—AA— SiJso—Oo—
Technical Report
Aircraft Technology Assessment:
Progress in Lp.w Emissions Engine
by
Richard Munt
May 1981
NOTICE
Technical Reports do not necessarily represent final EPA decisions
or positions. They are intended to present technical analysis of
issues using data which are currently available; The purpose in
the release of such reports is to facilitate the exchange of
technical information and to inform the public of technical
developments which may form the basis for a final EPA decision,
position or regulatory action.
Standards Development and Support Branch
Emission Control Technology Division
Office of Mobile Source Air Pollution Control
Office of Air, Noise and Radiation
U.S. Environmental Protection Agency
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Table of Contents
Table of Contents i
Summary .1
Section I. Forward 2
Section II. Introduction .3
Section II. Emissions Control Technology 19
Section IV. Industry Status .39
Bibliography 123
Appendices
A. Control Techniques .A-l
B. Engine Data B-l
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Summary
This report is the third in a series whose purpose is to evaluate
the potential of various control techniques to reduce emissions, to
/
assess the practicality of such techniques on commercially acceptable
engines, and to estimate the time frame in which such techniques can be
made available. As this document is the latest and, as such, reviews
the situation after the greatest development effort has been made, it
reflects more nearly the eventual outcome of the many industry and
government programs now in progress.
This report concludes:
1) NOx control for high pressure ratio engines remains in an
early stage of development with insufficient control of all four (HC,
CO, NOx and smoke) pollutants available and airworthy hardware uncertain.
2) Although substantial reductions in CO can be attained (>70%) ,
compliance with the proposed standard would require greater reductions
'r '
3) Smoke control is well understood, but its control is compromised
somewhat by the control for HC and, especially, CO.
4) HC control to the level required is readily achievable only if
sector burning is permitted for some engines.
5) The control of HC.and CO by sector burning at idle is very
effective, but possesses unresolved problems of reliability which, in
turn, impact the economics and potentially the safety of the engines.
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Section I ;-
/
FOREWORD
On July 17, 1973, regulations controlling the gaseous and smoke
emissions form aircraft engines were promulgated.(3) The fuel venting
requirement and certain specific smoke standards are already in force;
the principal gaseous pollutant standards originally scheduled for 1979
have now been delayed. The interval between the promulgation date and
the compliance date is intended to permit the development of the requisite
technology, and in the case of retrofit, for the orderly installation of
the new hardware onto the in-use engines.
Two previous reports have been issued (4 and 5) which review the
status of the development of the control technology and this report,
the third in the series, constitutes an update to the second. This
update is required to provide technical support data for the compre-
hensive review and revision of the standards that is now underway.(6)
•/** :'
Because this report is only an updating of the previous one (5), it
is abbreviated. It attempts only to correct obsolete data and analyze
particular points relevant to the unanswered issues confronting the
revision of the standards.
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Section II
INTRODUCTION
Section 231 of the Clean Air Act, as amended by Public Law 91-604,
directs the Administrator of the Environmental Protection Agency to:
(1) investigate emissions of air pollutants from aircraft to
determine the extent to which such emissions affect air quality and to
determine the technological feasibility of control, and
(2) establish regulations for the control of emissions from air-
craft or aircraft engines if such control appears warranted in the light
of the investigation referred to above.
Furthermore, the Clean Air Act states that any such regulations can
take effect only after sufficient time has been allowed to permit the
development and application of the requisite technology. '
The EPA has complied with both mandates of the Clean Air Act,
=*"
first, by publishing a report, "Aircraft Emissions: Impact on Air
Quality and Feasibility of Control" (1) , which concluded that the
impact on air quality by aircraft was sufficient to justify control
and that such control was technologically feasible, second, by publish-
ing a report, "Assessment of Aircraft Emission Control Technology" (2),
which offered the best projection at that time of the feasibility of
control with the knox^ledge then available, and third, by promulgating
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standards limiting the emissions from aircraft engines (3).
In keeping with the spirit of the instructions to determine- the
technological feasibility of control and to the time required to permit
the development and application of the technology, the EPA established
an Aircraft Technology Assessment Program for the purpose of monitoring
the many programs for the development of the low emissions technology
for aircraft gas turbine engines. This program for gas turbines was
begun in July 1974 and has produced two previous reports, "Aircraft
Technology Assessment - Interim Report on the Status of the Gas Turbine
Program, .December, 1975" (4), and "Aircraft Technology Assessment -
Status of the Gas Turbine Program, December, 1976" (5).
In March, 1978, the EPA published a Notice of Proposed Rulemaking
(6) offering considerable revisions to the existing regulations (3).
