jet-a reaction mechanism study

7/23/2019 Jet-A Reaction Mechanism Study

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E 6 a 7 9

NASA

VSCOM

Technical M emorandum 104441

echnical Report 91- C- 029

AIAA-91-2355

Jet-A Reaction M echanism Study

for Com bustion A pplication

Chi-M ing Lee and Krishna Kundu

L ew is Research Center

Cleveland Ohio

and

W aldo Acosta

Propulsion Directorate

U.S. A rm y A viation Sy stems Com m and

L ew is Research Center

Cleveland Ohio

Prepared for the

27th Joint Propulsion Conference

cosponsored by the AIAA, SAE, ASM E, and ASEE

Sacramento, California, June 24-27, 1991

US ARMY

NASA

SYSTEMS

AVIATION

COMMAND


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J E T -A R E A C T I O N M E C H A N I S M S T U D Y F O R C O M B U S T IO N A P P L IC A T I O N

C h i -M i n g L e e a n d K r ish n a K u n d u

National Aeronautics and Space Administration

L e w is R e s e a r c h C e n t e r

C l e v e la n d , O h i o 4 4 1 3 5

and

W a ld o A c o s t a

P r o p u ls i o n D i r e c t o ra t e

U . S . A r m y A v i a ti on S ys t e m s C o m m a n d

L e w i s R e s e a r c h C e n t e r

C l e v e la n d , O h i o 4 4 1 3 5

ABSTRACT

Simplified chemical kinetic reaction mechanisms for the combustion of Jet A fuel are studied. Initially

40 reacting species and 118 elementary chemical reactions were chosen based on the literature review

j

Analysis Code, 16 species and 21 elementary chemical reactions were determined from this study. This

mechanism is first justified by comparison of calculated ignition delay time with available shock tube data,

then it is validated by comparison of calculated emissions from plug flow reactor code with in-house flame

t u b e d a t a .

I N T R O D U C T I O N

A suc cess fu l mode l i ng o f combus t i on and em iss ions o f gas tu rb i ne eng ine comb us to rs requ i res an

adequate description of the reaction mechanism. For hydrocarbon oxidation, detailed mechanisms are only

available for the simplest types such as methane, ethane, acetylene, ethylene, and propane. 1,2

These

detailed mechanisms contain a large number of chemical species participating simultaneously in many

elementary kinetic steps. Current computational fluid dynamics (CFD) models involve chemical reactions,

turbulent mixing, fuel vaporization, and complicated boundary geometries, etc. To simulate these conditions

requires a sophisticated computer code, which usually requires a large memory capacity an take a long

time to simulate. To get around these problems, the gas turbine combustion modeling effort has frequently

been simplified by using a global approach that reduces chemistry to the specification of and overall global

reaction mechanisms, which can predict quantities of interest: heat release rates, flame temperature,

em iss ions , and i gn it ion d e lay t im e .

The s imp les t Je t -A reac t i on m echan i sm i s t he one -s tep m echan i sm:

C n H m

+ n

+410

2

- - > n C O

2

+ 2 H

2

O

1 )

w h e r e t h e c o e f f ic i e n t s n , m a r e t h e c a r b o n t o h y d r o g e n r a t io . T h e a d v a n t a g e o f th i s me c h a n i s m i s it s

s imp l ic i t y ; i t i nvo lves the so lu t ion o f the conse rva t ion e qua t ions fo r un burne d fue l an d m ix tu re f r ac t i on ,

the heat release and other species concentrations are obtained from linear functions of the amount of

fuel consumed. This mechanism, however, fail to predict the important characteristics of Jet-A oxidation,

i.e., the formation of intermediates and CO. As a result, this mechanism is overpredict the heat of reaction,

hence higher adiabatic flame temperatures.

A slightly more complex mechanism is the two-step mechanism proposed by Edelman and Fortune:3

C n H m + ( _

2+-E)

102->nCO+

2

H 2

O

2 )


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C O + 2 0

4 C O

3

his involves one global reaction describe the formation of CO and H

2

O, and a second global reaction

describe the formation of CO2

. However the formation of intermediates is still ignored and so this mecha-

nism cannot predict the time delay between the initial disappearance of fuel into intermediates and a

significant rise in temperature.

