NASA Technical Memorandum 103176

AIAA-90-2452

Introducing the VRT Gas Turbine Combustor

Jerry O. Melconian

SOL-3 Resources, Inc.

Reading, Massachusetts

Abdu A. Mostafa

Textron Lycoming

Stratford, Connecticut

and

Hung Lee Nguyen

Lewis Research Center

Cleveland, Ohio

Prepared for the

26th Joint Propulsion Conference

cosponsored by the AIAA, SAE, ASME, and ASEE

Orlando, Florida, July 16-18, 1990

INTRODUCING THE VRT GAS TURBINE CONBUSTOR

Jerry O. Melconian*

SOL-3 Resources, Inc.

Reading, MA 01867

Abdu A. Mostafa*

Textron Lycoming

Stratford, CT 06497

Hung Lee Nguyen*

NASA Lewis Research Center

Cleveland, OH 44135

ABSTRACT

An innovative annular ccmbustor configuration

is being developed for aircraft and other gas

turbine engines. This design has the potential of

permitting higher turbine irlet temperatures by

reducing the pattern factor end providing a major

reduction in NOx emission, lhe design concept is

based on a Variable Residenc_ Time (VRT) technique

which allows large fuel particles adequate time to

completely burn in the circumferentially mixed

primary zone. High durabilit) of the combustor is

achieved by dual function use of the incoming air.

The feasibility of the concept was demonstrated by

water analogue tests and 3-D computer modeling.

The computer model predicted _ 50 percent reduction

in pattern factor when compared to a state of the

art conventional combustor, lhe VRT combustor uses

only half the number of luel nozzles of the

conventional configuration. The results of the

chemical kinetics model require further

investigation, as the NOx predictions did not

correlate with the availa_le experimental and

analytical data base.

I.O INTRODUCTION

1.1 Future aircraft operatinu in both the subsonic

and supersonic regime will be powered by engines

operating at higher cycle eff ciencies than current

gas turbines. Increases in cycle efficim_cy of

simple aircraft gas turbines require an increase in

mean turbine inlet temperatures with

disproportionately large incr(}ases in peak turbine

inlet temperature and NOx fomation. To meet these

challenges, SOL-3 Resource:_ Inc. proposed an

innovative annular combustor _:oncept which is based

on a circumferentially mixed vortex flow principle

with variable residence time (VRT) for the fuel

particles. The salient features of the combustor

are:

* Circumferentially mixed vortex flow for low

pattern factor.

* Rich-Quench-Lean com)ustor for low NOx

emission.

* Variable fuel resilence time for high

efflciency at all operating conditions.

$MimblrAIAA

* Walls cooled by 60 percent of total

available air for liner durability.

Elimination of compressor exit flow

straightening vanes and reduction in the

number of fuel injectors and turbine nozzle

vanes for a simple engine.

1.2 The Phase I SBIR (Small Business Innovation

Research) effort supported by NASA Lewis Research

Center was aimed at demonstrating the feasibility

of the VRT combustor concept in an annular

configuration. A Textron Lycoming engine of the 10

Ib/sec class was chosen as host for the combustor

to provide a realistic physical boundary envelope.

The design operating conditions were based on the

Lycoming turbine flow numbers and temperatures and

pressures which would be experienced at a

supersonic cruise condition of 65,000 feet altitude

and Mach 3. The approach used a three-dimensional

mathematical model running both cold and reacting

flow conditions. The cold flow model was validated

experimentally by a clear plastic model in a water

analogue rig. The reacting flow conditions will be

validated by combustor laboratory tests which are

planned for Phase II.

2.0 THE VARIABLE RESIDENCE TIME (VRI) COMBUSIOR

2.1 Description of the Concept

In a conventional 9as turbine engine, ambient

air is compressed and given axial and

circbmferential velocities by a compressor. At the

exit of the compressor one or more sets of

straightening vanes are used to change the swirl

velocity into an axial direction. The air then

goes through a combustor where it is mixed with

fuel, ignited, and its temperature is raised. At

the exit of the combustor, a set of nozzle vanes

again turns the flow so that the hot gas is

directed at the correct an91e for entry to the

turbine.

The three basic combustor concepts currently

in wide use are a) the swirler stabilizer design,

b) the vaporizing baffled design and c) the slinger

design. The fundamental reason for these concepts

is that in a continuous combustion system, the

continually injected charge of fuel needs a source

of ignition. This is achieved by recirculatin9 the

burning hot gases. The recirculation ratio and

turbulent flame speed together with combustor inlet

pressure and temperature place an upper limit on

the through flow velocity in the orimary zone.
VAT CONFIGURATION

F[_iER INJECTORS

SHORTER nlFFUSER SIMPLER COHBUSTON

EXIT STATOR$

ELIMINATED

SHORTER

FEWE_ _OZZLE

VANES

CONVEHTIONAL COHFIGURATION •

FIGURE I - THROUGH FLOW ENGINE CONFIGURATION

$

l ii'i ill FittIll I '

FIGURE 2 - VRT COHBUSTOR MAIN FLOW PATTERN

With current designs the mean through flow velocity

in the primary zone is of the order of 20 feet per

second.

The VRT combustor in a straight through

configuration is shown in Figure I together with an

equivalent conventional design. The noticeable

differences are the elimination of compressor exit

straightening vanes and a shorter diffuser leading

to a shorter engine. Not as obvious are the

reduced number of fuel injectors and number of

turbine nozzle vanes.