These revisions were based upon reviews of aircraft air quality impact,
economic impact, and technology limitations, the latter being based upon
» •
reference 5 and additional material supplied by the industry, NASA, and
**''" •
the U.S. Air Force in the intervening period between the publication of
reference 5 and the NPRM (6). The NPRM proposes, among other things, to
restrict compliance with the gaseous emissions standards to commercial
jet engines of sufficient size and frequency of operation as to warrant
their control. This report limits itself to the assessment of the
technology involved in those commercial engines which are likely to be
affected by the proposed regulations. Emissions control technology for
engines not affected by the proposed standards for gaseous emissions as
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and II-2 present the existing and proposed standards in both the old,
english unit format and the new, metric unit format.
;:
/
The classes of engines referred to in the standards were estab-
lished by the EPA to categorize the engines according to technical,
economic, and safety constraints. In the proposed standards, the
subsonic jet classifications become less meaningful in that the gaseous
pollutant standards, like the smoke standards, are monotonic and analytic
functions of size (thrust) and are not discontinuous at the class
demarcations. For reference, the classes are listed in Table 11-3.
The emissions levels permitted by the standards are described by an
EPA parameter (EPAP) which is defined in the aircraft regulations.
Briefly, it is a measure of the total emission of a particular pollutant
produced by an engine over a typical landing-takeoff (LTO) cycle
normalized with respect to the useful output of the engine over that
cycle. As such, larger engines performing greater useful work are
permitted proportionally larger amounts of total emissions over smaller
engines. The proposed standards changed the definition of EPAP by con-
sidering the useful output of the engine to be its rated power rather
than, as originally, the total work (integrated thrust times time over
the cycle) or total engergy (integrated power times time over the cycle)
as appropriate to each class. As a result, the standards no longer give
implicit credit to a high idle point (given because a high idle point
increases the useful output term in the demonimator in the calcultation
of the total work based EPAP, thereby lowering the emissions rating
relative to the standards).
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Original Standard (All Engines)
TABLE II-l
Comparison Between Original and Proposed Standards
(English Units)
Newly Manufactured Engine Standards
Proposed Standard (Commercial Only)
1979
Class
Tl
T2,3,4
T5
P2
APU
Size
0-8,000 Ibs
>8,000 Ibs
All (1980)
All
All
Class
Tl
T2.3.4
T5
P2
HC
. 1.6*
0.8*
3.9*
4.9**
0 . 4***
Original
Size
0-8,000
CO
9.4
4.3
30.1
26.8
5.0
Standard
1981
HC
Ibs
>8,000 Ibs 0.4*
All (1984
All
) 1,0*
NOx Smoke
3.7 (1)
3.0 (1)
9.0 (1)
12.9 (3)
3.0
Newly Certified
(All engines)
Size
<6,000 Ibs.
6-20,000 Ibs
>20,000 Ibs
All (1980)
All
All
Engines Standards
Proposed
1981
HC
1984 1981 ( ,
CO NOx Smoke"""
No standard Same
2.1-0-8
0.8
Standard
12.9-4.3 4.0(2) Same
4.3 4.0(2) Same
Same
Deleted Same
Deleted
(Commercial Only) c»
1984
CO NOx
No standard
3.0 3.0
7.8 5.0
No standard
Size
<6,00.0 Ibs
>6,000 Ibs
All
>2,700 HP
HC
0.4
0.4
0.8**
CO NOx
3.0 4.0-,,-,,.
3.-0 4.0
Same
6.4 8.4
(1) SN » 331.8 (Ibs. thrust)' (Presented graphically in original standards).
(2) With additional allowance for PR > 25. .
(3) SN = 300.7 (HP)"0'280
(4) All engines, not just commercial.,
•"*'''»'
* Pounds of pollutant per 1000 pounds thrust-hour per cycle.
** Pounds of pollutant per 1000 HP-hour per cycle,
***Pounds of pollutant per 1000 HP-hour. ' ,
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Table II-2
Comparison Between Original and Proposed Standards
(Metric Units)
Newly Manufactured Engine Standards
Class
Tl
T2,3,4
T5
P2
APU
Original Standard
1979
Size HC
0-36 KN 13.4* .
>36 KN 6.7*
.All (1980) 30.7*
All 0.26**
All 0.24***
(All Engines)
CO NOx
78.9 31.1
36.1 25.2
237.0 70.8
1.43 0.69
3.0 1.8
Newly
Smoke Size
(1) <27 KN
(1) 27-90 KN
>90 KN
(1) All (1980)
(3) All
All
Proposed Standard (Commercial Only)
1981 1984 1981 ,,,
HC CO NOx Smoke ^ '
No Standard Same
17.6-6.7 106.6-36.1 33.0(2) Same
6.7 36.1 33.0(2) Same
Same
Deleted Same
Deleted
»
VO
Certified Engine Standards
Old Standard (All Engines)
Class
Tl
T2,3,4
T5
P2
Size HC
0-36 KN
>36 KN 3.3*
All (1984) 7.8*
1981
CO
No Standard
25.0
61.0
No Standard
NOx Size
<27 KN
25.0 >27 KN
39.0 All
>2000 KW
Proposed Standard (Commercial Only)
1984
HC CO NOx
No Standard
3.3 25.0 33.0
Same
0.045** 0.34 0.45
(1) SN=79 (Rated Kilonewtons) ' (Presented graphically in original standards).