The objective of this study is to define a mechanism that can explain most of the observed phenomena

in our flame tube experiment. The proposed mechanism involves 16 species and 21 elementary reactions.

The initial breakdown of the fuel molecule has been assumed to be the reaction of the fuel molecule with

oxygen; the chain carriers are CH

2 , 0 and OH radials, assumed Jet-A structure is C

13 H 27

nitiation:

C 1 3

H

27 +0 - - ) 13C H

+H 0

4

hese important steps in the chain propagation are:

C H

+0 ^ C H

0 +0 5 )

M + C H

2O - - ) C O + H

2

(6 )

0+H2 > O H +H (7)

The species CH

2

has been considered here as a representative of unburned hydrocarbon fragments.

The importance of this specie increases with increase in fuel concentration. The above reaction steps

have been combined with the existing mechanism of hydrogenair oxidation reported by Nguyen and

Bittker,

4some reaction rates were replaced by more recent values reported by Miller.

5

The activation

energy used for Jet-A oxidation was close to the value reported by Freeman.

6

The proposed mechanism

is listed in Table 1.

The proposed mechanism was first examined through a sensitivity analysis with the use of in-house

Sensitivity Analysis Program Code, the orders of importance for the species of interest and classification

of reactions in descending order of importance are determined. The resulting mechanism was then

validated by calculated ignition delay time with experimental ignition delay time. Then using this mechanism

to calculate results from plug flow reactor code were verified with in-house experimental flame tube data.

E XP E R I M E N T A L A P P A R A T U S A N D P R O C E D U R E

T e s t F a c i l it y

The combustor was mounted in Stand 2 of the test facility CESB located in the Engine Research

Building building 5) at NASA Lewis. Tests were conducted with combustion inlet air pressure ranging

up to 16 atm with the air indirectly heated to about of 811 K (1000 F). The temperature of the air was

controlled by mixing the heated air with varying amounts of cold by-pass air. Air flow through the heat

exchanger and by-pass flow system and the total pressure of the combustor were regulated by remotely

controlled valves.

T e s t R i g

The high pressure and temperature test section used in this experiment consisted of an inlet section

fuel injection and vaporization section, flame holder, and combustion section. The combustion test rig is

illustrated schematically in Fig. 1. The flow area is square having an area of 58 cm

2(9 in.

2

). The premixing

and vaporization section, and the combustion section were 27 cm (10.5 in.) and 74 cm (29 in.) long,

respectively. The rig is designed to allow changes in the mixing and vaporization lengths.

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Fuel Injector

Jet-A fuel was introduced into the airstream by means of a multiple-passage fuel injector shown in

Fig. 2. The fuel injector was designed to provide a good dispersion of fuel in the air stream by injecting

equal quantities of fuel into each of the individual air passages. The injector used in these tests had

16 square passages. Each passage was machined to form a converging diverging flow path. The

64 percent blockage helped to insure a uniform velocity profile over the duct cross section. The pressure

drop ranged between 3 and 6 percent of the inlet pressure.

Fuel was discharged from 16.5 cm 6.5 in.) long, 0.7 mm 0.027 in.) diameter tubes into the converging

upstream end of each of the air passages. The fuel tubes were routed through a 0.32 cm (0.125 in.)

diameter feedthru hole. The feedthru holes were routed through a plenum, and the plenum was air cooled

to prevent the fuel from heating and coking within the tubes. The cooling air was discharged into the main

airstream. The cooling air amounted to about 5 percent of the total air flow.

Flame Holder

The flame holder assembly is shown in Fig. 3. The flame holder is a water-cooled perforated plate.

The flame holder was made by brazing 36 tubes of 0.63 cm 0.25 in.) inside diameter between two cooper

nickel beryllium alloy plates. This resulted in an open area of 20 percent of the inlet duct cross-sectional

area. The total pressure drop across the burner ranged from 5 to 12 percent of inlet air pressure depending

on the operating conditions.