Referring to Figure 2, the air entering the

combustor at ST3 has both axial and swirl

velocities. Both velocities are reduced in the

diffuser section. As the flow path of the air is

the resultant vector of the axial and swirl

velocities, the diffuser effective angle is smaller

than the 9eometrlc angle and a shorter diffuser is

possible with the VRT configuration because the

\

P,_ IIICN

I:O,_tts T101_

FIOURE 3 - TRANSVERSE FLOW PATTERN IN PRIMARY ZONE

swirl is not taken out. The air enters the primary

zone of the combustor through axial slots which

maintain the swirl component to provide

circumferential mixing. Within the primary zone, a

combined vortex is established. Fuel is injected

in a tangential direction (i.e. in the plane of the

paper) to reinforce the vortex. As the fuel is

burnt, there is a natural stratification of the

droplets due to the centrifugal effect of the

vortex. The heavier fuel droplets are flung to the

outer circumference of the toroidal shaped primary

zone and the lighter particles remain at the inner

circumference. The shape of the waist forces the

heavy particles to remain in the vortex until they

are burnt and replaced by new heavy particles

entering. Opposed air Jets In the waist provide

rapid quenching of the fuel rich combustion

products leaving the primary zone. In the

secondary zone, air is introduced by means of

tangential slots to maintain the swirl and complete

the combustion of unburnt gaseous products. In the

dilution zone, air is again introduced by means of

tangential slots to tailor the swirl and the radial

temperature profile of the combustor exit gases to

that acceptable to the turbine.

ORIGINAL PAGE IS

OF POOR QUALITY
Conventional annular cof_bustors are like a

number of can combustors plac,_d side by side with

the side walls removed. Each injection point has

its own flame holding devise and the main flow

proceeds axially in the x-direction. Mixing of

fuel and air and ignition of freshly injected fuel

is by secondary vortices in the x-y plane.

Propagation of the flame from injector to injector

is at the interface of the two "cans".

Circumferential mixing, if any, also occurs at this

interface. Acceptable combusLor exit temperature

distribution is achieved by s_rict control of the

air apertures in the liner and the fuel passages in

the injectors.

The mixing process in the primary zone of the

VRT combustor is radically different from

conventional annular designs. In the VRT

configuration, the main flow Is a spiral in the y-z

plane with its axis in the axial direction. The

primary zone can be visualized as a series of

curved "can" combustors placed end to end to form a

circle. The axis of the fuel spray is in the y-z

plane and ignition of freshly injected fuel is by

the flame from the preceedin_ injector. Mixing of

fuel and air is by shear layers in the y-z plane

enhanced by secondary vortices in the x-y plane

(Figure 3). The net _ffect is a true

circumferentially stirred annLlar combustor. Flow

reversal no longer being ne¢essary to ignite the

freshly injected fuel, the through flow velocity

can be increased substantially.

2.2 Advantages of Concept

The proposed concept has the following

advantages over current gas tt;rbine designs:

.AerothermodTnamic Advant_._es

I. The circumferentially mixe(i vortex flow provides

the necessary conditions for a low temperature

pattern factor at the exit of the combustor. A

reasonable target for the developed hardware is

a pattern factor of 0.15 c(_mpared to a value of

0.25 for the best conventi(mal combustor.

2. The Rich-Quench-Lean burn technique used in the

combustor partitioning ha:_ been shown to be one

of the effective method:_ of limiting flame

temperatures and thus the formation of NOx.

Preliminary calculations indicate that an NOx

Emission Index of 5 g/kg is achievable at high

altitude supersonic _Ircraft operating

conditions.

3. The Variable Residence rime feature which

prevents liquid fuel droplets from leaving the

primary zone ensures a broad combustion

stability envelope and nearly 100 percent

efficiency at all operatin_ conditions. It also

provides a true multi-fuel capability.

Mechanical Advantages

I. Elimination of the compressor exit stator and a

shorter diffuser in an axial compressor

configuration leads to a shorter engine. With a

centrifugal compressor, the higher acceptable

diffuser exit velocity and elimination of

turning vanes leads to a smaller engine diameter

2.

3.

4,

Nearly 60 percent of the total available air

enters the combustor through tangential total

head apertures and will first cool the walls

before taking part in combustion or exit

temperature control. The effectiveness of any

cooling method will be further enhanced by the

larger quantity of available air.

The circumferentially stirred primary zone is

not limited by the need for flow reversal and

carl therefore be operated at much higher air

velocities. The result is a reduction in

combustor size and number of fuel injectors.

The high velocity, high swirl combustor exit

flow conditions will reduce the number of

turbine nozzle vanes required. Preliminary

calculations indicate a combustor exit swirl

angle of ?0 degrees at a Mach number of 0.23 for

the through flow configuration. For the reverse

flow configuration, the swirl angle is 78

degrees and the Mach number 0.38.

3.0 COHBUSTOR PRELIMINARY DESIGN

3.1 Combustor Dimensions

In arriving at the combustor dimensions, two

factors were considered to be of paramount

importance. The first was the capability of the

engine to relight at 30,000 ft altitude. Although

the aircraft is expected to cruise at 65,000 ft, it

would normally descend to 30,000 ft for relight in

the event of an engine flame out. The second

factor which was important for this program, was

the ability to use the Lycomin9 test rig without

major modifications.

The design methodology for gas turbine

combustors is based on semi-empirical correlations

developed over the past 40 years. The most widely

used loading parameter Is Lefebvre's Theta, e ,

parameter (Ref. I) which is an evolution of the

kinetic rate formulation and time.

A large base of experlmental data has shown

that to achieve combustion efflciencies exceeding

95 percent, Theta must equal 73 x 106 (SI units).