(2) With additional allowance for. PR > 25.
—0 280
. (3) SN = 277 (Rated Kilowatts) (Presented graphically in original standards).
(4) All engines, not just commercial.
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10
*
Table II-3
Summary of Aircraft Classes
Class
Type
Aircraft Application
PI
P2
Tl
T2
T3, T4
T5
APU
Piston engines
(exluding radials)
Turboprop engines
Small turbojet/fan
engines (0-36 KN
thrust)
Large turbojet/fan
engines intended
for subsonic
flight (greater
than 36 KN thrust)
Special classes
applying to spe-
cific engines for
the purpose of
instituting early
smoke standards
Turbojet/fan
engines intended
for supersonic
flight
Gas turbine auxil-
iary power units
Light general aviation.
Medium to heavy general
aviation; some commer—
cial air transport
General aviation
jet aircraft
Commercial subsonic
transports
Commercial subsonic
transports
SST
. r '
Many turbojet/turboprop
transports and business jets
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11
•
It is worthwhile to review the present aircraft emissions situation
in order to give the reader some perspective of the demand that the
>
regulations impose on the industry. The emissions performance of
current production engines is presented first. Table II-4 presents a
list of all production and development engines and for each engine, the
standards to which it must now comply, the proposed levels, the emis-
sions performance in the present production (or baseline) configuration,
the manufacturer, and an estimate of the engine's production potential,
as defined below, which crudely measures the likelihood that the manu-
facturer will attempt to comply. In addition, Figures 1-4 present in
graphical form the same emissions information. The relevant standards
for these engines are, of course, the standards for newly manufactured
engines, not those for newly certified engines as these engines are
either already certified or are expected to be certified prior to the
compliance date for such engines.
Production potential is not usually available in hard figures.
Generally, though, the production of all engines can be grouped intoi ;
four categories for EPA purposes: ^'.
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Table II-4
Summary of Engines and Their Emissions
Class
Tl
Engine
1979 Std.
Production Engines;
TFE 731-2
TFE 731-3
JT15D-4
JT12A-8
CF700-2D
CJ610-2C
ALF502D
M45H
Viper 600
New Engines;
RB401-07
ATF3
ALF502L
T2, T3, T4 1979 Std
Production Engines:
JT9D-7
JT9D-70
CF6-6D
CF6-50C
RB211-22B
RB211-524
SpeySll
(T3)
(T4)
Spey555
JT3D-7
JT8D-9
JT8D-17
Size
0-35.6 KN
15,
16,
11,
14,
19,
13.1
28.9
32.4
KN
KN
KN
KN
KN
KN
KN
KN
16.7 KN
24.6 KN
22.2 KN
33.4 KN
>35.6 KN
205 KN
228 KN
178 KN
222 KN
187 KN
218 KN
50.7 KN
43,8 KN
84.5 KN
64.5 KN
71.2 KN
Mfr****
AR
AR
PWAC
P&WA
GE
GE
LY
RR/SN
RR
RR
AR
LY
P&WA
P&WA
GE
GE
RR
RR
RR .
.RR
P&WA
P&WA
P.&WA''"' ...
HC
Var.
46. 6/* **
21.67*
123/*
47. I/*
97. I/*
159/*
14.8/17.0
162/16.2
156/*
1.9/*
52. 5/*
28.6/15.9
6.7
61.0/6.7
26.0/6.7
43.3/6.7
63.0/6.7
134.6/6.7
110.4/6.7
278.4/12.2
441/13.6
356/7.3
35.1/9.9
. 37.3/8.9
CO
Var.
159/*
129/*
330/*
770/*
861/*
1450/*
112/103
526/97.1
924/*
96. 9/*
153/*
136.0/95.5
36.1
98.5/36.1
87.5/36.1
96.5/36.1
119,5/36.1
172.3/36.1
145.1/36.1
435.8/70.9
420/79.8
294/39.7
124.5/55.9
112.7/49.9
NOx
Var.
43. O/*
52. 6/*
35. 8/*
29. O/*
20. 2/*
25. 2/*
28.8/33.0
31.7/33.0
16.3/*
34. 2/*
37. I/*
32.3/33.0
33.0 + Pres-
sure Corr.
61.8/33.0
54.3/33.0
65.7/33.0 "'
60.8/38.1
51,9/33.4
61.4/34.6 ...
68.1/33.0
49.5/33.0
31.0/33.0
52.2/33.0
60.1/33.0
Sk
Var.
47/38.2
51/37.6
14/41.7
/38.8
24/36.1
33/40
25/32.4
46/31.4
/37.5
/33.8
/34.7
25/32.2
Var.
