. These drops also vaporize more rapidly, as no drops 

are observed at the X = 9-mm location. Also, the higher temperature condition shows that a greater percentage of the 

drops measured have positive axial velocity. As inlet temperature increases, the fraction of drops that have negative 

axial velocity decreases. This suggests the strength of the 

recirculation zone decreases with higher inlet temperature. 

Figure 11.—Axial velocity profiles along Y = 0 for three drop 

size ranges, at X = 0-, 3-, 6-, and 9-mm. Inlet condition M1.

Figure 12.—Velocity vectors for three drop size bins at the 

X = 3-mm plane. Size bins: small drops <10 μm; medium, 

10 to 20 μm; large >20 μm. Inlet condition H2. Flow out of 

page.

A greater degree of atomization is expected with increased 

fuel injector pressure drop, and the bottom row results serve 

to demonstrate this behavior. Despite the higher pressure case 

flowing 0.0057-kg/s (0.0128-lbm/s) more fuel, the number 

of drops measured is lower by more than an order of magnitude, and at each axial position, the average drop diameter 

decreases. The reason this effect is not apparent for the high 

inlet temperature cases is because the fuel-to-air ratios are not 

equivalent. Condition H1 had  = 0.41 instead of 0.45, and 

disproportionately less fuel than its high pressure counterpart, 

for a difference of 0.0053 rather than 0.0049 kg/s. Another 

factor is that the fuel more readily vaporizes at this higher inlet 

U-Mean (m/s)

Z (mm)

0 50 100 -20

-10

0

10

20

X=6

U-Mean (m/s)

Z (mm)

0 50 -20

-10

0

10

20

X=9

U-Mean (m/s)

Z (mm)

0 50 100 -20

-10

0

10

20

0 < Red <= 10

10 < Green <= 20

20 < Blue <= Max

X=0

size bins, m: 

U-Mean (m/s)

Z (mm)

0 50 100 -20

-10

0

10

20

X=3

Y (mm)

Z (mm)

-5 0 5 10 -15

-10

-5

0

5

Reference symbols:

= 50 m/s

red = small drops

blue = medium drops

green = large drops
NASA/TM—2010-216365 10 

Figure 13.—Distribution of droplet mean diameter D10, with axial velocity for the high and medium inlet temperature 

conditions. Drops are located in the immediate downstream region of the center injector, within the Y = 0 plane, for 

various axial distances from dump plane. 

Figure 14.—Distribution of droplet mean diameter D10, with axial velocity, at various axial locations within the 

Y = 0 plane at inlet condition L1. Drops are located in the immediate downstream region of the center injector (left), or downstream of the lower injector (right). 

temperature. Despite these factors, the number of drops measured 

at conditions H1 and H2 is comparable, and supports the expectation that the larger fuel pressure drop enhances atomization. 

The drop size-velocity distribution for the lowest inlet 

temperature case, L1, is shown in Figure 14 for the upstream 

(left) and downstream configurations. We would expect to see 

both a higher number of droplets measured and larger drops 

at the upstream location, but instead record many fewer drops 

than at condition M1. The drops also are smaller. These discrepancies from expectation we attribute to the optical thickness and/or large density gradients of the flow field at this 

inlet condition that helped to obscure the signal. 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 3 mm

Avg D = 3.79 m

Counts = 101

10 

D ( m)

Axial Velocity (m/s)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 0 mm

Avg D = 4.40 m

Counts = 93

10  D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 3 mm

Avg D = 5.94 m

Counts = 93

10 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 6 mm

Avg D = 5.02 m

Counts = 33

10 

D ( m)

Axial Velocity (m/s)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 0 mm

Avg D = 6.12 m

Counts = 87

10 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 3 mm

Avg D = 11.92 m

Counts = 3931

10 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 6 mm

Avg D = 10.50 m

Counts = 1120

10 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 9 mm

Avg D = 4.41 m

Counts = 89

 10

D ( m)

Axial Velocity (m/s)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 0 mm

Avg D = 15.46 m

Counts = 4634

10  D ( m)

0 20 40 -100

-50

0

50

100

150

10 

X = 3 mm

Avg D = 6.13 m

Counts = 204

10 

D ( m)

0 20 40 -100

-50

0

50

100

150

10 

X = 5 mm

Avg D = 4.74 m

Counts = 50

10 

D ( m)