This level of efficiency at the 30,000 ft windmill

condition would virtually guarantee an engine

relight. Solving the equation gives us a required

liner cross-sectional area of 0.03052 sq m (47.32

sq in). The Lycoming high performance combustor

liner has a cross-sectional area of 0.04055 sq m

(62.83 sq in). The Lycoming cross-sectional area

was chosen for the VRT combustor as this gave us a

33 percent margin and compatibility with the test

rig and engine. The remaining combustor dimensions

shown in Table I were arrived at through geometric

considerations, air flow pattern visualization, and

engine space compatibility.

3.2 Supersonic Cruise Operation

The supersonic cruise operating conditions

were synthesized by using the Lycoming engine

combustor exit flow number, which determined the

mass throughput of the engine, and the NASA High

Speed Civil Transport Emission Data of Niedzwiecki

(Ref. 2), which provided the state of the fluid.

The lower end of the range of the NASA data was

used as this was considered more realistic of the

next generation of gas turbines.
1AC LI[ L

PHYSICAL OIH£NSIOHS OF VRT ¢O_SUSTOR LIH(R (INNER SURFACE)

Overall Length 4,1 tn 0.10414 m

Prtmery Zone Height 2.0 In 0.0508 m

|econdary Zone Hetght 1.9 |n 0,04826 m

Exit Height 0.9 tn 0.02286 m

Wall Perimeters

Outer S.188 tn 0.13178 m

Inner 4,751 tn 0,12088 m

Primary Zone 4.501 tn 0,11433 m

Second&ry Zone 3.125 in 0,07938 m

011ut(on Zone 2.313 tn 0.05875 m

Crose-Sect|onel Aria, X-Y Plane

Prtmsr¥ Zone 2.68 eq tn 0.001718 eq m

8e¢ondery Zone 2.75 eq tn 0.001774 e@ •

Dtlutton Zone 1.30 sq tn 0.0008387 sq m

Combuetor Volume 0.1lEt4 cuft 0.003178 cu m

TAOL( 2

SYNTHESIZED _JP(RSONIC CRUIBE OPERATING ¢OHOITIOH|

Atr Flow, • 5.017| ppo 2.2?8? kg/oec

3

Fuel Flow, • 0.1772 pps 0.08030 kg/sec

f

Gas Ftow, m 5.195 ppe 2,356 kg/eeo

3.8

Inlet Preoeure, p 176.4 pete 1.216 HPe

3

Pressure Drop

Ltner 3.0

Dtffuser Exit 0.6 S

Inlet Temperature, T 1459 R 010.6 K

3

Extt Temperature, T 3489 R 1921.7 K

3.8 -6

Extt Flo_ Function 1.795 (ENG) 06.02x10 (8I)

Hot Residence Time, _" 0.00274 cec

COttBUSTOR LOAD ING

6

Engltsh Unite 0.74x10 Btu/hr,cu, ft.st_,

12

8I Unite 3,90x10 W/hr.cu.m.stm.

4o

P. - l| at•

i J

T1 - 810.6 K

w

i

N

g

4 0,1 O_ I_ 1.1 IA I,I 1.1

EOUIVAL[HCE _TIO - _ - FAK/FAR.STOIC.

lt_l '" loll 2_TI 1S$1 _$1l 23il =201 |051

_AS T[t_ZRATUR[ - T9 - K

FIGURE 4 - PREDICTED NOx EMISSIONS

The pressure drop across the liner was set at

3 percent to match the Lycoming high performance

combustor pressure drop. Although the VRT

configuration could use a lower pressure drop, this

pressure drop also controls the turbine cooling

flow. Any changes to the pressure drop would

require a redesign of the turbine cooling passages

which at this stage of the program would prevent

the use of the Lycoming rig and engine.

The resultant synthesized supersonic cruise

operating conditions are shown in Table 2. The low

residence time of 2.74 millisecond will be a

significant factor in reducing the NOx emissions.

3.3 NOx Emissions

The NOx emissions of greatest concern are

those occurring at the supersonic cruise altitude

of 65,000 ft. Thls is where the greatest damage to

the Earth's ozone layer is likely to occur due to

the catalytic action of NOx. The target NOx

emission index at these operating conditions has

been set at 5 grams NOx per kilogram of fuel. A

rich burn, quick quench, lean burn configuration

was chosen to minimize the NOx emission.

T_L£ 3

COHBUSTOR [_JIVALEHC( P_TIO & RIR OI|TRIBUTIOH

COHEUSTOR [QUIVALEHC( RATIO t

For Emtsoton Index EZ,NOx m 5 ct 8upersontc Cruise

Overell, 0.62

Primary Zone, 1.72

8econdsry Zone, 0.62

AIR PARTITIONIHG

Prtmery Zone, 30.23 percent

Quench, 87.18 percent

Dilution 12,59 percent

PHYSICAL AIR DISTRIBUTION

Atrblast InJectors (8) 2.0 percent

Prtmery Zone Tengentte| 27.43 percent

Secondary Zone Radtal 40.47 percent

8econdsry Zone Tangential 16.71 percent

011utton Zone Tangenttsl 1E.89 percent

To estimate the NOx Emissions Index, the

correlation derived by Odgers (Ref. I, Vol. If) was

used. Figure 4 shows the estimated emission index

versus equivalence ratio at the altitude cruise

combustor operating conditions. The curve

correlates with the data given by Nguyen et al

(Ref. 3). To achieve an EI(NOx) of 5 g/kg, the

rich primary zone will have to operate at an

equivalence ratio of 1.72 and the lean secondary

zone at an equivalence ratio of 0.62. These

results together wlth the required air partitioning

are given in Table 3.

3.4 Combustor Air Distribution

The air partitioning calculations show that to

obtain a primary zone equivalence ratio of 1.72

will require 30 percent of the available air flow.