4/19.3
8/18.8
16/20
13/18.9
10/19.8
12/19
66/27.9
/29.0
/24.4
23/26.2
24/25.5
Production
(1981)
III
III
III
I
II
I
IV
III-IV
I
IV
IV
IV
III
III
III
III
III
I-II
I-II
II
II
-------
Table 11-4 continued
Class
New Engine
(T4)
(TO
Engine
s :
RB410
RB432
RH21 1-535
CF34
CFM56
CF6-32
CF6-80
JT10D-4
JT8D-209
JT8D-217
Size
68.5 KN
71.2 KN
.163 KN
40.0 KN
107 KN
157 KN
213 KN
129 KN
82.3 KN
92.0 KN
M fr****
RR
RR
RR
GE
GE/SN
GE
GE
P&WA
P&WA
P&WA
HC
/19.3
78.9
32.4/6.7
53.1/15.4
12.0/6.7
48.1/6.7
55.0/6.7
/6.7
/7.5
76.7
CO
/52.2
749.9
96.6/36.1
205/85.2
79.5/36.1
102.1/36.1
736.1
736.1
741.2
736.1
NOx
733. 0
733. 0
49.0/33.0
24.9/33.0
42.8/33.0
69.1/33.0
745.2
/***
733.0
733. 0
Sk
725.8
725.5
720.5
20/30.7
722.9
721.1
719.1
721.8
724.5
723.8
Procluc I ion
(1981)
IV
IV
TV
IV
IV
IV
IV
IV
IV
IV
T5 1980 Std All
Production Engines;
OLY593-610 171 KN
RR/SN
30.7
129/30.7
237.0
530/237
70.8
Var.
70.1/70.8 32/25
II
* No Standard to be met
** xx/xx - (Actual performance)/(Proposed NME Standard)
*** Insufficient data to compute standard, probably 33.0.
**** AR = AiResearch
PWAC = Pratt & Whitney Aircraft of Canada
P&WA = Pratt & Whitney Aircraft
GE » General Electric
LY *> LycominR
RR c Rolls Royce
SN = SNECMA
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14
*
Production Category Situation
I Engines already out of production; engines
certain to be out of production by the com-
pliance date for newly manufactured engines.
II Engines at or near the end of their pro-
duction, run by the compliance date. The
few, if any, units produced after that would
not be sufficient to amortize the develop-
ment and certification cost of a low emissions
combustor.
Ill Engines in the broad middle of their pro-
duction run. It is possible to amortize the
necessary development and certification costs
for emissions control over the remaining pro-
duction. It is equally possible to consider
a cost-effective retrofit of the units pro-
duced prior to the compliance date; there
are sufficient units to amortize that develop-
ment and certification costs and to realize
air quality gains.
IV Engines beginning their production run
shortly before or after the compliance date
for newly manufactured engines. There would
likely be insufficient engines built prior
to the deadlines to warrant a retrofit
program.
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15
HYDROCARBON EMISSIONS VS. RATED THRUST-PRODUCTION ENGINES
1000
100
HC
EPAP
(gms/KN)
10
0
0
G
o
D
O
Proposed standard
k JT15D-4
$TFE731-2
&TFE731-3
ALF502L
0 CF34
0 SPEY-555
G SPEY-511
X JT8D-9
&JT8D-17
^ CFM56
ORB211-535
O CF6-32
D CF6-6
QRB211-22B
QRB21 1-524
OCF6-50C
AJT9D-70
50
100
150
200
250
300 350
Rated Thrust (KN)
Figure 1
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16
CARBON MONOXIBE EMISSIONS VS. RATED THRUST -
PRODUCTION ENGINES
1000
0Q
CO
EPAP
(gms/KN)
0
Q
O
100
cr
A
JT15D-4
ATFE731-3
BkALF502D
S ALF502L
0 CF34
0 SPEY-555
OSPEY-511
X JT8D-9
DJT8D-17
ts.JT3D
K CFM56
0RB21 1-535
O CF6-32
D CF6-6
QRB211-22B
OJT9D-7
QRB211-524
OCF6-50C
AJT9D-70
Proposed Standard
10
50 100 150 200
Rated Thrust (KN)
Figure 2
250
300
350
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18
SMOKE EMISSIONS VS. RATED THRUST
PRODUCTION ENGINES
70
60
k JT15D-4
$TFE731-2
ATFE731-3
8&ALF502D
Bl ALF502L
50
EPA
Smoke
Number
40
0 CF34
0 SPEY-555
ClSPEY-511
X JT8D-9
&JT8D-17
fc^JTSD
h CFM56
ORB211-535
QCF6-32
D CF6-6
QRB211-22B
30
20
Proposed standard
ORB211-524
OCF6-50C
AJT9D-70
0
D
10
Q
A
O
o
0
50
100 150
Rated Thrust (KN)
Figure 4
200
250
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17
OXIDES OF NITROGEN EMISSIONS VS.
RATED THRUST -PRODUCTION: ENGINES.
