Yolanda R. Hicks, Sarah A. Tedder, Robert C. Anderson, and Anthony C. Iannetti

Glenn Research Center, Cleveland, Ohio

Lance L. Smith and Zhongtao Dai

United Technologies Research Center, East Hartford, Connecticut

NASA Numerical and Experimental Evaluation of

UTRC Low Emissions Injector

NASA/TM—2014-218493

September 2014
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 443–757–5803

• Phone the NASA STI Information Desk at

 443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Yolanda R. Hicks, Sarah A. Tedder, Robert C. Anderson, and Anthony C. Iannetti

Glenn Research Center, Cleveland, Ohio

Lance L. Smith and Zhongtao Dai

United Technologies Research Center, East Hartford, Connecticut

NASA Numerical and Experimental Evaluation of

UTRC Low Emissions Injector

NASA/TM—2014-218493

September 2014

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

50th Joint Propulsion Conference

cosponsored by the AIAA, ASME, SAE, and ASEE

Cleveland, Ohio, July 28–30, 2014
Acknowledgments

This work was supported by the Supersonics Project of the NASA Fundamental Aeronautics Program and by the NASA 

Environmentally Responsible Aviation Program. We also thank Changlie Wey and Randy Locke for their assistance in conducting 

the experiments.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 

This report is a formal draft or working 

paper, intended to solicit comments and 

ideas from a technical peer group.
NASA/TM—2014-218493 1 

NASA Numerical and Experimental Evaluation of 

UTRC Low Emissions Injector 

Yolanda R. Hicks, Sarah A. Tedder, Robert C. Anderson, and Anthony C. Iannetti 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Lance L. Smith and Zhongtao Dai 

United Technologies Research Center 

East Hartford, Connecticut 06108 

Abstract 

Computational and experimental analyses of a PICS—Pilot-In-Can-Swirler technology injector, 

developed by United Technologies Research Center (UTRC) are presented. NASA has defined 

technology targets for near term (called “N+1”, circa 2015), midterm (“N+2”, circa 2020) and far term 

(“N+3”, circa 2030) that specify realistic emissions and fuel efficiency goals for commercial aircraft. This 

injector has potential for application in an engine to meet the Pratt & Whitney N+3 supersonic cycle 

goals, or the subsonic N+2 engine cycle goals. Experimental methods were employed to investigate 

supersonic cruise points as well as select points of the subsonic cycle engine; cruise, approach, and idle 

with a slightly elevated inlet pressure. 

Experiments at NASA employed gas analysis and a suite of laser-based measurement techniques to 

characterize the combustor flow downstream from the PICS dump plane. Optical diagnostics employed 

for this work included Planar Laser-Induced Fluorescence of fuel for injector spray pattern and 

Spontaneous Raman Spectroscopy for relative species concentration of fuel and CO2. 

The work reported here used unheated (liquid) Jet-A fuel for all fuel circuits and cycle conditions. 

The initial tests performed by UTRC used vaporized Jet-A to simulate the expected supersonic cruise 

condition, which anticipated using fuel as a heat sink. 

Using the National Combustion Code a PICS-based combustor was modeled with liquid fuel at the 

supersonic cruise condition. All CFD models used a cubic non-linear k-epsilon turbulence wall functions 

model, and a semi-detailed Jet-A kinetic mechanism based on a surrogate fuel mixture. Two initial spray 

droplet size distribution and spray cone conditions were used: (1) an initial condition (Lefebvre) with an 

assumed Rosin-Rammler distribution, and 7 degree Solid Spray Cone; and (2) the Boundary Layer 

Stripping (BLS) primary atomization model giving the spray size distribution and directional properties. 

Contour and line plots are shown in comparison with experimental data (where this data is available) for 

flow velocities, fuel, and temperature distribution. The CFD results are consistent with experimental 

observations for fuel distribution and vaporization. 

Analysis of gas sample results, using a previously-developed NASA NOx correlation, indicates that 

for sea-level takeoff, the PICS configuration is predicted to deliver an EINOx value of about three for the 

targeted supersonic aircraft. Emissions results at supersonic cruise conditions show potential for meeting 

the NASA goals with liquid fuel. 

Nomenclature 

AST Advanced Subsonic Technology 

CFD computational fluid dynamics 

CFL Courant-Friedrichs-Lewy condition 

Cp specific heat 
NASA/TM—2014-218493 2 

CPU central processing unit 

 turbulent dissipation 

EINOX emission index for oxides of nitrogen 

ERA environmentally responsible aviation 

FAR fuel-to-air ratio 

FWHM full width at half maximum 

JST Jameson-Schmidt-Turkel dissipation scheme 

k turbulent kinetic energy 

LTO landing-takeoff cycle 

 molecular vibrational energy 

NCC National Combustion Code 

PICS Pilot-In-Can-Swirler 

P3 combustor inlet pressure 

PLIF planar laser-induced fluorescence 

PLS planar laser scatter 

T3 combustor inlet temperature 

SLTO sea level takeoff 

SRS spontaneous Raman spectroscopy 

Vdrop liquid phase droplet group velocity vector 

Vgas gas phase velocity vector 

Vgas turbulent fluctuations of the gas phase velocity vector 

Introduction 

For more than 40 years, NASA has sustained programs to reduce the environmental effects of 

aviation. A major focus of these programs has been reducing the emissions of nitrogen oxides (NOx). 

NOx emissions decrease the protective ozone layer in the stratosphere and increase smog and ozone in the 

lower troposphere (Ref. 1). To prevent damage to the protective ozone layer, NASA programs have 

focused on reducing NOx emissions at cruise conditions for supersonic flight. To reduce the emissions of 

NOx in the lower troposphere, NASA programs have also focused on reducing NOx emissions during the 

landing-takeoff cycle in subsonic flight. 

In addition to decreasing NOx emissions, NASA has also addressed the reduction of carbon dioxide 

emissions by increasing engine fuel efficiency. Increased fuel efficiency is typically achieved by 

increasing the engine operating pressure ratio, which increases combustor inlet temperature; however, 

NOx formation rates increase with higher temperatures. Without an improvement in combustor 

technology, higher efficiency engines will have higher, not lower, NOx emissions: improved low-NOx 

combustor technologies need to be developed. NASA has defined technology targets for the near term 

(called “N+1”, circa 2015), midterm (“N+2”, circa 2020) and far term (“N+3”, circa 2030) that specify 

realistic emissions and fuel efficiency goals for commercial aircraft. In the far term, the emissions goal for 

supersonic cruise is less than 5 g/kg of fuel burned, with fuel efficiency targeted at between 3.5 to 4.5 

passenger-miles per pound of fuel (Ref. 2). The Supersonics Project of the NASA Fundamental 