This air would be divided between the airblast fuel

injectors and the tangential apertures in the

primary zone.

At this time, a choice had to be made on the

number of circumferential sectors and the type of

fuel nozzles. The Lycomlng configuration has 18

fuel nozzles which flow 16 percent of the total

air. The decision was made to go with 8 sectors

for the VRT configuration and use available fuel

nozzles from the Lycoming ALF 502 engine.

Unfortunately these nozzles limit the quantity of

alrblast to 2.8 percent.

ORIGINAL PAGE IS

OF POOR QUALITY
c \

\

\

B

_"'D L.T _ , ,

mR_

FIGURE S - VRT COI4BU_TOR GEHERAL ASSEMBLY

The alr partitioning calculations also showed

that to achieve a transition from an equivalence

ratio of 1.72 in the primary zone to 0.62 in the

secondary zone would require 57 percent of the

total air flow for quenching. Thls quantity of air

was split between the radial jets at the waist (40

percent) and the secondary zone tangential

apertures (17 percent). The remaining 12 percent

of the available air was all,}cated to the dilution

zone. The detailed physical air distribution is

shown in Table 3.

3.5 Liner Wall Apertures

Next, the zone open arenas were calculated and

converted into physical oper_Ings. A number of

iterations was needed to obta n an optimum aperture

shape, size and spacing. The final VRT combustor

configuration for manufactur(_ in plastic is shown

in Figure 5.

3.6 Gas Temperatures

For these calculations the combustor was

divided into three zones. Th(_ rich primary zone is

where the fuel and air are intimately mixed and the

fuel vaporized. The lean secondary zone is where

the gases are quenched and ¢ombustion completed.

In the dilution zone th(_ remaining air is

introduced to reduce the gas temperatures to that

acceptable to the turbine. TPe temperature in each

zone Is assumed to be the mean between the inlet

and exlt temperatures of the zone. The temperature

rise is obtained from standard charts. The

resultant calculated temperatures in each zone is

given in Table 4.

Tl_ll* [ 4

TDIpEAATURER

Inlet Te4@ereture, T 1459,0 R 810.1 K

3

Prtmsry Zone Average zBgs.I R 1442.0 K

Primary Zone Extt 3?Zg,6 R 2072.0 K

Secondary Zone Average 3692,3 R ZOII.S K

8econdery Zone ExH; 3655.1 R 2030.0 K

Dtlutt0n Zone Average 3ssg,2 R lg?g.| K

Dilution Zone Ex|t, T 3489.0 R Igll.? K

3.6

TAISLE S

COtIGUSTOR Wi_LLS t'l_XZt_JM TEHPERKruREs

Primary Zone Outer

Primary Zone Inner

legendary Zone Outer

legendary Zone Inner

Ollu¢lon Zone Outer

Dilutton Zone Inner

1053.6 R 91$.? g

1631,3 R 90g.3 K

2313.3 R 1285.2 K

Z176,1R 1208.g g

Zg54,0 R 168g.l K

2698.4 R 1499.1K

3.7 Combustor Wall Temperatures

In a conventional combustor, the wall cooling

apertures are distinctly separate from the

aerodynamic apertures. The former are designed to

provide a wall hugging air blanket, whereas the

later are penetrating Jets. For the VRT

configuration, except for the quenching Jets at the

waist, the wall cooling and aerodynamlc apertures

are the same. The air exiting the tangential

apertures hugs the wall to provide the cooling and

the driving force for the vortex.

ORIGINAL PAGE IS

OF POOR QUALITY
The present VRT combustor was designed with

eight fuel injectors. For symmetry, the number of

sectors must be a factor of the number of

injectors. For simplicity in the first design, the

number of sectors was also set at eight. The

resultant estimated maximum wall temperatures using

the techniques described in Reference I are given

in Table 5.

3.8 Air Flow Vector Analysis

One of the unique features of the VRT

combustor is the circumferential vortex. A

preliminary analysis of the expected velocity and

swirl angles was performed using the principles of

conservation of angular momentum and continuity in

the axial direction. The effects of frictional and

mixing losses were neglected in these calculations.

The results of this cold flow analyses can be

compared to the water analogue flow fields and the

non-reactlng computer model. The results of the

analysis are given in Table 6 together with the

cold and hot exit Mach numbers.