70
O
D
60
O
50
Q
40
NOx
EPAP
(gss/Kn)
Proposed standard
30
0
20
10
0
50
100 150
Rated Thrust (KN)
Figure 3
k JT15D-4
$ TFE731-2
ATFE731-3
&kALF502D -
m ALF502L
^ K45H
0 CF34
0 SPEY-555
Ci SPEY-511
X JT8D-9
DJT8D-17
K CFM56
ORB21 1-535
O CF6-32
D CF6-6
QRB211-22B
OJT9D-7
QRB211-524
O CF6-50C
AJT9D-70
200
250
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19
Section III
EI4ISSIONS CONTROL TECHNOLOGY
There are four pollutants of interest here: the chemical species,
CO; the combination of the species NO + N02, collectively called NOx
(the latter species is the actual pollutant insofar as the formation of
smog and toxicity are concerned, but the former will eventually combine
with atomospheric D£ to form the latter at atmospheric temperature due to
equilibrium chemistry considerations); the collective group of species
of various non-oxygenated hydrocarbons, either raw fuel compounds or
compounds created by the cracking, decomposition, pyrolyzing, or polymer-
ization of the fuel, all of which are simply called hydrocarbons or
HC.
HC and CO are both products of and occur principally under low
power operating conditions conducive to incomplete combustion. ~Lo\
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20
Beyond that:, the low gas pressure and the low fuel flow requirement lead
to poor amortization of the fuel and poor mixing with the air which in
turn cause pockets of excessively ruch (no oxygen for burning) or ex-
cessively lean (too weak to burn) mixture of fuel and air, both of which
will lead to incomplete combustion and hence the emission of unburned or
partially burned fuel (HC and CO). .
Smoke also is a product of incomplete combustion, but its formation
is likely to occur at high temperature and pressure (i.e., high power)
if pockets of excessively rich (insufficient C^) mixture occur. Under
this circumstance, the fuel in the pocket cannot burn (because of lack
of 62), but instead pyrolyzes in the hot, high pressure environment
laving basically microscopic carbonaceous matter, possibly also coated
with a heavy hydrocarbon residue.
NOx, on the other hand, is a product of an unintentional reaction.
which occurs only at high temperatures (i.e., high power) and with ample
Q2> namely the oxidation of nitrogen either from the air itself (the
usual case) or from nitrogen bound in the -fuel. ."Unlike the others;4 bnce
fonned it cannot be consumed as it is. a final reaction. Fortunately, at
the temperatures experienced in gas turbines, the reaction proceeds
relatively slowly so as the gas is exhausted, the NOx levels are usually
well below equilibrium.
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21
•
The control of these pollutants is achieved through various mech-
anisms depending upon the nature of their formation. HC and CO, both
being products of incomplete combustion are often treated together.
However, HC consumption is fast (producing CO and H«0, plus intermediate
radicals) and occurs in the primary zone of the combustor, whereas CO
oxidation is a relatively slow reaction and occurs in the intermediate
zone downstream where additional air has been added. Thus, it becomes
possible and indeed necessary occasionally to consider the two separ-
ately. In any case, control is achieved by enhancing the combustion
rates, increasing the residence time in the environment most favorable
to combustion, or by improving the mixing of the fuel-air mixture to
better utilize the existing potential for reaction. Proper fuel prep-
aration, including thorough atomization and correct spray distribution
is particularly useful in controlling HC, but it may also influence the
CO levels to a lesser degree.
On the other hand, NOx control is achieved by mechanisms which
discourage the oxidation reactions of the fuel bound or atmospheric *
nitrogen. This is achieved by avoidance of high temperatures (>1800°K)
v/ith an ample supply of oxygen. While theoretically simple, such an
approach is difficult to implement practically for such will generally
lead to operational problems (e.g., flame stability) or excessive low-
power emissions or possibly both.
Smoke, while also a pollutant of incomplete combustion like HC and
CO, arises from the presence of a different set of conditions and
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22
•
requires a different means of control. Control is achieved either by
avoidance of the condition which forms smoke (hot and rich), or by
careful tailoring of the airflow distribution in the aft portion of the
combustor so that the particulate matter is consumed after it has
formed.
Table III-l lists all the relevant control techniques for aircraft
gas turbine engines. The techniques are grouped into HC and CO control
on one hand and NOx control on the other. Control techniques can be
classified into four broad categories: (1) operational control, (2)
fuel preparation, (3) airflow distribution, and (4) staging. Water
injection and catalysis are specialized categories which are not of
practical significance at this point. A detailed description of each
control method is also presented in Appendix A.
Each technique differs in its capacity (or effectiveness) to
control each of the pollutants. Similarly each technique has its own
level of sophistication (complexity). Beyond these factors, it is also
imperative to consider the breadth of utility of each control method"
(i.e., the extent of application in the inventory of engines) and its
impact on both the engine and the airframe. For this purpose, a rating •
system is established. The system evaluates each of the criteria
mentioned above on a scale of one to four. The implications of each
number for each of the criteria are summarized in Table III-2. Finally,
in Table III-3 are the actual ratings for the various control methods as
assessed by EPA. In addition, the expected effect of each technique on
smoke production is noted.