0 20 40 -100

-50

0

50

100

150

10 

X = 7 mm

Avg D = 4.60 m

Counts = 33

10 

D ( m)

Axial Velocity (m/s)

0 20 40 -100

-50

0

50

100

150

10 

X = 1 mm

Avg D = 6.64 m

Counts = 39

10 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 3 mm

Avg D = 7.82 m

Counts = 728

10 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 6 mm

Avg D = 5.70 m

Counts = 779

10 

D ( m)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 9 mm

Avg D = 10.92 m

Counts = 1184

10 

D ( m)

Axial Velocity (m/s)

0 20 40 -150

-100

-50

0

50

100

150

10 

X = 0 mm

Avg D = 7.66 m

Counts = 688

10  D ( m) Axial Velocity (m/s)

0 20 40 0

20

40

60

80

100

10 

X = 9 mm

Avg D = 9.39 m

Counts = 7846

10  D ( m)

0 20 40 0

20

40

60

80

100

10 

X = 12 mm

Avg D = 9.90 m

Counts = 3102

10 

D ( m)

0 20 40 0

20

40

60

80

100

10 

X = 15 mm

Avg D = 10.09 m

Counts = 2700

10 

H1 H2 

M2 M1 
NASA/TM—2010-216365 11 

At the downstream location, the drops measured are primarily from the upper half of the lower injector, though there 

is probably contribution from the lower half of the middle 

injector, as well. This location should correspond with an area 

downstream of the recirculation zone, where there may be 

interaction between adjacent injectors. Many more droplets 

were measured at this downstream location, all with positive 

velocity. 

The contour plots in Figure 15 display the influence of 

drop diameter on axial velocities at the downstream locations 

of X = 9-, 12-, and 15-mm. Data represent the L1 case. For 

these positions, only two size bins are used. The top row 

contains the average axial velocities for drops with linear 

diameters between 0- and 10-μm, non-inclusive. The bottom 

row portrays the same information for drop sizes in the range 

of 10- to 20-μm, inclusive. For each size range, the span of 

axial velocities decreases with downstream progression, indicating the flow is approaching a steady, uniform distribution. 

By comparing velocity spans of the top and bottom rows, it 

appears evident that the small drops more closely track the 

actual airflow characteristics during combustion. 

PIV Results 

Figures 16 and 17 present PIV results. The field of view 

used for PIV isolates the center row of injectors. The presented fields are the average of only 250 image pairs, and are 

not necessarily a good statistical sampling. However, limiting 

the number of pairs was convenient for this first attempt at 

PIV in a combusting case, allowing us to consider issues of 

timing, using a shutter, and data storage. Figure 16 shows the 

resultant average droplet velocity and RMS fields from PIV 

for inlet condition M1 at the Y = 0-mm position. For comparison, the PDI average velocity data from the same region is 

plotted on the same scale. The range in RMS for PIV is similar to what we encountered for the individual PDI channels. 

Using the PIV vector plot (center image) we measured a full 

spray angle of approximately 80°, which is consistent with 

the PDI results. 

Taking into account the relatively small PIV sample size, 

the PIV and PDI results compare well and show similar 

trends. One key difference between the PIV and PDI vector 

fields is in the core central region, where the PIV appears 

X = 9-mm X = 12-mm X = 15-mm 

Figure 15.—Results from point L1 at X = 9, 12, and 15-mm from injector exit. Top row: average axial velocities represented 

by contour for droplets between 0 and 10-μm. Bottom row: average axial velocities represented by contour for droplets 

between 10 and 20-μm. Flow direction is out of the page. 

Y (mm)

Z (mm)

-15 -10 -5 0 5 10 15 -30

-25

-20

-15

-10

-5

Y (mm)

Z (mm)

-15 -10 -5 0 5 10 15 -30

-25

-20

-15

-10

-5

Y (mm)

Z (mm)

-15 -10 -5 0 5 10 15 -30

-25

-20

-15

-10

-5 U-Mean (m/s)

70

65

60

55

50

45

40

35

30

25

20

Y (mm)

Z (mm)

-15 -10 -5 0 5 10 15 -30

-25

-20

-15

-10

-5

Y (mm)

Z (mm)