Test Section

The water cooled combustion section had a square cross-section like the inlet section and was

74 cm (29 in.) long, because of availability. At the downstream end quench water was sprayed into the

gas stream to cool the exhaust. A cross section schematic of the combustor is shown in Fig. 4. The

flow path was casted in place by using a high temperature castable refractory material. A high tem-

perature insulating ceramic fiber paper was placed between the hard refractory material and the stainless

steel water cooled housing. The paper served two purposes, first to reduce the heat loss and minimize

cold-wall effects, and second to compensate for the difference in thermal expansion between the ceramic

and the housing.

Instrumentation

The combustion gases were sampled with six water-cooled sampling probes located at the axial

positions shown in Fig. 1, 10.2, 3 0.5, and 5 0.8 cm 4, 12, and 20 in.) downstream of the flame holder. There

were two probes at each axial location, 1.57 cm (0.62 in.) from the center line. The probes were 1.57 cm

(0.62 in.) in diameter with five 1 mm (0.040 in.) diameter sampling tubes manifolded together. Remotely

operated solenoid sampling valves permitted the selection of the sample gas from one probe at a time.

The probes were mounted on pneumatic operated cylinders interconnected with the solenoid sampling

valves so that only one probe was in the airstream at a time.

In addition to gas analysis, pressure and temperatures were measured along the test rig. At the exit

of the bellmouth, a rake containing five total pressure probes and a wall static tap were used to determine

the air velocity profile. The inlet temperature was measured with two thermocouples at the inlet to the rig.

Pressure and temperature were also measured upstream of the flame holder to determine the presence

of upstream burning and the fuel injector pressure drop. The temperature of the combustion gases was

measured using a Type B thermocouple located approximately 40.6 cm (16 in.) downstream of the flame

holder. A pressure tap at the exit of the combustor was used to calculate the combustor pressure drop.

The fuel used for this work is specified by ASTM Jet-A turbine fuel disignation. This is a multicompone nt

kerosene type fuel commonly used in gas turbine engines. Jet-A with a HJC ratio of 1.96, was metered

to the reactor from a pressurized fuel tank. Flow rates measured with a calibrated turbine flow meter were

varied from 0.1 to 4.0 GPM, depending on the equivalence ratio desired.

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Standard procedures were followed for each run. These included a warm-up of at least 2 hr with 1000

to 1100

F hot air to the desired test conditions. This procedure assure steady-state temperature in the

reactor. After the reactor reached a steady-state temperature, start-up was initiated by adding fuel to the

hot air and igniting the mixture with a spark igniter. Gas samples were drawn sequentially from one of the

six probes, sample gases then were passed through the following analyzers: nondispersive infrared carbon

monoxide, carbon dioxide, and hydrocarbon units, a chemiluminescent nitrogen oxides unit, and an

electrochemical oxygen unit. Each analyzer unit was zeroed and calibrated with known concentration gas

prior to test run.

COMPARISONS OF PROPOSED MECHANISM WITH EXPERIMENTAL DATA

The worth of any reaction mechanisms is determined by its ability to predict experimental data from

various sources. This section evaluates the proposed Jet-A mechanisms with chemical equilibrium

calculation, ignition delay times, and in-house flame tube experimental data.

Equilibrium Calculation

The combustion mechanism we started with had 118 reaction steps and 40 reaching species, but it

could be divided into three parts (1) oxidation and breakdown of the fuel; (2) hydrogen-oxygen reaction;

and (3) oxidation of carbon monoxide. To reduce the size of the mechanism, the important reaction steps

were computed by senisitivity analysis. Normalized sensitivity coefficients were computed using decoupled

direct method reported by Radhakrishnan.7

In the present work sensitivity coefficients of several species concentrations plus temperature and

pressure were used to determine important reactions.

The predictions of sensitivity calculations were tested by indirect methods. The rate constants for

individual reactions were changed and the ignition delay calculations were repeated. Using this technique,

a few steps which were not very important in the fuel-lean combustion, were eliminated.

This mechanism was further tested by comparing the computed combustion temperature and the

concentrations of different species with those obtained by using chemical equilibrium code.

8Table 2 shows

that the predictions of temperature and species concentration by using present mechanism agree very well

with the results from chemical equilibrium calculation. The proposed mechanism has reduced to 16 species

and 21 reaction steps.

Ignition Delay Time

The ignition delay time was defined at those corresponding to the advent of significant increase in

temperature and pressure. Figure 5 shows the calculated ignition delay time for Jet-A and air is 36 msec.