-------
HC and CO Control Techniques
Control Technique
Operational
(1) Idle speed
(2) Air bleed
Fuel Preparation
(3) Spray Improvement
(4) Air blast
(5) Air assist
(6) Preraix
Air Flow Distribution
(7) Advanced cooling
(8) Rich primary*-!
Basis
Higher pressure ratio and hotter flame.
Hotter flame and longer residence time.
Atomization and distribution of fuel to
eliminate excessively rich or lean spots
flame.
Atomization and distribution of fuel to
eliminate excessively rich or lean spots
in flame through use of air jets driven
by liner pressure difference.
Atomization and distribution of fuel to
eliminate excessively rich or lean spots
in flame through use of externally sup-
plied air jets.
Distribution and vaporization of fuel to
eliminate excessively rich or lean spots
in flame.
Prevents quenching the CO > C0£ reaction
at the liner wall by cooling air.
Hotter flame for consuming CO.
Comment
Avoids overly rich pockets which create
smoke and HC.
(9) Lean primary*-!
(10) Delayed dilution Longer residency at high temperature.
Staging
(11) Sector burning
This approach provides, in actuality,
spray improvement (3) and a richer
primary (8) at idle without affec-
ting the combustor design at higher
p owe r s.
Fuel penalty, noise, exessive braking.
Fuel penalty, noise, extra valving and
ducting for excess air.
Very simple if effective. Formation of
carbon and smoke must be watched. May
aid in NOx control (See (15)).
Usually must be applied in conjunction
with air flow redistribution in liner
to maintain stoichiometry.
Complex external plumbing and air pump.
u>
Flashback, flame stabilization problems.
Possibly major revision of engine required
to accommodate the necessary geometry.
Has advantages beyond that of lower
emissions.
Causes smoke formation, but has good
flame stability and relight.
Relight and stability problems.
Pattern factor and temperature profile
adjustment difficult.
Fuel penalty at idle due to lower tur-
bine efficiency. Additional fuel con-
trol complexity.
* Stoichiometry refers to high power condition* Rich at high power means near perfect (f/a " .067) at idle.
Lean at high power...means..v.erv lean at idle.
-------
24
While all of these methods have potential application in at least
a few engines, only a few seem to be prominent at this point in meeting
the proposed newly manufactured engine (NME) standards for 1981 and
f
1984. The first is sector burning (method 11) for HC, CO control which
was investigated by General Electric and Rolls Royce for use on the GE,
CF6, and CFM56 (joint manufacture with SNECMA) engines and the Rolls
Royce RB211 engines. In some cases, it alone is inadequate or faulty
and requires additional, minor modifications to the spray or to the
airflow distribution in order to achieve the full emissions control and
proper operational and mechanical performance. In the case of the
CFM56, an increase in the idle speed (method 1), an operational technique,
V
is used. The major performance concerns are fuel control maintenance
and turbine distress due to the sector burning at idle. New injectors,
if needed, may introduce carbon deposition problems^ The principal
advantage of sectoring is the lack of influence of the control upon the
engine and combustor performance in flight. The difficulty with this
technique lies in its mechanical complexity which could adversely affect
reliability and lead to in-flight malfunction (sectoring in flight is;
considered dangerous because of possible engine damage or inability to •
accelerate, depending on the power level).
The.second major approach for HC, CO control is selective azimuthal
burning (method 22). This is closely related to sector burning, but by
dividing the annular into a sufficient number of sectors, the in-flight
hazards of sector burning disappear because there are enough burning
zones to preserve symmetry: in flight operation is then acceptable.
Also, the complexity of the fuel control system and the need for a
failsafe mechanism is removed. It offers less emissions control than
-------
l;iLiny, ol CoiUrul TuchiiL
-------
Table III-3
Control Technique
Idle Speed
Air Blued
Spray Improvement
Air Blase
Air Assist
Premix-1
Advanced Cooling
Rich Primary-1*
Lean Primary-1*
Delayed Dilution
Sector Burning
Selective AZinn thai
Burning
Quick Quench
Rich Primary-2*
Lean Primary-2*
Premix/Prevap
Fuel Staging (1
Variable Geometry
Catalysis
Water Injection
Ref.
No.
(1)
(2)
(3)
(15)
(4)
(5)
(7)
(3)
(8)
(9)
(10)
(11)
(22)
(12)
(13)
(14)
(16)
18)
(19)
(20)
(21)
Control
Capacity
2(HC,CO)
2(HC,CO)
KHC), 2(CO)
4(NOx)
2(HC,CO)
KHC)
2(CO)
2(HC)
3(CO)
3(HC)
2(CO)
2(IIC,CO)
3(HC,CO)
3(CO)
KHC)
2(CO)
2(1IC) .