TAaLE 6

COLD AIR FLOW VECTOR ANALYSIS

PRIMARy ZONE MEAN - STATION 3.2

Tangential Velocity, w 395.2 ft/seC 120.5 m/sec

3,2

Axtet Velocity, u 6.O5 ft/aec 1.8 fft/se_c

3.2

Swtrl Angle, _ 89.1 degrees

3.2

PRIMARY ZONE EXIT - 9TATIOH 3.3

Tangential Velocity, w 395.2 ft/sec 120.6 m/sec

3,3

Axtsl Velocity, u 16.0 ft/oec 4.9 m/sac

3.3

Swirl Angte, D 87.7 degrees

3.3

SECC_OARY ZONE IMLET - STATION 3.4

Tsnosnttel Velocity, w 395.2 ft/sSc 820.6 m/see

3.4

Axtel Velocity, u 37.5 ft/|e¢ 11.4 m/sec

3,4

Swirl Angle, p 77,E degrees

3,4

SECONDVMIy ZONE flESH - STATION 3.5

Tangential Velocity, w 395.2 ft/sec 120.5 m/sic

3.5

Axtsl Velocity, u 32.9 ft/eo¢ 10.O m/lee

3.5

9wlr1 Anglo, B 80.3 degrees

3.9

OZLUTICR4 ZOtt[ MEAN - STATION 3.&

Tsngsnttsl Velocity, w 395.2 ft/eoc 120,5 rn/8sc

3.8

Axis| Yetoctty, u E8.E ft/8oc 20.4 m/8ec

3.6

|wlrl Angls_ D 73.4 degrees

3,6

COi18USTOR EXIT - STATION 3.7

Tangential Velocity, w 395.2 ft/0ec 120.5 m/see

3.7

Axial Velocity, u 98.6 ft/ss¢ 30.0 m/sac

3.7

|wlrl Angle, B 67.3 degrees

3.7

COtlBUSTC)A EXIT MACH MUtlBER - STATION 3.7

Cold Msch Number 0.120

Not Ni©h Number 0.180

4.0 WATER ANALOGUE TESTS

4. I Installation

The Lycomlng water analogue test facility

consists of a vertical glass testway with a 12 x 24

Inch (0.3 x 0.6 m) rectangular test section.

Because of the depth limitations of the testway,

the model was mounted horizontally and bolted to a

collector. The flow approached the model

vertlcally from below, was turned horizontal within

the model and returned to the vertical inside the

collector. Shop air was available to 25

individually controlled small bore tubes within the

testway. A large holding tank and reclrculating

pump allow a controllable water flow with a maximum

rate of over I00 pps.

Observations were carried out at flow rates of

15 pps (Reynolds No. = 0.89 x 105 ) and 30 pps

(Reynolds No.: 1.8 x I0S). Except for the higher

flow speed, the flowpaths at the two flow rates

were virtually identical.

The model was instrumented with ten 0.040 inch

O.D. tubes feeding the apertures shown in Figure 6.

Eight fuel nozzles were inserted into the fuel

nozzle holders in the header and retained with

silicone rubber adhesive. One of the fuel nozzles

was instrumented with 0.125 inch O.D. tube to allow

feeding air or dye to the main fuel passage.

4.2 Test Observations

The tests consisted of allowing time for the

water flow to stabilize and then introducing air

bubbles at each of the injection points In turn.

Written and video records were made at three series

of tests. The observations were as follows:

Test Series I - Water flow rate 15 pps

The flow went about 350 degrees in the primary

zone. The flow hugged the wall, was temporarily

deflected at each fuel nozzle location but

reattached itself to the wall. The reattachment

after each nozzle was due to a superimposed swirl.

The quench jets produced a secondary

counter-clockwise swirl in the outer and a

clockwise swirl in the inner section of the

secondary zone. The flow did one further

revolution in the secondary and dilution zones

before exiting the combustor. The main flow

pattern is shown in Figure 6.

® ®

0 AIR IUIIL|S IIUI¢TI{Wl MILTS

FIOURE E - VRT COHBUSTOR WATER ANALOGUE FLOW PATTERNS

Test Series 2 - Water flow rate 30 pps.

The flow pattern was essentially the same as

Test Series 1. Turbulence and mixing were

increased, and the flow path around the

circumference appeared to have increased slightly

to 370 degrees.

ORIGINAL PAGE !_

OF POOR QUALITY
Test Series 3 - Water flow rates 15 pps and 30 pps.

The third series of tests was with red dye

being fed into the main fuel passage of the fuel

injector. The circumferential flow pattern was

essentially the same as seen in Test Series I, with

one complete turn in the primary zone. It took

about 90 degrees circumferentially for substantial

mixing between the dye exiting the fuel nozzle and

the bulk of the primary zone flow entering

tangentially. After this initial 90 degrees

however, the dye filled the whole cross-section of

the primary zone and was totally dissipated by the

time it reached the secondary zone.

5.0 THREE-DIMENSIONAL COMPUTER MODEL

This section presents the 3-D computer model

analysis conducted by Textron Lycoming under

subcontract to SOL-3 Resources. In addition to the

performance analysis of the VRT combustor, a

comparison was made with the results obtained on a

state of the art Lycoming designed combustor of

conventional configuration utilizing 18 swirler

stabilized fuel injectors.

5.1 Description of the 3-D Combustor Model

The 3-O combustor performance model computer

program ("FLUENT") that forms the basis of the

present work is described ir Reference 4. "FLUENT"

is a general purpose computer program that solves

steady state continuity, momentum, and energy

equations in 2D/3D Cartesiar or polar coordinates

for laminar or turbulent flows. Turbulence is

modeled using the standard two-equation K-E model

or Algebraic Stress Model (ASM). "FLUENT" also has

a six equation radiation model that allows

calculation of radiative heet transfer. A single

step global chemical reaction is used to model

reacting flows. The conservation equations for

three chemical species are solved to calculate the

local reacting flow field. Four reaction rates are

computed at each node point in the flow field; the

kinetic reaction rate, and three mixing controlled

rates based on the laminar cr turbulent diffusion

of the three chemical species respectively. The

local reaction is then allo_ed to proceed at the

lowest of these four rates.

The liquid spray is modeled using a Lagrangian

trajectory technique for groups of droplets based

on the computed continuous _as flow field. Effect

of turbulence on the trajectories of the droplets

is simulated using a stochastic tracking technique

which takes into account the local turbulence

characteristics as the droplet traverses the flow

field. The calculations of the liquid spray and

continuous phase are fully coupled so that

evaporation and combustion of liquid fuel droplets

can be modeled in a realistic manner. A partial

equilibrium Zeldovich mechanism is used in the NOx

calculation. The NOx reaction sequence is solved

under the assumption that it has no impact on the

predicted flowfield or tem[erature field in the

combustor.

The NOx reaction sequen(e is solved decoupled

from the flowfield and temrerature field in the

combustor. It is also asstmed that the reaction

rates are infinitely fast so that the species

concentrations may be ottalned via chemical

equilibrium. The "CREK" (Ref.5) program is used to

solve the conservation equations for the species

involved in the NOx production.