Aeronautics Program is charged with managing the challenge to develop the technologies required to 

meet these goals. The technical approach taken incorporates using physics-based modeling for 

combustion CFD to improve supersonic cruise emissions predictions, and using validation experiments to 

assure the models work. These models are applied to promising low emission concepts to aid in design 

and development. NASA has partnered with engine and fuel injector manufacturers to develop practical 

technology to meet the stringent emissions goals requirements. 
NASA/TM—2014-218493 3 

One such partnership is with United Technologies Research Center (UTRC), which devised a fuel 

injection/mixing concept to meet the far term criteria using a fuel-lean approach compatible with the 

cycle conditions devised for a Pratt & Whitney supersonic transport notional engine. The concept is called 

the Pilot-In-Can Swirler (PICS). In the PICS injector concept, each swirler contains a pilot “can” 

concentrically located inside the main stage swirler (which dumps into an annulus), and in which the pilot 

stage flame is isolated from the main stage flame to help reduce interaction between zones for improved 

turndown capability. One key aspect of the concept is that the fuel for the main fuel stage is injected as a 

gas (Ref. 3). Gas phase fuel air mixing theoretically has the benefit of reducing the additional mixing time 

that would have been required for fuel atomization and vaporization. Additionally, if the fuel is able to be 

preheated by using it as a heat sink for cooling the aircraft, other efficiencies can be gained. 

Previous evaluation of the PICS injector performed at UTRC (Ref. 3) included comparison of 

cold-flow gas-gas mixing measurements with analytical predictions for supersonic cruise conditions, and 

combustion tests at both supersonic cruise and idle conditions (pilot only) using vaporized Jet-A fuel for 

the main stage and non-heated, liquid Jet-A for the pilot. The gas-gas mixing results showed acceptable 

variation in mixture fraction at the primary swirler exit. The emissions results at supersonic cruise 

satisfied the NOx emissions target with combustion efficiency  99.99 percent, and at idle conditions also 

had high combustion efficiency with EINOx  6. 

This paper reports results from CFD analysis and combustion testing conducted at the NASA Glenn 

Research Center, to explore PICS injector performance when using liquid (not gaseous) fuel. Testing of 

the PICS injector at NASA used unheated liquid fuel for all fuel circuits and all power settings, and 

expanded the previous test matrix. For these liquid-fuel tests at NASA, UTRC designed and fabricated a 

PICS injector having a reduced number of main fuel-injection orifices as compared to the earlier 

vaporized-fuel design. The smaller number of injection orifices ensured adequate fuel-injection velocity 

when using liquid fuel, which has a lower volume flow rate than vaporized fuel because liquid is more 

dense than vapor. Optical diagnostics techniques were applied at the combustor inlet to examine injector 

performance for fuel liquid versus vapor distribution and mixedness, and exhaust gas analysis was applied 

at the combustor exit to measure emissions. The test results are compared with CFD results produced 

using the National Combustor Code (NCC), specifically at supersonic cruise conditions. 

PICS Hardware 

The PICS injector concept, illustrated in Figure 1, consists of a pilot stage and a main stage, with 

turndown controlled by shifting the fuel distribution between the stages. To achieve this, the injector 

contains a pilot ”can” centrally located within the main swirler, in which the pilot flame is isolated from 

the main-stage flame. When installed within a single-annular combustor dome, each injector is identical 

and receives the same fuel flow allocation. All tests and computations at NASA used a single PICS 

swirler, installed in a single 4- by 4-in. sector. The single PICS injector hardware is shown in 

Figure 2 and the computational envelope used for CFD is shown in Figure 3. 

Figure 1.—PICS swirler-injector 

concept. 

Figure 2.—Photograph of the 

single PICS injector, from 

aft-looking-forward view. 

Figure 3.—Computation envelope 

used by the National Combustion 

Code for CFD simulations of PICS. 
NASA/TM—2014-218493 4 

Experiment Setup 

PICS Installation 

The combustor subcomponent facility at NASA Glenn supplies heated, non-vitiated air to the test rig. 

The test rig uses a castable ceramic to form the combustor “liner”, so the four inch by four inch PICS 

module had adiabatic walls, rather than a water-cooled liner as used for the UTRC tests. Another 

difference in the NASA installation was that the test stand on which the PICS injector was installed is 

best suited for higher air and fuel flow rates. Because the two facilities are designed to accommodate 

different flows, the two sets of tests could not completely match the Pratt & Whitney supersonic cycle 

conditions explored in this study. Table 1 shows the Pratt & Whitney supersonic N+3 cycle conditions. 

Table 2 those lists the conditions used at NASA Glenn that incorporate the laser diagnostics 

measurements. 

Laser and Optical Diagnostics 

To determine the fuel spray pattern, we acquired two-dimensional images of fuel via planar laserinduced fluorescence (PLIF) and planar laser scatter (PLS); to measure major combustion species (N2, O2, 

CO, CO2, fuel, and H2O), we acquired one-dimensional images using Spontaneous Raman spectroscopy 

(SRS). We also acquired chemiluminescence images of Swan band C2

*

., which are visible in the reaction 

zone and show where carbon-carbon electronic transitions occur. (CH*

 and C2

*

 excited species emit light 

in the visible region of a hydrocarbon flame spectrum.) Setup and data acquisition are described next. 

The fuel PLIF, PLS, and C2

*

 images were obtained using the same receiving optics and intensified 

CCD (ICCD) camera. A second set of fuel PLIF images was gathered using a second ICCD camera, 

located on the opposite side of the combustor. The setup is similar to that illustrated in Figure 4, in which 

the cameras were positioned on the sides, to image light from an angle perpendicular to the combustor 

flow direction. A remotely controlled filter wheel allowed us to select the appropriate filter (FWHM of 

10-nm, typical) to pass light through a UV-grade, f = 105-mm, f/4.5, macro camera lens. The light was 

then focused onto a gated, 16-bit, 1k  1k pixel array, ICCD camera having a Gen II Super-Blue-SlowGate intensifier. The gate time used was 100-ns. The laser used was a 10-Hz, frequency-doubled 

Nd:YAG-pumped dye laser/frequency mixer system set up to produce wavelengths around 282-nm. The 

laser beam was formed into a sheet using a pair of cylindrical lenses, to obtain a sheet approximately 

300 m thick. We typically acquired images with on-chip averages of 200 gates, and traversed across the 

flow along the Y axis in 1-mm increments. The laser sheet and collection optics were traversed together 

so as to maintain focus on the laser sheet. For PLIF, we collected fluorescence using a filter centered near 

334-nm on one camera and near 313-nm on the other. For the elastically-scattered light of PLS, we used a 

filter centered near 280-nm for each camera. For chemiluminescence, we did not use the laser and 

collected light when positioned at Y = 0 using a filter centered at 514-nm. We collected PLIF, PLS and 

C2

*

 data for all test matrix conditions. 