This calculation is performed by in-house shock tube code integrated with the proposed mechanism. The

experimental data of Jet-A ignition delay times were taken from Freeman and Lefebure's

6

work for equivalence

ratio of 0.5. Figure 6 shows very good agreement between computed results and experimental data.

Flame Tube Experiment

Jet-A fuel has been studied over the equivalence ratio range 0.471 to 0.588 (F/A=0.032 to 0.040),

with inlet air maintained at 1000 F (810 K). Adiabatic flame temperature ranging from 2940 to

3265 F (1889 to 2069 K).

The Jet-A fuel is pre-mixed with air and prevaporized, so that transport effect can be neglected. The

amount of fuel injected is less than 1 percent on a molar basis, and the inlet air flow is highly turbulent,

thus, the effects of longitudial diffusion of mass and energy are negligible. The reactor is insulated with

ceramic material, as a result, the reactor can be characterized as one-dimensional adiabatic plug flow

reactor.

4


6/14

The concentrations of CO and CO

2

were recorded at three probe locations; the temperature was

recorded at a location between probe 2 and probe 3. The combustion was practically 99 percent complete

at all three locations, based on emission data.

Since an ignitor was used to start combustion, it was very difficult to identify the zero time of reaction

in plug flow type calculations, we assumed the time of reaction started at the ignitor.

Figures 7 to 9 shows CO, CO

2

, and flame temperature plotted against equivalence ratio. Judging

from these figures, it appears that the experimentally measured CO

2

concentrations were consistently higher

than computed, it is possible that there was an air leak in the system, as a result the actual equivalence

ratio was higher than what used in the computation. The computed flame temperatures was also slightly

higher than the experimental results. It is possible because of an air leak in the system, in addition, the

thermocouples were installed about 1/8 in. into the flame tube wall, it could be affected by boundary layer

temperature. This mechanism explains that carbon monoxide concentration increases with increase in

equivalence ratio, but no quantitative correlation could be found.

ON LUS ON

This work presents the results of fuel-lean combustion of Jet-A with inlet air temperature around

1100

F and pressure around 10 atm. Combustion temperature and concentrations of CO and CO2

at three probe locations have been reported.

A simplified mechanism to explain the experimental results, is also presented in this work. This

mechanism has 21 steps of reactions and 16 reaching species; CH

2

is the only intermediate hydrocarbon

fragment assumed in this mechanism. The equilibrium temperature and the concentration of species

predicted by this mechanism, agrees very well with the results calculated by using equilibrium code by

Gordon and McBride. Good agreement was found between the computed and experimental ignition delay

times measured by Freeman and Lefebure over a considerable range of temperature.

This mechanism satisfactorily computes the in-house experimental combustion temperatures. The

computed carbon dioxide concentrations also compare, satisfactorily with the experimental results. This

mechanism explained the increased carbon monoxide concentration with increase in equivalence ratio, but

no quantitative comparison could be made.

References

1.

Westbrook, C.K. and Pitz, W.J., A Comprehensive Chemical Kinetic Reaction Mechanism for Oxidation

and Pyrolysis of Propane and Propene,

Combustion

Science and T echnology

Vol. 37, Nos. 3-4,

1984, pp. 117-152.

2. Jachimowski, C.J., Chemical Kinetic Reaction Mechanism for the Combustion of Propane,

Combustion

and Flame, Vol.

55, Feb. 1984, pp. 213-224.

3. Edelman, R.B. and Fortune, O.F., A Quasi-Global Chemical Kinetic Model for the Finite Rate Combustion

of Hydrocarbon Fuels with Application to Turbulent Buming and Mixing in Hypersonic Engines and

Nozzles, AIAA Paper 69-86, Jan 1969.

4.

Nguyen, H.L., Bittker, D.A. and Niedzwiecki, R.W., Investigation of a Low No

X

Staged Combustor

Concept in High Speed Civil Trasport Engines, AIAA Paper 89-2942, June 1988. (Also, NASA

TM-101977).

5.

Miller, J.A. and Bowman, C.T., Mechanism and Modeling of Nitrogen Chemistry in Combustion,

Progress in Energy and Combustion Science, Vol.