The equilibrium assumption reduces the number

of kinetic equations to be solved and was used in

several premixed combustion systems. However, this

assumption may not be quite valid in all regions of

a gas turbine combustor.

5.2 Results

The results include velocity vectors,

temperature isotherms, species_ and emission

concentrations. The analysis covers one of the

derivatives of the Lycoming combustors and SOL-3

VRT combustor. The cruise condition given in Table

2 was used in the calculations and JP4 fuel was

used as a fuel type in the spray calculation.

The VRT Combustor

A 45 degree sector in polar coordinates with

cyclic boundary conditions at the repeating planes

was used to model the entire combustor. The

computational grid comprised 35x36x31 elements in

axial, radial, and tangential directions

respectively. The fuel spray was modeled with

droplets injected at 8 discrete angular positions

with 10 discrete drop sizes. Figure 7 illustrates

a 3-D view of the fuel droplet trajectories with

crosses indicating the point of complete

evaporation. This figure also shows that all

droplets are fully evaporated within a distance

approximately equal to half the circumferential

spacing between fuel injectors. The smallest

droplets evaporate very near the fuel nozzle while

the largest travel a longer distance. The droplets

are heated to the vaporizations temperature, at

which point they begin to evaporate by convection

and diffusion into the gas phase. When the droplet

temperature reaches the boiling point, a boiling

rate equation governs the mass transfer. As the

droplet evaporates, it enters the gas as a chemical

species "f" and reacts with oxygen present in the

gas. This process creates interphase coupling of

heat, mass, and momentum transfer between the gas

and liquid phases.

°.-o

Looking Dovnstre&m * DoLnt O| COtpl|t,t ovipor|t.lon

FIGURE 7 - VRT COI48USTOR MEAN DROPLET TRAJECTORIES

ORIGINAL PAGE IS

OF POOR QUALITY
(e) Temperature contours 0oR (b) velocity vectorsa 3-0 14agnitude & DirectS[on

(c) Unburned fuel contours (d) Oxygen

,mass fraction

FIGURE E - VRT COH6USTOR PREDICTIONS AT AN AXIAL PLANE

(e) Temperature contours •°R (b) Velocity vectorss 3-D Magnitude & Direction

(c) Unburned fuel contours

rmass fraction

(d) Oxygen

,mess Erectiol_

FIGURE 9 - VRT COI4BUSTOR PREDICTIONS IN PRIMARY ZONE

ORIGINAL PAGE IS

OF POOR QUALITY
Figure 8 illustrates velocity vectors,

contours of temperature, and lhe concentrations of

unburned fuel and oxygen "n an axial plane

intersecting with the fuel noz;le centerline at the

combustor inlet plane. Maximum gas temperatures

occur in the middle of the primary zone. Gas

temperature decreases near the air tnlets and lower

levels of temperature occur in regions close to the

liner walls due to the high tzmgential velocity of

the injected air. Figure 8 al_,o shows the contours

of the concentration of unbulned fuel vapor and

oxygen. The distributions of these two quantities

are also quite consistent wth the temperature

distribution.

Figure 9 shows velocity vectors, contours of

temperature and the concentrations of unburned fuel

and oxygen in 8 vertical plane midway through the

combustor primary zone. It can be seen that max

temperature and unburned fuel concentration occur

just downstream of the fuel nozzle. Gas

temperature as well as the fuel concentration

decrease in regions between nczzles.

The Lycomin9 Combustor

A 20 degree sector of the combustor was

selected for modeling and divided into 38x33x35

finite difference nodes along axial, radial, and

circumferential directions, respectively. In order

to concentrate calculation cells near the tnlets, a

non uniform grid spacing was Lsed in the analysis.

Figure 10 shows a plot cf predicted velocity

vectors in an axial plane in line with'the injector

centerline. As can be seer the fuel injector/

swirler that has a swirl number of 0.8 produces a

large on-axis recirculation 2one that sweeps back

burnt gas products to ignite the incoming

reactants.

Figure 11 shows the cor_tours of temperature

and fuel concentration in l_ne with the injector

centerline. Maximum gas temperatures occurs toward

the end of the primary zone. The secondary jets

that are in line with the injector cause a

reduction in gas temperature in this plane. Lower

levels of gas temperature occur in regions

surrounding the air inlets which is marked by low

level of fuel concentration.

_pl¢llture

r-

_el Hall Friction

FIGURE 11 - IYCONING CONBUSTOR AXIAL CONTOUI_3

:_

i":.N3 ,..... ,,,.

' _:._ _ -...._

FIGURE 12 - LYCONING COMBUSTOR TRANSVERSE VELOCITY VECTORS

(

/

=...

FIGURE 13 - LYCOMINQ CONBUSTOR TRANSVERSE TEMPERATURE CONTOURS

ORIGINAL PAGE. IS

OF POOR QUALITY
Figures 12 and 13 illustrate the predictions

at the axial plane of the primary zone holes.

Figure 12 displays the velocity vectors while

Figure 13 depicts the temperature contours. It can

be seen from these two figures that the primary

Jets are well impinged and bring the gas

temperature down rapidly. Gas temperature

decreases In regions between injectors due to

primary zone jets and lower gas temperature occur

in regions close to the inner wall of the liner.

(a) L¥comlng Conventional Coe_stor (Side VlewJ

P,F. - 0.27t

(b) VI_T Combustor (Isometric V_ew)

P.F. - 0.1]1

kverage radial

_max. - T4, avq. pKof£1e

P_tte=n Factor _T 3

a 5000 I

; 4000 '__

H ...........