TABLE 1.—PRATT & WHITNEY SUPERSONIC N+3 CYCLE (REF. 3) 

 P3, psi T3, °F FAR/FARSLTO

Supersonic cruise 50,000 ft, M1.8 174 1087 1.10 

SLTO 329 890 1.00 

Climb-66% 235 767 0.82 

Approach-32% 149 634 0.66 

idle 78 475 0.58 

TABLE 2.—INLET CONDTIONS USED FOR PICS TESTS 

WITH LASER DIAGNOSTICS 

Nominal cycle P3, psi T3, °F FAR/FARSLTO

Supersonic cruise, N+3 174 975 1.24 

~ Subsonic cruise, N+2 250 1000 0.85 

Approach, subsonic, N+2 205 716 0.75 

~ Idle, subsonic, N+2 100 425 0.85 
NASA/TM—2014-218493 5 

Figure 4.—Schematic drawing that shows the key 

components used for planar laser-induced 

fluorescence and planer laser scattering. 

Figure 5.—llustration that shows the key components 

used for spontaneous Raman scattering experiments. 

The spontaneous Raman species measurements were obtained using an optical arrangement similar to 

that shown in Figure 5. In this case, 532-nm light from a frequency-doubled Nd:YAG laser operated at 

15-Hz provided the light-scattering source. The probe volume was formed using a spherical lens, to 

provide a narrow, high-energy-density region from which molecules can scatter light (the SRS signal is 

~10-6 smaller than for laser induced fluorescence). The shape of the probe volume was defined by a 

combination of input and collection optics which produces a cylinder 6.67-mm high—aligned in the 

vertical (z-)—with approximately a 2-mm diameter. The scattered light from the molecules in the probe 

volume passed through a 550-nm long pass filter and was collected using an f = 60-mm spherical lens 

coupled to a vertically-oriented, linear fiber array, consisting of thirty-seven, 200-m-diameter fibers. The 

fiber transmitted the light to an f/4, f = 300-mm spectrometer with a 600 groove/mm grating blazed at 

500 nm. The light dispersed by the grating was focused onto a High Q Blue, gated, intensified camera 

having an array size 1024-pixel-wide by 256-pixel- high. The camera CCD pixel size combined with the 

600 line/mm grating produced a ~0.1667 nm/pixel spectral resolution. The camera array was binned in 
NASA/TM—2014-218493 6 

the spatial (vertical) dimension into five rows, each 40 pixels high, with the remaining pixels not used. 

These five rows provide for simultaneous measurement at five discrete locations in the z-direction. Sets of 

100, 10-gate on-chip averages were collected at each position traversed in x and y. The five binned rows 

created five spectra in which the peaks were used to identify the relative species concentrations. The use 

of Raman was restricted to those inlet conditions expected to have little or no liquid fuel so that the 

interference from Mie scatter is minimized; Mie scatter has a signal roughly a million times greater than 

SRS. 

For all measurements, there may be a slight difference in the true spatial coordinates because the 

heated inlet air causes the test rig will to shift axially (x-), horizontally (y-) and rotationally about its 

centerline. We took measures during the tests to account for translations, but the rotation was hard to 

quantify and not addressed. 

PICS Modeling 

The National Combustion Code 

The National Combustion Code (NCC) is a state of the art CFD program specifically designed for 

combustion processes. A short summary of the features of the NCC pertaining to this paper are: the use of 

unstructured grids (Ref. 4), massively parallel computing—with almost perfectly linear scalability 

(Refs. 5 and 6) on non-spray cases up to four thousand central processing units (CPU), a dynamic wall 

function with the effect of adverse pressure gradient (Ref. 7), low Reynolds number wall treatment 

(Ref. 8), and a cubic non-linear k-epsilon turbulence model (Refs. 9 and 10), lagrangian liquid phase 

spray model, and stiff laminar chemistry integration. Recently, viscous low-speed preconditioning 

(Refs. 12 and 13) has been added to improve the low-speed convergence of the NCC in viscous regions. 

The combination of these features is usually not available in other CFD codes and gives the NCC an 

advantage when computing recirculating, turbulent, reacting, spray flows. Previously, the NCC has 

undergone extensive validation studies for simple flows (Ref. 14), complex flows (Ref. 15), NOx

emissions prediction performance (Ref. 16), and traditional gas turbine combustor/injectors (Ref. 17). 

Geometry and Mesh Generation 

UTRC created the orginial PICS solid model using NX (Ref. 18) and transferred it to NASA in 

ParaSolid format as an “air-solid”, the solid geometry that only represented the fluid flowing through the 

PICS combustor and NASA test cell. The file was then imported into SolidWorks, “regenerated” and 

exported as an ACIS (Spatial Corporation, Broomfield, CO) file. The ACIS file was then imported into 

Cubit (Ref. 19) for automatic geometry clean up and manual simplification, while also removing surface 

artifacts (like holes where tubes once existed) from the simplification process. 

The single, simplified “air-solid” was then imported as an IGES file into Pointwise (Ref. 20). Because 

ignition transients tended to greatly increase wall clock computational time as mesh size increased, an 

optimized mesh of 5.5 million isotropic tetrahedral elements was generated. We believe this mesh balanced 

accuracy and reasonable solution time for this engineering effort. Non-reacting simulations were run as a 

check using a stretched tetrahedral mesh of 18 million elements. The 5.5 million element matched the global 

pressure drop to within 20 percent of the 18 million element mesh value. 

Chemistry Modeling 

Ideally, we would prefer to use detailed chemical kinetic models. There are two problems with this 

approach: (1) Jet-A is a multi-component fuel and not a substance, and there are no universally accepted 

surrogate fuel models for Jet-A; (2) the computational costs associated with these models make them 

impractical when fine computational grids are used. Originally, a single-step, global chemistry model was 

used. This model was based on propane kinetics (Ref. 21), which are close to Jet-A’s reaction rates. The 

Jet-A fuel is modeled as single species (C11H21) that represents a hydrocarbon mixture of decane, hexane 
NASA/TM—2014-218493 7 

and benzene in both gas-phase and liquid-phase solvers. The gas is treated as an ideal mixture with 

five-coefficient curve fits for Cp(T) of each species and a CHEMKIN treatment of transport properties 

(species and ideal mixture rules). The single-step model allowed an easier start up in the solution process, 

by reducing the computational requirements during the ignition phase. Single-step models do not allow 

emissions calculations, only heat release. Because of this, a reduced twenty-step, fifteen species model 

was used based on the mechanism published in Ajmani-Kundu (Ref. 22). The reduced mechanism also 

describes the formation of Carbon Monoxide and Nitrogen Oxide. However, only one nitrogen-oxide 

species, NO, has been used in the reduced mechanism. NO in the reduced mechanism represents the 

whole family of nitrogen oxides including nitric oxide by Zeldovich (Ref. 23) reactions, prompt NO 

reactions by Fenimore (Ref. 24), and nitrogen oxide formation through nitrous oxide. 