3(CO)
• 2(NOx)
3(NOx)
2(NOx)
2(HC,CO,NOx)
2(HC,CO,NOx)
2(HC,CO,NOx)
l(HC,CO,NOx)
l(NOx)
Effect
on Smoke
None
None
None
Increase
None
None
Decrease
None
Increase
Decrease
None
None
Increase
Decrease
Decrease
Increase
Decrease
?
, None?
Impact
Complexity
1
1
1-2
2
1
3
3-4
3
3
3
2
1
3
3
3
4
3-4
4
4
.
2
Application
• 1
1
4
3
3
2
2
2
4
4
2
2
2
2
1
2
2 '
2
2
2
Engine
1
1
1
1-2
3
3
1-2 .
2
, 2
4
2
1
1
2
2
3
3-4
4
4
2
Airfrarae
2
3
1
1
2
1
1
1
1
1
1
1
1
1
1
1
2
1-3
4
3
Comment
Limited to higher pressure
rates at idle. May need to
change the stoichiometry.
The external compression is a
difficult mechanical problem.
•
May receive non-emissions
benefit.
In-flight malfunction is a con-
cern.
Safe in-flight; less effective
than sector burning (11).
Low power emissions may be high
Low power emissions may be high
Flame stability questionable,
low power emissions may be high
Flashback is a concern.
May not scale down to small
engine size.
Mechanical reliability in
question.
Not readily avnilnble to
small engines. •
Not practical.
-------
27
does sector burning because it is less able to optimize the stoichiometry
and because of the numerous quenching zones between several burning
zones.
(
The third major control technique- for HC, CO control is the use of
airblast nozzles, (method A) combined with airflow redistribution
(methods 7-10) to achieve the necessary stoichiometry. This approach is
expected on the Pratt and Whitney JT8D, JT9D, and, if built, JT10D
engines. The use of airblast nozzles alone on production type com-
bustors gave inadequate performance because of the alteration of the
stoichiometry and forced the use of additional techniques to tune the
combustor to acceptable emissions and performance (e.g., altitude
relight, durability, temperature profile, etc.). In total, this
approach leads to changes of the stoichiometry and cooling air patterns
in flight which lead usually to increased smoke and degraded combustor
performance.
One of the major features of the NPRM is the required retrofit of
in-service T2 and T4 class engines to achieve compliance with the .1981
newly manufactured engine levels. Unfortunately, installation into in-
service equipment may not be quite as simple to achieve as installation
into newly manufactured equipment because of the need to replace or
modify parts that would be properly installed new on a new engine or
aircraft. Examples would be nozzles, fuel controls, igniters (causing
perhaps a rework of the outer casing), combustor liners, and for sector
burning, squat switches sensing the aircraft on the ground so the fuel
control can distinguish between flight and ground modes because (sector
burning is in general prohibited in flight.
-------
28
For NOx control (NME 1984^, the leading technology is fuel staging,
(methods 17 and 18) due largely to the joint NASA-industry Experimental
Clean Combustor Program (1973-1977) wherein, first, single stage tech-
niques such as lean burning (method 14) we're found inadequate if em-
ployed alone, and second, fuel staging x^as investigated to resolve the
deficiencies of the single stage controls. The latter investigation was
carried through to a technology demonstration in two test engines (not
flightworthy). Two different approaches to fuel staging were inves-
tigated, axial staging (method 17, used on the JT9D-7) and radial
staging (method 18, used on the CF6-50). Application or transfer of
this technology to other engines, even related ones, is not always easy,
however, and has not been pushed by the manufacturers thus far. For
instance, the radially staged combustor developed in the CF6-50, called.
the "Double Annular" can be installed on a CF6-6 only with considerable
modification to the basic engine, including the casing, although it can
be employed directly into the CF6-50 with only direct changes in the
fuel plumbing and fuel control. This is because the double annular
airflow requirement into the double dome calls for a dump type diffuser
from the compressor as is found on the CF6-50. The smooth diffuser
-,"i :
found on the CF6-6 would have to be replaced by a dump type in order to
utilize the double annular idea (Figure 5). Similarly, use of the
double annular concept in the CFM56 is questionable because its much
smaller size (Figure 6) does not lend itself readily to staging in-
volving physical separation of the zones. This separation increases
considerably the S/V ratio and reduces simultaneously the residence
time. When combined with the overall smaller geometry of a smaller
engine, both factors adversely impact the HC and CO emissions.
-------
SIZE COMPARISON BETWEEN THE CFM56 AND Ci'6-6 COMllUSTOUS
-.
.,,^_../" ~~^—*~~~—*jii__' "'"^"^'•''•'-'^v/l-^-^''"r
CFM56
CF6-6
Figure 6
-------
29
COMPARISON OF DIFFUSERS BElWEEN THE CF6-6 AND CF6-50 ENGINES
Diffuser
Diffuser
Double Annular Corabustor in CF6-50
Figure 5
-------
31
Rolls Royce did not benefit in this technology development by
direct association with the program and consequently has not yet engine
tested a fuel staged combustor in an RB211.,S Its recent investigations
have, however, explored the benefits of both radial and axial staging
and it has leaned in favor of the former because of the packaging con-
straints imposed by the relatively short combustors of the R3211 family
(a similar situation to the CF6).