15, No. 4, 1989, pp. 287-338.

6.

Freeman, G. and Lefebure, A.H., Spontaneous Ignitition Characteristics of Gaseous Hydrocarbon-Air

Mixtures,

Combustion and

Flame

Vol. 58, Nov. 1984, pp. 153-162.

7.

Radhakrishnan, K., Decoupled Direct Method for Sensitivity Analysis in Combustion Kinetics, NASA

CR-179636, 1987.

8.

Gordon, S. and McBride, B.J., Computer Program for Calculation of Complex Chemical Equilibrium

Compositions, Rocket Performance, Incident and Reflected Shocks, and Chapman-Jouget

Detonations, NASASP-273, 1971.


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T A B L E 1 . -T H E P R O P O S E D J E T - A K IN E T I C M E C H A N IS M *

E

6098.

16400.

13750.

-1000.

96000.

118020.

2126.

1070.

0 .

41000.

-758.

22930.

0.

41380.

45500.

0 .

0 .

-479.

42000.

9000.

14595.

A

B

H2

+

OH

=

2O

+

H

4.74E+13 0 .

H + 02 =

H +

0

1.89E+14

0 .

0 +

H2 =

H +

H

4.20E+14 0 .

H +

02

=

02

+

M

1.46E+15

0 .

THIRDBODY

N2 3.0 02

.3

H2O

21.3

ND

M +

H2

=

+

H

2.20E+14

0 .

M +

02 =2.00

1.80E+18 -1.

H

+

H02 =

2

+

02 2.20E+14 0 .

H02 +

H =2.00H

4.24E+14

0 .

H02 +

OH

=

2O +

02 8.00E+12 0 .

CO

+

02

=O2 +

0

1.60E+13

0 .

CO

+ OH

=

O2

+ H

1.51E+07

1.3

CO

+

H02 =

O2

+

OH

5.80E+13 0 .

N +

NO =

2

+

0

3.27E+12

0.3

0

+ NO

=

+

02

3.80E+09

1 .

02 +

NO

=

O2

+ 0

1.00E+12 1 .

N

+ OH =

O +

H 3.80E+13 0 .

H

+ NO2 =

O

+

OH 3.00E+13 0 .

H02 +

NO =

O2

+

OH

2.11E+11

0 .

02

+

C 1 3 1 - 1 2 7

>13.CH2

+

H02

6.00E+14

0 .

CH2

+

02 =

H2O

+

0

2.00E+13

0 .

M

+

CH2O =

O

+

H2

2.50E+14 0 .

* Forward

reaction

rate constants

expressed as

T

xp(-E/RT).

A= Frequency

factor

(cm-mol-s units)

B=

Temperature

coefficient

unitless)

E=

Activation

Energy

cal/mol)

T A B L E 2 .- -C A L C U L A T E D R E S U L T S F R O M

E Q U IL IB R I U M C O D E A N D F R O M T H E

P R O P O S E D M E C H A N I S M

S p e c i e s

Calculated b y Calculated b y

proposed

equilibrium

mechanism

c o d e

(t im e = 1 s e c )

C O

82

p p m

7 7 p p m

C O

2

6 8 per cent

6 8 per cent

H 2

18 p p m 18 p p m

Tempera t ure

1985 K

1 9 78 K

I n i ti a l m i x tu r e :

E q u i v a l e n c e r a t io : 0 . 5 1 , P = 9 .5 3 a t m , T i n = 84 1 K

6


8/14

O

E

(0

4)

70

C

O

C

C

O

U

U

B

C

E

a )

m

io

O

O_

N

N j

E

C 5

C

Q

E

(0

O

C

O

C

C

O

R3

U

U

75

C

O 5

E

E

L

x

a)

N

rn

O

I I

2

0

Z

W

J

W

r

_ X

Ir

Q

Q

W

O

U -

0

W

F --

}

W

0

Z

O

H

Z

0

0

O

W

Z

Q

f- -

Z

O

L L

V

W

M ^

W

Q

^W

LL

O

`29

COOr-

no

9 ar cvrnrnT

T

LO O f rl 0 0 N

T T T

Ot1) ( C7 C7 t

a) 0

2N^t

c v D

Gin

rn rn CD 0

0 0

N 9 ON V N O

r- O N LO O CO

T T T T

E U) O CO O) Lo

93 C

l)

V ct

to

r-

O

I I

2

0

Z

NmTcn

O O

t

NTOCOr-

Y)

T T T

r N OO

C7

It

L 0 CO

r

l

-

NCOLO(OCn

N

T- Cl) rNN

2 C

I I

v

rl-Omr- NOCn

ON n NN m

co ti O to ct C7

Cnrnrnrn

C)

C

7


9/14

min

M

in.