Liquid Phase Modeling 

The liquid Jet-A spray is evolved using a dilute spray Lagrangian solver (LSPRAY) (Ref. 4) which 

neglects any drop-drop interactions. Turbulence effects follow the KIVA-II approach (Ref. 25) of adding 

a turbulent fluctuation velocity to gas-phase velocity when calculating droplet drag and vaporization. The 

evaporation model includes solution of internal drop temperature distribution (thirteen point onedimensional mesh, finite-difference solution of a Hill vortex model) and a droplet regression rate 

employing three different correlations depending on droplet Reynolds number, Redrop (where Redrop is 

based on the relative speed Vgas + V

gas – Vdrop). 

Jet-A fuel was represented as a single component, C11H21. Fuel injection, atomization, and 

evaporation were handled in the following manner. Fuel injection in the model was handled by using 

Lefebvre’s (Ref. 26) airblast correlations for a plain orifice. The correlation provided mean fuel droplets 

sizes of 16.3μm for the main fuel circuit and 13 μm for the pilot. These particles are then injected, 

assuming a Rosin Rammler drop size distribution (Ref. 27), in the appropriate locations in the model, as 

jets in crossflow. These jets were presumed to have a solid cone angle of 7°, with each cone represented 

by eight circumferential droplets groups, with eight polar droplets groups at each circumferential location, 

stochasitally injected (with respect to polar and circumferential locations) every 200 spray solver time 

steps (40 sec). The main and pilot circuit drops were injected with velocities of 42 and 13 m/s, 

respectively. Once in the flow, the particles were tracked 0.2 msec before evaporation was switched on. 

Computational Procedure 

The simulations are steady-state and solve the Favre-averaged transport equations for species, 

momentum and energy, as well as a two-equation k- turbulence model. Staging was used in the solution 

process; cold-flow calculations and initial combustion calculations were performed using a single-step 

chemistry model with lagrangian spray until a steady state solution was obtained. The final stage of CFD 

calculations was performed by switching from the one-step chemistry model to the reduced chemistry 

model; this was done by changing the input chemistry-parameters of the code. It is important to note that 

no turbulence–chemistry interaction model was used for this case, so called “laminar chemistry”. So, 

averaged temperature and averaged mass fractions were used to compute the reaction rate. We believe 

this is appropriate as an engineering assumption, because for this particular case, turbulent kinetic energy 

(k) is below 20 m2

/s2, with the peak on occurring at the tip of the fuel injector. 

Calculation of the source term due to chemical kinetics uses the explicit “reference species” approach 

described in the KIVA-II manual. In the current implementation, the explicit “reference species” 

integration of chemical kinetics source is performed using ten sub-iterations within each pseudo-time step 

of gas-phase Runge-Kutta integration in pseudo-time. 

The NCC computations for reacting and non-reacting flow were run in general until the flow residuals 

were reduced three orders of magnitude. The mass flow rates at the boundary conditions were also 

monitored as a convergence criterion. Dissipation (JST type) was set at 0.0 for second order dissipation 

(2) and 0.05 for fourth order dissipation (4) (Ref. 28). The value of k2, the constant that scales the second 
NASA/TM—2014-218493 8 

order dissipation gradient switch, was set at 0.70. Setting the second order dissipation to zero is absolutely 

necessary to accurately resolve flow features, like jets. A CFL number of 1.0 was used. A cubic, nonlinear k-epsilon model with a variable Cmu coefficient was used. This model was selected because of the 

swirling flow. A dynamic wall function with pressure gradient effects was used to model near wall 

turbulent flow effects. 

Computations were performed on the SGI ICE computer “Pleiades” at NASA Ames. The “Pleiades” 

supercomputer with “Ivy Bridge” processors was preferred because its processor cores had considerably 

higher computational performance and because of its high speed, low-latency interconnect. This 

interconnect was important because the Lagrangian spray model created a load unbalance in compute 

nodes. The high speed interconnect seemed to mitigate the load imbalance. It takes approximately one 

week (wall clock) to complete a single PICS combustion case (starting from no initial solution) using 

960 processors. 

Results and Discussion 

Along with the emissions results obtained by gas sampling at the combustor exit, we present the CFD 

results for supersonic cruise and compare those with measured fuel patternation and Raman fuel and CO2

concentrations relative to nitrogen. We also compare cases related to the ERA subsonic N+2 cycle for 

approach and elevated “idle” conditions, for fuel patternation via PLIF and PLS only. 

Supersonic Cruise 

Figures 6 to 9 show highlights of the CFD results at the N+3 supersonic cruise conditions run at 

NASA Glenn using fuel at ambient temperature. Later, we compare the simulated results with the laserbased measurements. The mean velocities and vorticity are shown in Figure 6, while Figure 7 shows the 

temperature distribution. Figures 8 and 9 show the OH and fuel profiles, respectively. 

Figure 6.—CFD results: normalized velocity and vorticity for N+3 supersonic cruise 

at center of combustor. Flow is from left to right. a) total velocity; b) axial velocity; 

c) radial velocity; d) vorticity. 
NASA/TM—2014-218493 9 

Figure 7.—CFD results: Predicted contours of 

temperature within the Y = 0 plane of the PICS 

combustor at the N+3 supersonic cruise condition. 

Flow passes from left to right. 

Figure 8.—CFD results: Predicted contours of OH mole 

fraction within the Y = 0 plane of the PICS combustor 

at the N+3 supersonic cruise condition. Flow passes 

from left to right. 

Figure 9.—CFD results: Predicted contours of fuel mole 

fraction within the Y = 0 plane of the PICS combustor 

at the N+3 supersonic cruise condition. Flow passes 

from left to right. 

The PICS generates high radial velocity and vorticity within the swirler, which translates to a strong 

central recirculation zone immediately downstream from the swirler to help promote flame stability. The 

highest temperatures, from about 2200 to 2400 K extend as far downstream as the recirculation zone and 

are bounded radially by the mid-level vorticity (with values on the order of 5500 s–1). 

Figures 8 and 9 show the predictions for OH and fuel. We see high levels of OH—which show areas 

of reactivity and temperature—in the regions where fuel and air mix, so the regions of Jet-A (as 

represented by C11H21) coincide with high OH. Most of the OH occurs within the recirculation zone, but 

secondary and tertiary reactions persist downstream until reactions are complete. Fuel from the main 

stage projects downstream approximately to the axial location coinciding with the maximum reverse flow 

velocity and radially outside of the recirculation zone. This allows us to compare the fuel spray as 

measured using PLIF, PLS, and SRS to the CFD analysis. 