-15 -10 -5 0 5 10 15 -30

-25

-20

-15

-10

-5

Y (mm)

Z (mm)

-15 -10 -5 0 5 10 15 -30

-25

-20

-15

-10

-5 U-Mean (m/s)

70

65

60

55

50

45

40

35

30

25

20
NASA/TM—2010-216365 12 

Figure 16.—Results from PDI (left) and PIV (center and 

right columns) at Y = 0-mm for inlet condition M1. 

sparse. One key advantage of PDI for droplet velocity measurement is that the 30° off-axis forward scattering angle is 

optimal for Mie scattering, whereas the 90° scattering angle 

used for PIV and other planar imaging techniques is less 

efficient. On the other hand, because PIV is a whole-field 

technique, we obtain 2–D snapshots of droplet and velocity 

that would require much more time—and multiple optical 

arrangements—to acquire using PDI. 

Another issue this comparison serves to highlight is the 

optical density problem, mentioned earlier with respect to the 

point L1 PDI data. Figure 17 shows a series of X-plane slices 

from the data block we derived by combining results of all 47 

U–W vector fields. In looking at these end views we see an 

obvious difference between the spray pattern on the camera’s 

near-side (Y ≥ 0) and the opposite side of the flow field. 

Although the distance from the camera to the probe region 

was fixed, the light scattered had to travel through more fluid 

as the distance from the camera increased. Additionally, Mie 

scatter intensity goes as the square of drop size surface area, 

so the signal from the smallest drops may be more obscured 

than from larger drops. Field thickness, obscuration, and 

multiple scattering effects contribute to degrade the initial 

images and the final processed result. 

Nevertheless, this whole field approach of PIV allows us to 

conclude the presence of symmetry in the flow field structure, 

as well as highlight the injector spray pattern using a greater 

field of view than that attainable through PDI. The images in 

Figure 17 tell us that the adjacent fuel sprays begin mixing by 

X = 3-mm and are fully involved by 5-mm. The farthest 

extent of the spray (most drops have vaporized) we found to 

be around 11-mm at this inlet condition, which is consistent 

with the range we found using PDI. 

X  1-mm

X  2.5-mm

X  5-mm

Figure 17.—Axial velocity contours at selected axial 

planes obtained via PIV, inlet condition M1. Flow 

out of page.

Summary 

To aid in the process of building the next generation of aircraft powerplants that will meet the increasingly stringent 

restrictions on emissions, researchers at the NASA Glenn 

Research Center have been investigating a novel 9-point lean 

direct injector. In order to meet the challenges involved in 

taking this injector from test lab to commercial aircraft 

engine, a more thorough understanding of the chemical as 

well as physical processes occurring within the flowfield of 
NASA/TM—2010-216365 13 

this injector is required. Recent advances in optical diagnostic 

methods offer the greatest opportunity to provide this information. Given our facility and its available optical access, 

PDI and PIV were chosen as complementary candidate techniques to fully describe parametrically the flowfield of this 

injector. 

Our results demonstrate that PDI was successful at locating 

the core recirculation zones and at determining the injector 

full cone spray angle which was found to be roughly 80o in 

each case. PDI was also particularly useful in ascertaining 

droplet size, local turbulence levels and velocity vector components. Average droplet diameters, D10, were typically less 

than 20-μm. Mapping the droplet horizontal and vertical 

velocity vectors showed the clock-wise circulation (in the 

stream-wise sense) that the 60o air swirlers imparted to the 

flow. The droplet diameter data was also significant in showing that the 9-point LDI hardware is highly effective at atomizing the Jet-A fuel and showing that all but the largest 

droplets were completely vaporized within the first 15-mm 

from the injector exit. PDI measurements revealed that the 

highest velocities occurred where the jets from adjacent injectors mix, while the lowest velocities were found within each 

injector’s hollow fuel cone. Local turbulence levels were also 

shown to be at a maximum where adjacent fuel jets mixed 

and at a minimum in the boundary region of the spray cones. 

PIV results, like those of PDI, also revealed an 80o full cone 

spray angle and yielded similar results for average droplet 

velocity and RMS fields. Both techniques were also found to 

be similarly affected by the optical thickness of the region 

being examined with the amount of data being reduced as the 

optical thickness at the measurement location increased. 