+

3 in.

G A S S A M P L I N G

P R O B E ( 6 )

F U E L / A I R S A M P L I N G P R O B E

I W OU T )

(TRAVERSE)

A D V A N C E D D I A G N O S T I C S :

A S E R I N D U C E D

_ - - - -

_------------------------

I N L E T P L E N U M

V

rm

I

FUEL INJECTOR

L O C A T I O N S

L A M E H O L D E R

C O M B U S T I O N

F i g u r e 1 . H i g h p r e s s u r e a n d t e m p e r a t u re s q u a r e w a v e f la m e t u b e .

.+ i 1 in. -.L.

in.

,

.25 in.

F i g u r e 2 . M u l ti p le t u b e f u e l i n j e c t o r .

8


10/14

Figure 3.Water cooled flame holder.

Figure 4.Combustor cross section.

' i pe

f le S i C

J e t A + a i r , e q u ' n a l e n c e r a t i o = 0 . 5

E x p e r i m e n t o

C a lcu la t e d

25

2000

a

Y 1500

a i

m

E 1000

a ^

H

500

0

0

0

20

60

00

75

0 0 0

0 2 5

0 50

0 7 5

T i m e , m i l li s e c

/ te m p e ra tu re , 1 / K

Figure 5.Ignition delay time for Jet A + air.

i g u r e 6 . S p o n t a n e o u s i g n it io n d e l a y t im e s f o r J e t A - a i r.

9


11/14

250

200

E 150

Q _

Q

O

U

100

50

0

0.46 0.48 0.50 0.52 0.54 0.56 0.58 0.60

Equivalence ratio

9 . 0

C

8 . 5

a

C 8 . 0

0

0

7 . 5

a ^

p

7 . 0

E

c ,

j

6. 5

O

U

6 .0

0.46 0.48 0.50 0.52 0.54 0.56 0.58 0.60

Equivalence ratio

3 . 4

C O

. 3

o_

X 3.2

Y

3 . 1

cis

Q 3.0

E

~ 2 9

2 . 8

0 . 4 6 0 . 4 8 0 . 5 0 0 . 5 2 0 . 5 4 0 . 5 6 0 . 5 8 0 . 6 0

Equivalence ratio

F i g u r e 7 . -E x p e r im e n t a l a n d c a l c u l a t e d s p e c i e

c o n c e n t r a ti o n s a n d t e m p e r a tu r e s f o r J e t A

o x i d a t i o n a t t h e p r o b e 3 l o c a t io n .

400

350

exp

300

alc

E250

C L

a

200

U 150

100

50

0

0.46 0.48 0.50 0,_2 0.54 0.56 0.58 0.60

Equivalence ratio

8 . 5

c

c

i

.0

a

a .

7 . 5

o

7 . 0

a >

o

E. 5

N

O

6 . 0

0.46 0.48 0.50 0.52 0.54 0.56 0.58 0.60

Equivalence ratio

3 . 5

C O

.4

o _

X 3 3

Y

E 3 2

it s

CL 3.1

E

. 0

2 . 9

0 . 4 6 0 . 4 8 0 . 5 0 0 . 5 2 0 . 5 4 0 . 5 6 0 . 5 8 0 . 6 0

Equivalence ratio

F i g u r e 8 .- E x p e r i m e n t a l a n d c a l c u la t e d s p e c i e

c o n c e n t r a ti o n s a n d t e m p e r a t u r e s f o r J e t A

o x i d a t i o n a t t h e p r o b e 2 l o c a t io n .