 
NASA/TM—2014-218493 10 

The optical measurements occurred beyond the PICS “dump plane” which is the injector’s 

downstream-most vertical end surface. Figures 10 to 15 show the results obtained at supersonic cruise 

conditions. The PLIF signal—representing fuel liquid and vapor—arises from the family of compounds 

having double ring aromatics (naphthalene—C10H8), as its base, such as methylnaphthalene. PLS signal 

arises from the Mie-type scatter presumedly of liquid phase fuel droplets on the order of 280 nm and 

larger. The SRS fuel signal is derived from the spectral region of the main C—H stretch, having a Raman 

shift near 3015 cm–1, near methane 1 and 2 and ethylene 1 vibrational energies. The SRS signals for 

fuel and CO2 are presented as mole fractions, normalized to molecular nitrogen. The PLIF and PLS 

signals are measured in counts of signal and presented on a linear scale. The signals are scaled separately 

to ease comparison. 

Figure 10.—PICS Fuel pattern in an axial slice at the 

supersonic cruise condition: CFD prediction as line 

contour, with measured PLIF as color contour. Aftlooking-forward perspective. 

Figure 11.—Comparison at supersonic cruise condition 

between predicted Jet-A (line contour) and measured 

fuel PLIF (color contour) distribution in the center 

vertical plane, at Y = 0. Flow is from left to right. 

Figure 12.—Supersonic cruise result: Comparison 

between total fuel signal represented by Fuel PLIF 

(red-yellow contour) and liquid fuel from PLS (blue 

contour). Perspective view is aft-looking-forward. 

Figure 13.—Supersonic crise result: Comparison 

showing the decay of liquid fuel (laser scatter, 

dashed, blue line) to Fuel PLIF (liquid+vapor, solid 

red line) at supersonic cruise. Each plot is normalized 

using its maximum signal. 

 
NASA/TM—2014-218493 11 

Figure 14.—Fuel fraction contours determined using CFD (grayscale, with lines) and SRS (color) at supersonic 

cruise condition. a) and b) are presented as axial slices, from the perspective of aft-looking-forward, with a) near 

the dump plane and b) 10 mm downstream. The 14c) image shows a center vertical slice with flow from left to right. 

Figure 15.—Predicted and measured CO2 signals within two axial slices, supersonic cruise condition. 

a) near the PICS dump plane, and b) 10 mm downstream. CFD results are presented as labeled, gray 

scale contours and the SRS results are scatter plots, wit size represented by color. The perspective is 

aft-looking-forward. 

Figures 10, 11, and 14 show the fuel pattern for PLIF or SRS, each overlaid with CFD results at the 

corresponding spatial locations. The PLIF and PLS images were obtained on the same day, within 

minutes of each other; the SRS data were obtained on a different day. In each case, fuel signal emanates 

from a different “species”. Results from the optical techniques compare well with the numerical 

predictions, and show the fuel-air mixture is not uniform, which is to be expected because the fuel was 

injected at room temperature. There was not enough time for complete vaporization and mixing before 

exiting the main swirler. The amount of liquid compared to vapor can be seen qualitatively in Figure 12, 

which shows the fuel PLIF signal (in red-yellow tones), comprised from both liquid and vapor signals; 

along with the PLS signal (blue tones), which is from the liquid fuel. Of the cases tested, this supersonic 

cruise point exhibited the highest signal from both PLIF and PLS. The two signals overlap almost totally 

to show that the fuel at this axial location is mostly in the liquid phase. Figure 13 shows the decay of fuel 

signal with axial distance from the dump plane for PLIF( solid red line) and PLS (dashed blue line). The 

mean signal was computed by calculating the average total signal at each axial location within the data 
NASA/TM—2014-218493 12 

blocks, then normalizing based on the respective maximum signal. The liquid signal decays at a faster rate 

than the total signal. We assume the fuel is effectively vaporized when the PLS signal reaches 5 percent 

of the maximum. Using this assumption, the fuel is evaporated an axial distance approximately 60 percent 

of the distance it takes the total remaining fuel to be consumed. 

Figures 11 and 14(c) show, respectively, the fuel PLIF and SRS signals, along with the CFD fuel in the 

y = 0 plane, just downstream from the PICS dump plane. One may notice in Figure 11 that the fuel PLIF 

signal contour is flat at the top; this marks the top of the optical field-of-view. All show similar results in 

that the fuel is projected outward and downstream from its origin. The slight displacement between 

measurements might be for a number of reasons. For example, the inlet conditions can vary slightly from 

one test day to the next. Setup of the two imaging techniques uses different equipment and necessarily 

different fiducials to designate a home reference position for the optics. SRS measurement results may also 

appear displaced because of the low spatial resolution of the measurement volume compared to the CFD or 

PLIF resolution; thus for example, the SRS fuel signal in Figure 14(c) appears somewhat lower compared to 

the CFD. Also noted earlier, the test rig shifts depending on the inlet conditions, and some misalignment 

may occur based on rig growth. Although the CFD requires compromise in choices of models, mesh, and 

computational procedure, among others, it represents a physical ideal, while the research hardware, in 

particular, is likely to have some machining non-uniformities or imperfections and undergoes wear and tear 

for each test installation and as time progresses. 

Figure 15 shows the results for carbon dioxide. The CFD mass fractions are shown as grayscale line 

contours, with the contours labeled. The SRS data are shown as a color scatter plot, with the symbols 

sized based on the concentrations relative to molecular nitrogen. From the CFD, we see that CO2

concentration peaks in two regions: centrally, immediately downstream from the pilot; and in the annular 

region downstream from the main. The greater amount is near the pilot. For the Raman measurement, we 

found more CO2 near the pilot, with decreasing amounts as one moves radially outward to the main 

region. The discrepancy of there being very little relative CO2 concentration near the main may be due to 

an increase in luminosity in the background. The increase in background luminosity for the top two rows 

of the detector is approximately thrice and twice that of the bottom three rows. The source of the 

increased luminosity might be flame chemiluminescence arising from the main flame that is just above 

the field of view, or from blackbody radiation emitted by the casting, which is also just above the field of 

view. This increase in background luminosity decreased the signal to noise ratio to a level such that the 

signal was obscured. The plot in Figure 15(b) is located 10 mm downstream from Figure 15(a), and 

shows an increase in the amount of carbon dioxide including the area immediately downstream from the 

main annulus; the result is consistent with expectation and also with the thought that fuel interferences 

have some effect on the measurements in the region of the main injector. 

Subsonic N+2 Cases 

NASA’s N+2 subsonic engine is expected to have a larger core than the supersonic N+3 engine, and 

higher operating pressures. Thus, it has higher air and fuel flows in its cycle. We provide a brief synopsis of 

results. 