Future work will include testing to better assess interactions between LDI elements. Plans include testing both 

seeded cold (unfueled) and combusting cases at the same air 

inlet conditions. Plans also include testing with fuel to the 

center element only without fuel to neighbors, which will enable 

us to more readily control boundary conditions so that we may 

directly compare single and multiple injector configurations. 

References 

1. Rizk, Nader K., 2004, “Fuel Atomization Effects on Combustor Performance,” AIAA–2004–3540. 

2. Penner, J., Lister, D., Griggs, D., Dokken, D., and McFarland, M., “Regulatory and Market-Based Mitigation 

Measures,” Aviation and the Global Atmosphere: A Special Report of IPCC Working Groups, Cambridge University Press, Cambridge, UK, 1999, pp. 335–348. 

3. Tacina, R., Wey, C, Laing, P., and Mansour, A., 2002, 

“Sector Tests of a Low-NOx, Lean Direct Injection, Multipoint Integrated Module Combustor Concept,” ASME 

Paper No. GT–2002–30089. 

4. Tacina, R., Lee, P., and Wey, C., 2005, “A Lean-DirectInjection Combustor Using a 9 Point Swirl-Venturi Fuel 

Injector,” ISABE Paper No. ISABE–2005–1106. 

5. Hicks, Y.R., Locke, R.J., Wey, C.C., and Bianco, J., 1993, 

“A Unique Optically Accessible Flame Tube for Lean 

Combustor Studies,” NASA/TM–107073. 

6. Davoudzadeh, F., Liu, N.-S., and Moder, J.P., 2006, 

“Investigations of Swirling Air Flows Generated by Axial 

Swirlers in a Flame Tube,” ASME Paper No. GT–2006–

91300. 

7. Hicks, Y.R., Locke, R.J., and Anderson, R.C., 2007, “Optical Measurements in a Combustor Using a 9-Point SwirlVenturi Fuel Injector,” ISABE Paper No. 2007–1280. 

8. Cai, J., Jeng, S.-M., and Tacina, R., 2005, “The Structure 

of a Swirl-Stabilized Reacting Spray Issued From an Axial 

Swirler,” AIAA–2005–1424. 
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Optical Characterization of a Multipoint Lean Direct Injector for Gas Turbine Combustors: 

Velocity and Fuel Drop Size Measurements 

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6. AUTHOR(S)

Heath, Christopher, M.; Anderson, Robert, C.; Locke, Randy, J.; Hicks, Yolanda, R. 

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WBS 561581.02.08.03.16.03 

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John H. Glenn Research Center at Lewis Field 

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This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Performance of a multipoint, lean direct injection (MP-LDI) strategy for low emission aero-propulsion systems has been tested in a Jet-A 

fueled, lean flame tube combustion rig. Operating conditions for the series of tests included inlet air temperatures between 672 and 828 K, 

pressures between 1034 and 1379 kPa and total equivalence ratios between 0.41 and 0.45, resulting in equilibrium flame temperatures 

approaching 1800 K. Ranges of operation were selected to represent the spectrum of subsonic and supersonic flight conditions projected for 

the next-generation of commercial aircraft. This document reports laser-based measurements of in situ fuel velocities and fuel drop sizes for 

the NASA 9-point LDI hardware arranged in a 3×3 square grid configuration. Data obtained represent a region of the flame tube combustor 

with optical access that extends 38.1-mm downstream of the fuel injection site. All data were obtained within reacting flows, without 

particle seeding. Two diagnostic methods were employed to evaluate the resulting flow path. Three-component velocity fields have been 

captured using phase Doppler interferometry (PDI), and two-component velocity distributions using planar particle image velocimetry 

(PIV). Data from these techniques have also offered insight into fuel drop size and distribution, fuel injector spray angle and pattern, 

turbulence intensity, degree of vaporization and extent of reaction. This research serves to characterize operation of the baseline NASA 9-

point LDI strategy for potential use in future gas-turbine combustor applications. An additional motive is the compilation of a

comprehensive database to facilitate understanding of combustor fuel injector aerodynamics and fuel vaporization processes, which in turn 

may be used to validate computational fluid dynamics codes, such as the National Combustor Code (NCC), among others. 

15. SUBJECT TERMS

Laser diagnostics; Fuel injection; Fuel-air mixing; Aircraft gas turbine combustion; Velocity 

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