_ 3000

EO00 --_ _ LycomtnvCo_vlntionBI ¢oed_ntoc

lOGO

0

0 l 2 3 • 6

fUt_81 Oistanoo. Am

FIGURE 15 - AXIAL DISTRIBUTION OF SECTOR AVERAGE TEMPERATURE3

,6| •

14 _-- ji_1_ Ll_som|Js"J CoamUonal C0mbustOf q

. 12 _-'--; . _ VJItT bbuitor

lO '_, ._

_ 0 '" •

z .....

0 ! 2 3 4 S

Imial Dist_aoe. /a

FIGURE 16 - AXIAL DISTRIBUTION OF SECTOR AVERAGE NOx

FIGURE 14 - TEMPERATURE CONTOURS AT COt48USTORS EXIT PLANE

5.3 Comparison Between the VRT and Conventional

Combustors

Figure 14 shows the temperature contours at

the exit plane of both combustors. The maximum

exit temperature is 3720 R for the VRT combustor

compared to 4000 R for the conventional combustor.

This corresponds to a pattern factor of 0.131 for

the VRT combustor compared to 0.271 for Lycoming's

conventional combustor.

Figures 15 and 16 show the axial distribution

of the mass average values of gas temperature and

NOx concentration, respectively. It can be seen

from these two figures the strong relation between

NOx level and gas temperature. These two figures

also show that most of the conventional combustor

NOx Is formed in the primary zone where the maximum

temperature occurs. In the case of the VRT

combustor, the NOx is mainly generated in the

secondary zone where most of the reaction occurs

and produces the maximum temperature.

6.0 DISCUSSION OF RESULTS

6.1 Water Analogue Tests

The water analogue tests clearly demonstrated

the predicted flow within the combustor model. In

the primary zone, the flow did one complete

circumferential circuit. This is equivalent to a

swirl angle of about 87 degrees. The preliminary

vector analysis carried out in the design phase,

predicted an angle of 89 degrees. The flow did one

further circumferential circuit in the secondary

and dilution zones before exiting the c_bustor.

This is equivalent to a swirl angle of 85 degrees.

The prediction was 80 degrees in the secondary zone

and 73 degrees in the dilution zone. The close

correlation of the swirl angles also means that the

velocity vectors have to correlate by geometric

similarity.

In the primary zone, there appeared to be a

weak secondary swirl superimposed on the

circumferential swirl. The first combustor design

required an airblast mass flow of 6 percent.

However because of hardware availability, the mass

flow was reduced to 2.8 percent. The rate of

diffusion of the red dye indicated an inadequacy in

the fuel nozzle airblast mass and swirl number.

The higher airblast mass flow would have been a

superior choice.

In the secondary zone, the radial quench jets

set up two strong secondary swirls superimposed on

the circumferential swirl. One rotating

counterclockwise in the upper half of the secondary

zone and the other rotating clockwise in the lower

half. These secondary swirls had been observed in

the can version of the VRT combustor and were the

basis of the predicted quick quench. The secondary

zone dld demonstrate the vigorous mixing required

to go from the rlch burn In the primary zone to the

lean burn in the secondary zone.

lO ORIGINAL PAGE IS

OF POOR QUALITY
6.2 Velocity Fteld

The velocity vectors predicted by the 3-0

computer model can be compared to the flow observed

in the water analogue tests. The typical example

given in Figure 8 shows in sor_e areas exceptionally

good correlation. The predicted primary zone flow

in the circumferential direction is identical to

that observed. The computer model even shows the

inferred weak secondary swirl

In the secondary zone, the two models are not

quite identical. The compu'.er model shows the

circumferential pattern near ':he walls and the two

secondary swirls. However ._t the center of the

secondary zone, the predicted velocity vectors are

now axial which is in disagreement with the water

analogue flow. It would ai)pear also that the

degree of penetration of the radial quench jets is

higher in the water analogue than predicted by the

computer.

In the dilution zone, the two models are

radically different. The com)uter model shows the

flow to be predominantly axiil. Whereas the flow

in the water analogue is predominantly

circumferential. In the vater analogue, the

circumferential flow path continues into the

transition section and can _ven be seen in the

collector. The artificial high numerical diffusion

in the "FLUENT" program could well be the cause of

the axial/circumferential flo_ pattern discrepancy

between the calculated and observed flows.

6.3 Thermal and Reacting Flo_s

The thermal results presented in this report

were arrived at after some modifications to the

input boundary conditions to best depict the flow

fields. In the case of the conventional combustor,

the swirler flow field in the primary zone was

adjusted until the desired recirculation was

obtained. For the VRT configuration, the airblast

mass flow rate was increased from two to four

percent and the inner quench jets moved axially

downstream by 0.I inch. These Jets will have to be

moved further downstream In future work to meet the

intent of the design.

Figure 15 shows the axial distribution of

section average temperatures for the conventional

and VRT combustors. The maximum average

temperature in the VRT combustor is lower and

occurs further downstream than the conventional

configuration. At the exit, the average

temperature of the VRT is nearly 100 degrees R

higher than the conventional. This means that

either the VRT combustor is _ore efficient, or that

the total air/fuel ratios for the two

configurations are slightly eifferent. In spite of

this discrepancy, the pattern factor for the VRT

combustor is predicted to be 0.131 compared to

0.271 for the conventional ccnfiguration.

Figure 14 shows the temperature contours at

the exit planes. In the ccnventional combustor,

the isotherms are randomly located with a fairly

flat radial profile pe_king at 20 percent turbine

blade height. In the VFT configuration, the

isotherms are more organize¢! with a steep radial

profile peaking at 70 percent turbine blade height.