Subsonic Cruise 

Figure 16 shows an axial slice (aft-looking-forward perspective) near the dump plane, for the subsonic 

cruise condition (Table 2). We saw earlier that test conditions using fuel at ambient temperatures still exhibit 

a good deal of fuel in the liquid state after entering the combustor. However, for this condition, we 

determined using the method described above, as well as considering the total PLS and PLIF relative 

signals, that at most 10 percent of the total fuel signal was from the liquid. The higher inlet temperature and 

pressure, along with the commensurate higher flow rates and lower FAR contributed to higher vaporization 

and mixing, which is reflected in the more uniform fuel distribution compared to that shown in Figure 12 for 

the supersonic cruise condition. This case also performed well for emissions, with EINOX  5. 
NASA/TM—2014-218493 13 

Figure 16.—Result for N+2 Subsonic Cruise. Fuel 

pattern near dump plane for subsonic cruise 

condition as tested at NASA Glenn. The 

perspective is aft-looking-forward. 

Approach 

Figures 17 and 18 show results from two cases for approach conditions (Table 2), one with fuel 

staging between pilot and main (Approach-1), the other using pilot only (Approach-2). The gas analysis 

results showed higher NOx levels for Approach-2 with EINOX 27 percent higher than for Approach-1. 

Approach-1 used approximately 8 percent more fuel than Approach-2, with all other inlet parameters 

equal. In this comparison, the local equivalence ratio was important. Figures 17(a) and 18(a) show fuel 

patternation at the same axial position, using PLIF and PLS, and show that Approach-2 is less mixed than 

Approach-1. We also noted for Approach-1 that the PLS raw signal exceeded the PLIF signal; this is the 

only point tested for which this was the case, and indicates the majority of fuel was in the liquid phase. 

Figures 18(b) and 19(b) show the corresponding C2

*

 chemilumenescence, obtained by computing the 

mean from 200 instantaneous images, with exposure time for each image of 100 s. Flow is left to right. 

These images give an indication of where much of the carbon-carbon breakup occurs. Reactions are 

occurring closer to the dome for Approach-1 (implying reactions will be completed sooner), and without 

as much total intensity, than for Approach-2. These results indicate that if possible, fuel staging will be a 

better option for emissions during approach than running with pilot only. 

Idle 

The results at the elevated pressure idle condition (Table 2) are presented in a similar fashion as for 

the approach cases, in Figures 19 (case Idle-1) and 20 (Idle-2). The idle cases use only the pilot and case Idle1 has FAR 1.7 times case Idle-2, with both running fuel-rich. In this comparison, Idle-1 has less uniformity in 

fuel pattern and much more reactivity farther downstream, and resulted in higher EINOX levels. 

Emissions Measurements 

NOx emissions results from the NASA liquid-fuel tests are presented in Figure 21, for PICS injector 

operation over a range of inlet conditions. For these emissions tests, combustor inlet pressure was varied 

from 150 to 255 psia (including the 175 psia supersonic cruise inlet pressure), and combustor inlet 

temperature was varied from 850 to 1004 F (about 80 F below the supersonic inlet temperature). The 

measured EINOx data for all of these conditions are plotted, and show only weak dependence on inlet 

pressure and temperature over this range. Combustion efficiency was greater than 99.95 percent for all 

data points shown. Based on the raw measured NOx data, the PICS injector appears to deliver about 

5 EINOx at the supersonic cruise fuel-air-ratio, FAR/FARSLTO = 1.1, when operated on liquid-fuel only 

(i.e., unheated fuel that is not vaporized) in the NASA rig, at the combustor inlet temperatures and 

pressures listed. 
NASA/TM—2014-218493 14 

a) 

b) 

Figure 17.—Results from Approach-1. a) Total fuel signal 

represented by Fuel PLIF (red-yellow contour) and 

liquid fuel from PLS (blue contour), viewed aft-lookingforward. b) Mean C2* chemiluminescence signal within 

the vertical central plane (Y = 0) obtained by computing the average of 200 individual images. Flow is left to 

right. 

a) 

b) 

Figure 18.—Results from Approach-2. a) Total fuel signal 

represented by Fuel PLIF (red-yellow contour) and 

liquid fuel from PLS (blue contour), viewed aft-lookingforward. b)Mean C2* chemiluminescence signal within 

the vertical central plane (Y = 0) obtained by computing the average of 200 individual images. Flow is left to 

right. 

 
NASA/TM—2014-218493 15 

a) b)

a) 

Figure 19.—Results from Idle-1. a) Total fuel signal 

represented by Fuel PLIF (red-yellow contour) and 

liquid fuel from PLS (blue contour), viewed aft-lookingforward. b)Mean C2* chemiluminescence signal within 

the vertical central plane (Y = 0) obtained by computing the average of 200 individual images. Flow is left to 

right.

b) 

Figure 20.—Results from Idle-2. a) Total fuel signal 

represented by Fuel PLIF (red-yellow contour) and 

liquid fuel from PLS (blue contour), viewed aft-lookingforward. b)Mean C2* chemiluminescence signal within 

the vertical central plane (Y = 0) obtained by computing the average of 200 individual images. Flow is left to 

right. 
NASA/TM—2014-218493 16 

Figure 21.—Measured NOx emissions from the 

single-injector PICS combustor testing at NASA. For 

these tests, unheated liquid Jet fuel was supplied to 

both the main and pilot circuits, and emissions samples 

were obtained at several different operating conditions, 

including variations in inlet temperature, pressure, and 

fuel/air ratio. Emissions are plotted against FAR in the 

combustor, as calculated from total fuel flow and 

combustor airflow, normalized by the FAR at sea-leveltakeoff (SLTO) for this engine cycle. FAR/FARSLTO = 

1.10 at supersonic cruise, as indicated by the red lines. 

The blue lines indicate the FAR/FARSLTO = 1.0 sealevel-takeoff condition. 

An improved estimate of supersonic cruise EINOx can be obtained from the correlation developed by 

NASA under the Advanced Subsonic Technology (AST) Program, for various fuel-lean aircraft 

combustor configurations (Refs. 29 and 30). In this correlation, EINOx scales with inlet pressure as P30.595

(roughly square-root of P) and inlet temperature as exp(∆T3/194 K), or exp(∆T3/350 R) in English units. 

Using this correlation, EINOx data from 175 to 180 psia are shown in Figure 22(a) after scaling to the 

supersonic cruise combustor inlet temperature of 1087 F. The scaled data show an EINOx value of about 

6 at FAR/FARSLTO = 1.1 for supersonic cruise. We note, however, that in the NASA rig the combustor 

residence time was significantly longer than expected in an actual engine application, as a result of the 

12-in. distance between the PICS injector exit (the combustor dome) and the emissions-probe location. 