10


12/14

140

120

100

E

a

n

0

O

U

B0

xp

alc

40

20

0.46 0.48 0.50 0.52 0.54 0.56 0.58 0.60

Equivalence ratio

9. 0

C

8. 5

a ^

CL

8. 0

0

7. 5

a>

0 7.0

E

cv 6 5

O

U

6. 0

0.46 0.48 0.50 0.52 0.54 0.56 0.58 0.60

Equivalence ratio

3 5

3. 4

c^

0

3 3

x

Y

3 2

ai

3. 1

aD

3. 0

2. 9

2. 8

0.46 0.48 0.50 0.52 0.54 0.56 0.58 0.60

Equivalence ratio

Figure 9.-Experimental and calculated specie

c o n c e n t r a ti o n s a n d t e m p e r a t u r e s f o r J e t A

o x i d a t io n a t t h e p r o b e 1 lo c a t i o n .


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0503.

1.

A G E N C Y U S E O N L Y Leave blank)

2 R E P O R T D A T E

3 . R E P O R T T Y P E A N D D A T E S C O V E R E D

Technical Memorandum

4 . T IT L E A N D S U B T IT L E 5 . F U N D IN G N U M B E R S

Jet-A Reaction Mechanism Study for Combustion Application

WU-537-01-11

PE- 1L1622I IA47A

. A U T H O R (S )

Chi-Ming Lee, Krishna Kundu, and W aldo Acosta

7 . P E R F O R M I N G O R G A N I Z A T IO N N A M E ( S ) A N D A D D R E S S ( E S )

8. P E R F O R M I N G O R G A N I Z A T I O N

NASA Lewis Research Center

R E P O R T N U M B E R

Cleveland, Ohio 44135-3191

and

E - 6279

Propulsion Directorate

U.S. Army Aviation Systems Command

Cleveland, Ohio 44135-3191

9. S P O N S O R I N G /M O N I T O R I N G A G E N C Y N A M E S (S ) A N D A D D R E S S (E S )

1 0 . S P O N S O R I N G /M O N I T O R I N G

A G E N C Y R E P O R T N U M B E R

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

an d

NASA TM -104441

U.S. Army Aviation Systems Command

A V SCOM TR - 91 -

C - 029

St. Louis, Mo. 63120

1 7 9 8

11. S U P P L E M E N T A R Y N O T E S

Prepared for the 27th Joint Propulsion Conference cosponsored by the AIAA, SAE, ASM E, and ASEE, Sacramento,

California, June 24 - 27, 1991. Chi-Ming Lee and Krishna Kundu, NASA Lewis Research Center. Waldo Acosta,

Propulsion D i

rectorate, U.S. Army Aviation Systems Command. Responsible person, Chi-Ming Lee, 216) 433 - 3413.

1 2a . D IS T R IB U T IO N /A V A IL A B IL IT Y S T A T E M E N T 1 2b . D IS T R IB U T IO N C O D E

Unclassified -Unlimited

Subject Category 07

1 3 . A B S T R A C T Maximum 200 words)

Simplified chemical kinetic reaction mechanisms for the combustion of Jet A fuel are studied. Initially 40 reacting

species and 118 elementary chemical reactions are chosen based on the literature review of previous works. Through a

sensitivity analysis with the use of LSE NS G eneral Kinetics and Sensitivity Analysis Code, 16 species and 21 elemen-

tary chemical reactions are determined for this study. This mechanism is first justified by comparison of calculated

ignition delay time with available shock tube data, then is validated by comparison of calculated emissions from plug

flow reactor code w ith in-house flame tube da ta.

1 4 . S U B J E C T T E R M S

1 5 . N U M B E R O F P A G E S

Jet engine fuels; Reaction kinetics; Combustion-, Jet engine

1 6 . P R I C E C O D E

A0 3

1 7 . S E C U R I T Y C L A S S I F I C A T I O N

1 8. S E C U R I T Y C L A S S I F IC A T I O N

19. S E CUR I TY CL A S S I F I CA TI O N 2 0. L IM I T A T I O N O F A B S T R A C T

O F R E P O R T O F T H I S P A G E O F A B S T R A C T

U nclassified U nclassified

Unclassified

N S N 7 5 4 0 - 0 1 -2 80 - 5 5 0 0

tandard Form

298 Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102


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S p a c e A d m i n is t ra t io n

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N A S A

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