Both conventional and VRT combustor exit

temperatures would be amenable to developmental

adjustments, with the VRT being the more manageable

of the two due to Its more regular isotherm

pattern.

6.4 NOx Emissions

For conventional combustor configurations,

both experimental and analytical data are available

for comparison. The data used must be adjusted for

the actual pressure, inlet temperature and

residence tlme of our combustor. The predicted

emission shown In Figure 4 was obtained from the

empirical relationship derived by Odgers in

Reference I and is based on a large experimental

data base. The data given by Nguyen, Bittker and

Niedzwiecki in Reference 3 are analytical but have

been validated experimentally.

For the conventional Lycoming combustor with a

near stoichiometric primary zone, the predicted NOx

emissions are:

Source NOx ppm EI NOx g/kg

Empirical (Ref 1) 1,040 33

Analytical (Ref 3) 1,076 34

Lycoming Analysis 3,676 117

The empirical and analytical predictions

correlate very well, but the Lycomln9 analysis is

high by a factor of 3.5.

Figure 16 shows the axial distribution of

sector average NOx concentration for the

conventional and VRT combustors. The maximum

average NOx concentration in the VRT combustor Is

nearly half that of the conventional configuration

and occurs in the secondary zone.

The NOx predictions provided by the "CREK"

model (Ref 8) lack credibility when applied to the

state of the art conventional combustor and the VRT

configuration. Possible reasons for the

discrepancy could be:

I, The simple three equation reaction

mechanism. (The analytical model of

Nguyen et al uses 102 reactions.)

2. The assumption that the reactions are

infinitely fast so that chemical

equilibrium is achieved at each node.

3. The decoupling of the NOx chemistry from

the hydrodynamics.

7.0 CONCLUSIONS

* The feasibility of a combustor based on the VRT

principle was demonstrated by both the water

analogue tests and the 3-D computer model. The

flow pattern within the combustor was virtually as

predicted.

* The VRT configuration showed the feasibility of

reducing by half the number of required fuel

nozzles when compared to a conventional design.

The thermal performance analysis conducted by

the 3-D computer model on the VRT configuration

predicted a 50 percent reduction in pattern factor

when compared to a state of the art conventional

combustor.

11
* The results of the chemlcal klnatlcs model w111

require further investigation, as the NOx

predictions did not correlate with the available

experimental and analytical data base.

* Because of the significant benefits which can

derive from the VRT combustor concept and the

encouraging results obtained in this Phase I SBIR

program, NASA is funding a Phase II which would

prove the concept under actual operating conditions

in an Allison engine configuration.

REFERENCES

1,

2.

3.

4,

5,

Melconian, J.O., Editor, "The Design and

Development of Gas Turbine Combustors," Volumes

I and II, Northern Research and Engineering

Corporation, 1980.

Niedzwieckl, Richard W., "HSCT Emissions,"

Presentation for Environmental Impact Oversight

Committee, NASA, Lewis Research Center,

December 20, 1988.

Nguyen, H., Blttker, D., and Niedzwiecki, R.

NASA, Lewis Research Center. "Investigation of

Low NOx Staged Combustor Concept in High Speed

Civil Transport Engines," 25th Joint Propulsion

Conference, Monterey, CA, July 10-12, 1989.

"FLUENT" Manual, CREARE, Inc., Hanover, NH,

1987.

Pratt, D.T. and Wormeck, J.J., "CREK, A

Computer Program Calculation of Combustion

Reaction Equilibrium and Kinetic Laminar or

Turbulent Flow," Report WSU-ME-TEL-76-1,

Washington State University, Pullman, 1976.

12
National Aeronauticsand Report Documentation Page

Space Administration

1. Report No. NASA TM-103176 2. Government Accession No. 3. Recipient's Catalog No.

AIAA-90-2452

4. Title and Subtitle 5. Report Date

Introducing the VRT Gas Turbin.: Combustor

7. Author(s)

Jerry O. Melconian, Abdu A. M,_stafa, and Hung Lee Nguyen

9. Performing Organization Name and Add ess

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington. D.C. 20546-0001

6. Performing Organization Code

8. Performing Organization Report No.

E-5554

10. Work Unit No.

537-01- i 1

11. Contract or Grant No,

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Prepared for the 26th Joint Propulsion Conference cosponsored by the AIAA, SAE, ASME, and ASEE, Orlando,

Florida, July 16-18, 1990. Jerry :). Melconian, SOL-3 Resources, Inc. Reading, Massachusetts 01867; Abdu A.

Mostafa, Textron Lycoming, Stra ford, Connecticut 06497; Hung Lee Hguyen, NASA Lewis Research Center.

16. Abstract

An innovative annular combustor configuration is being developed lor aircraft and other gas turbine engines. This

design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and

providing a major reduction in N()x emission. The design concept is based on a Variable Residence Time (VRT)

technique which allows large fuel particles adequate time to completely burn in the circumferenfially mixed

primary zone. High durability of 1he combustor is achieved by dual function use of the incoming air. The

feasibility of the concept was demonstrated by water analogue tests and 3-D computer modeling. The computer

model predicted a 50 percent reduction in pattern factor when compared to a state of the art conventional

combustor. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. The

results of the chemical kinetics model require further investigation, as the NOx predictions did not correlate with

the available experimental and analytical data base.

17. Key Words (Suggested by Author(s))

Variable Residence Time (VRT); Annular combustor;

Gas turbine engines; Reduced pattt:rn factor; NOx

emission; Water analogue test; 3-I) CFD 18. Distribution Statement

Unclassified- Unlimited

Subject Category 31

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Unclassified Unclassified 14

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