Thus, in an actual engine combustor we expect the liquid-fuel PICS NOx emissions to be well below the 

6 EINOx “worst-case” scenario predicted here. 

Figure 22(b) plots EINOx values at 850 to 900 °F inlet temperature, scaled to the 329 psia 

sea-level-takeoff pressure for this supersonic engine cycle. Thus, at 329 psia and 890 °F inlet conditions, 

for sea-level takeoff, the liquid-fuel PICS configuration is predicted to deliver an EINOx value of about 

3 (or less, given that the combustor residence time in the NASA rig is longer than expected in an aircraft 

combustor application). 

 
NASA/TM—2014-218493 17 

Finally, we note that the FAR values plotted in Figures 21 and 22 are based on a calculated airflow to 

the combustor. Two calculations are accounted for: First, because some air leaked around the combustor 

dome in the NASA rig (bypassing the primary combustion zone), and because this air leakage was not 

constant (it worsened during the course of testing), combustor airflow was calculated from the known 

effective area of the PICS injector and the measured pressure drop across the injector during testing. 

Thus, the leakage air (which did not participate in combustion) was ignored. Second, as in the UTRC 

testing (Ref. 3), the combustor airflow was set 14 percent below the cycle specification (making FAR 14 

percent higher than the cycle specification for each power condition) to account for an expected 14 

percent liner cooling-air requirement for a full-scale annular PICS combustor in an engine. This 

essentially corrects for the fact that the NASA flametube rig was ceramic-lined and not air-cooled, and 

therefore has a lower total air requirement than an engine combustor. We note that reducing the liner 

cooling-air requirement below 14 percent, such as enabled by new materials or cooling technologies, will 

improve NOx performance (reduce emissions) by enabling increased process air for leaner combustion. 

The NOx data in all cases are consistent in being reported without correction to standard humidity, since 

high-altitude supersonic flight occurs in dry air, and the facilities at both NASA and UTRC supplied dried 

air to the combustion test rigs. 

Conclusion 

We have presented computational and experimental analyses of the UTRC PICS injector, which has 

potential for application in an engine to meet the Pratt & Whitney N+3 supersonic cycle, or a subsonic 

N+2 engine cycle. NASA has defined technology targets for near term (called “N+1”, circa 2015), 

midterm (“N+2”, circa 2020) and far term (“N+3”, circa 2030) that specify realistic emissions and fuel 

efficiency goals for commercial aircraft. Experimental methods were also employed for select points of 

the subsonic cycle engine; cruise, approach, and idle with a slightly elevated inlet pressure. For the 

supersonic case, a PICS-based combustor was modeled using the National Combustion Code. 

Experiments at NASA employed gas analysis and a suite of laser-based measurement techniques to 

characterize the combustor flow downstream from the PICS dump plane. Gas analysis results showed that 

at both supersonic and subsonic cruise conditions, EINOx met the target goal of below 5. 

a) b)

Figure 22.—Predicted NOx emissions from the single-injector PICS combustor testing at NASA, based on scaling the 

measured results of Figure 21 to cruise conditions (left-hand panel) and sea-level-takeoff conditions (right-handpanel) for the supersonic cycle (Table 1) by applying the NASA correlation (Refs. 29 and 30) for NOx emissions 

from lean-burn aircraft combustors. 
NASA/TM—2014-218493 18 

We saw that for the supersonic cruise point, use of ambient temperature fuel at the de-rated air inlet 

temperature, did not result in fully vaporized fuel before exiting the main fuel-air mixing passage. Results 

from the optical techniques compared well with the numerical predictions, and show the fuel-air mixture 

is not uniform, which is to be expected because the fuel was injected at room temperature. Analysis of 

fuel PLIF and PLS results indicate that the liquid fuel is evaporated an axial distance approximately 

60 percent of the distance it takes the total remaining fuel to be consumed. Overall, the results show 

potential for meeting NASA’s supersonic cruise emissions goals with liquid fuel in all circuits. 

The subsonic cruise point performed well for emissions, with EINOX  5. For this point we 

determined, that at most 10 percent of the total fuel signal was from the liquid. The higher inlet 

temperature and pressure, along with the commensurate higher flow rates and lower FAR contributed to a 

higher vaporization rate and better mixing, which is reflected in the more uniform fuel distribution 

compared to that for the supersonic cruise condition. 

For the approach points, PLIF and PLS results show that Approach-2 with fuel staging between pilot 

an main is better mixed than Approach-1 which uses pilot only and that it produced lower NOx 

emissions. Analysis of the Approach-1 case also indicated that the PLS raw signal exceeded the PLIF 

signal, indicating that the majority of fuel was in the liquid phase. 

The idle cases use only the pilot. Results indicated that idle case with the higher FAR exhibits less 

uniformity in fuel pattern and much more reactivity farther downstream. This also resulted in higher 

EINOX levels. 

Using scaling methods previously developed at NASA, analysis of gas analysis results indicates that, 

at 329 psia and 890 °F inlet conditions, for sea-level takeoff, the liquid-fuel PICS configuration is 

predicted to deliver an EINOx value of about 3. 

References 

1. Lee, D.S., Pitari, G., Grewe, V., Grierens, K., Penner, J.E., Petzold, A., Prather, M.J., Schumann, U., 

Bais, A., Berntsen, T., Iachetti, D., Lim, L.L., and Sausen, R., Atmospheric Environment, Vol. 44, 

2010, pp. 4678–4734. 

2. Bulzan, D., “Supersonics Project High Altitude Emissions Overview,” 2012 Fundamental 

Aeronautics Program Technical Conference, 2012. 

3. Smith, L.L., Dai, Z., Lee, C., Fotache, C.G., Cohen, J.M., and Hautman, D.J., “Advanced Combustor 

Concepts for Low Emissions Supersonic Propulsion,” Proceedings of ASME Turbo Expo 2012, 

GT2012-69078, 2012. 

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

5. Quealy, A., Ryder, R., Norris, A., and Liu, N.-S., “National Combustion Code: Parallel 

Implementation and Performance,” NASA/TM—2000-209801, 2000. 

6. Quealy, A., “National Combustion Code Parallel Performance Enhancements,” NASA/CR—2002-

211340. 

7. Shih, T.-H., Povinelli, L.A., Liu, N.-S., and Chen, K.-H., “Generalized Wall Function for Complex 

Turbulent Flows,” NASA/TM—2000-209936, 2000. 

8. Chien, K.Y., “Prediction of Boundary-Layer Flows with a Low-Reynolds-Number Turbulence 

Model,” AIAA Journal, Vol. 20, No. 1, 1982, pp. 33-38. 

9. Shih, T.-H., Chen, K.-H., Liu, N.-S., Lumley, J.L., “Modeling of Turbulent Swirling Flows,” 

NASA-TM–113112, 1998. 

10. Shih, T.-H., Chen, K.-H., and Liu, N.-S., “A Non-Linear k-epsilon Model for Turbulent Shear 

Flows,” AIAA Paper 98-3983, 1998. 

11. Raju, M.S., “LSPRAY-II: A Lagrangian Spray Module,” NASA/CR—2004-212958, 2004. 

12. Venkateswaran, S., Weiss, J.M., Merkle, C.L., Choi, Y.-H., “Propulsion-Related Flowfields Using 

Preconditioned Navier-Stokes Equations,” AIAA-92-3437, 1992. 

13. Venkateswaran, S., Merkle, C.L., “Efficiency and Accuracy Issues in Contemporary CFD 

Algorithms,” AIAA–2002–2251, 2002. 
NASA/TM—2014-218493 19 

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

Combustion Code,” AIAA Paper 98-3855, 1998. 

15. Iannetti, A., Tacina, R., Jeng, S.-M., and Cai, J., “Towards Accurate Prediction of Turbulent, ThreeDimensional, Recirculating Flows With the NCC,” NASA/TM—2001-210761, AIAA–2001–0809, 

2001. 

16. Shih, T.-H., Norris, A., Iannetti, A., Marek, C.J., Smith, T.D., Liu, N.-S., and Povinelli, L.A., 

“A Study of Hydrogen/Air Combustor Using NCC,” AIAA–2001–808, 2001. 

17. Iannetti, A.C., Chen, K.-H., “An Initial Comparison of National Combustor Code Simulations Using 

Various Chemistry Modules With Experimental Gas Turbine Combustor Data,” AIAA–2000–0330, 

2000. 

18. NX, formerly Unigraphics NX, http://www.plm.automation.siemens.com/en_us/products/nx/. 

19. Cubit, https://cubit.sandia.gov/, Sandia National Labs, Albuquerque, NM, USA. 

20. Pointwise, http://www.pointwise.com/pw/, Pointwise, Inc., Fort Worth, TX, USA. 

21. Westbrook, C.K. and Dryer, F.L., “Simplified Reaction Mechanisms For The Oxidation Of 

Hydrocarbon Fuels In Flames,” Combust. Sci. Technology, Vol. 27, 1981, p. 31. 

22. Ajmani, K., Kundu, K., and Penko, P., “A Study on Detonation of Jet-A Using a Reduced 

Mechanism,” AIAA–2010–1515, 2010. 

23. Zeldovich, Y.B. “The Oxidation of Nitrogen in Combustion and Explosions,” Acta Physicochimica 

URSS, Vol.21, 1946, pp. 577–628. 

24. Fennimore, C.P., 13th Symposium (Int) on Combustion, 1971, p. 371. 

25. Amsden, A.A., O’Rourke, P.J., and Butler, T.D., “KIVA-II: A computer program for chemically 

reactive flows with sprays,” Report Number LA-11560-MS, Los Alamos National Lab, 1989. 

26. Lefebvre, A.H., Atomization and Sprays, Hemisphere Publishing Corporation, New York, 1989. 

27. El Banhawy, Y. and Whitelaw, J.H., “Calculation of the Flow Properties of a Confined KeroseneSpray Flame,” AIAA Journal, Vol. 18, No. 12, 1980, pp. 1503-1510. 

28. Swanson, R.C., and Turkel, E., “Multistage Schemes with Multigrid for Euler and Navier–Stokes 

Equations,” NASA-TP-3631, 1997. 

29. Tacina, R., Wey, C., and Choi, K., 2001, “Flame Tube NOx Emissions Using a Lean-Direct-WallInjection Combustor Concept,” NASA Technical Memorandum no. NASA/TM—2001-211105 and 

AIAA Paper AIAA–2001–3271, 2011. 

30. Tacina, R., Wey, C., Laing, P., and Mansour, A., “A Low NOx Lean-Direct Injection, Multipoint 

Integrated Module Combustor Concept for Advanced Aircraft Gas Turbines,” NASA/TM—2002-

211347, 2002. 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and

maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including

suggestions for reducing this burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite

1204, Arlington, VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it

does not display a currently valid OMB control number.

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YYYY)

01-09-2014

2. REPORT TYPE

Technical Memorandum

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

NASA Numerical and Experimental Evaluation of UTRC Low Emissions Injector

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Hicks, Yolanda, R.; Tedder, Sarah, A.; Anderson, Robert, C.; Iannetti, Anthony, C.;

Smith, Lance, L.; Dai, Zhongtao

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 699959.02.09.03.02

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18953-1

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2014-218493

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Category: 07

Available electronically at http://www.sti.nasa.gov

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

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Computational and experimental analyses of a PICS-Pilot-In-Can-Swirler technology injector, developed by United Technologies Research Center (UTRC) are presented.

NASA has defined technology targets for near term (called “N+1”, circa 2015), midterm (“N+2”, circa 2020) and far term (“N+3”, circa 2030) that specify realistic emissions

and fuel efficiency goals for commercial aircraft. This injector has potential for application in an engine to meet the Pratt & Whitney N+3 supersonic cycle goals, or the

subsonic N+2 engine cycle goals. Experimental methods were employed to investigate supersonic cruise points as well as select points of the subsonic cycle engine; cruise,

approach, and idle with a slightly elevated inlet pressure. Experiments at NASA employed gas analysis and a suite of laser-based measurement techniques to characterize the

combustor flow downstream from the PICS dump plane. Optical diagnostics employed for this work included Planar Laser-Induced Fluorescence of fuel for injector spray

pattern and Spontaneous Raman Spectroscopy for relative species concentration of fuel and CO2

. The work reported here used unheated (liquid) Jet-A fuel for all fuel circuits

and cycle conditions. The initial tests performed by UTRC used vaporized Jet-A to simulate the expected supersonic cruise condition, which anticipated using fuel as a heat

sink.Using the National Combustion Code a PICS-based combustor was modeled with liquid fuel at the supersonic cruise condition. All CFD models used a cubic non-linear

k-epsilon turbulence wall functions model, and a semi-detailed Jet-A kinetic mechanism based on a surrogate fuel mixture. Two initial spray droplet size distribution and spray

cone conditions were used: (1) an initial condition (Lefebvre) with an assumed Rosin-Rammler distribution, and 7 degree Solid Spray Cone; and (2) the Boundary Layer

Stripping (BLS) primary atomization model giving the spray size distribution and directional properties. Contour and line plots are shown in comparison with experimental

data (where this data is available) for flow velocities, fuel, and temperature distribution. The CFD results are consistent with experimental observations for fuel distribution and

vaporization.

15. SUBJECT TERMS

Laser diagnostics; Gas turbine combustor modelling; Combustion

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU

18. NUMBER

 OF

 PAGES

26

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) a. REPORT

U

b. ABSTRACT

U

c. THIS PAGE

U 19b. TELEPHONE NUMBER (include area code)

443-757-5802

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18