NOx control by fuel staging is in the exploratory development stage
and is not yet ready for final development into specific engines for
certification of airworthiness. Developmental problems for which solu-
tions have not yet been identified are (1) the achieving of all four
emissions goals simultaneously, (2) insufficient cooling air for ac-
ceptable durability performance, and (3) engine performance degradation,
notably transient response. Other shortcomings in the concepts at
present are considered normal for this stage of development and would be
expected to be resolved in due course.
As indicated in the preceding paragraphs, manufacturers have often
found it necessary to combine more than one emission control concept.
This is, in fact, the rule rather than the exception. Such compounding
may be necessary because of the inadequacy of a single control concept
to sufficiently reduce the emissions, or it may be necessary because of
combustor performance deficiencies brought about by the use of a single
control scheme. In the former category would be the combining of
airblast nozzles with liner airflow redistribution. In the latter would
-------
32
•
be the utilization of fuel staging with separate lean and rich primary
zones. Often the rationale is a combination of the two.
f •
The adding on of one control scheme upon another is not guaranteed
to produce a geometric compounding of results; there may be, in fact, no
improvement at all despite the fact that each separately may be quite
effective. For example, while redesigned nozzles and sector burning may
i
individually produce emissions reductions, the first because of improved
atomization, the second because of atomization and stoichiometry, their
joint use is likely to be no better than the sector burning alone be-
cause that in itself already accomplishes that which the redesigned
nozzles purports to do (i.e., improve atomization).
On the other hand, complementary forms of control may achieve
synergetic results: redesigned nozzles combined v;ith advanced cooling
techniques may together serve to reduce emissions that result from dif-
ferent mechanisms of formation (i.e., poor atomization and mixing .in the
primary vs. wall quenching on the liner). It is also possible thatj the
use of only one control will aggravate a condition which will lead &Q
the formation of more pollutants and hence will require two or more
control schemes to balance each other. A significant example of this
situation is the use of airblast nozzles which while providing better
fuel atomization and mixing, also leans the primary zone stoichiometry
by its own airflow. This may result in an excessively lean mixture so
that an airflow redistribution to richen the primary is also required.
Together, the two approaches provide improved atomization, better
mixing, and optimal stoichiometry. Similarly, the conflicting require-
ments of NOx control and HC, CO control require a combination of control
-------
33
techniques, itsost notably through fuel staging in which the two separate
portions utilize control techniques applicable to the particular pol-
lutants which are expected from them.
The above discussion makes repeated reference to the conflict
between emissions and combustor performance and this reference is con-
tinued in Section IV. Therefore, a brief explanation of combustor
performance is appropriate. The criteria by which combustor per-
formance is judged are related to both economic and safety consider-
ations. The economic criteria are created by the users while the safety
criteria are dictated by the users, the FAA, and common sense. Com-
bustor performance itself is two-fold: operational and mechanical.
Operational performance is measured primarily by ground ignition
and engine acceleration, altitude relight, and flame stability (com-
bustor response to engine transients, either intentional or accidental).
Mechanical performance standards are largely determined by economic con-
siderations and the principal criteria are durability, coking, and
carbon deposition. The first two are obviously considerations for the
cost of maintaining the system, but the third may not be so apparent.
Carbon deposition impacts the engine durability and hence maintenance
cost first through the turbine erosion which occurs when particles are
broken off the combustor or nozzle surface and sent downstream, and,
second, by its adverse effect on the combustor cooling (due to a change
in the radiative emissivity).
-------
34
Table III-4 lists the best emissions performance that has been
achieved in each engine. These data are also presented graphically in
Figures 7-9. However, in a few cases the ca~a nay represent combustor
rig data rather than engine data (e.g., CJ&10) or possibly educated
extrapolation from data of a related engine (e.g., RB211-535 figures
originated from RB211-22B data). Most importantly, though, the data may
be iron a configuration which has been found unflightworthy (e.g. Spey
511) or otherwise projected as unsafe. The purpose of this table is
nerely to present in concise form the kind of control that is achievable
by the control methods listed. The standards proposed for each engine
are presented also as a point of comparison. A more accurate interpre-
tation of the situation of emissions reduction is to be found in Section
IV, Industry Status.
-------
Tablu I!1-4
Summary of Boot Kmius ions Performance
Claag Engine
Tl Production Engine!
TFK731-2
TTK731-3
JTI5D-4
JT 1 2 A -8
CK 700-20
CJ6IO-2C
A1.K502U
M45II
Viper600
New F,nf»ine8
ALF502L
Size
15.5 KN
16.5 KN
11.1 KN
1 / t VM
IH * / KN
l'J.2 KN
13.1 KN
28.9 KN
32.4 KN
16.7 KN
rt/ L i»ij
*H . D KN
fj 1 tf ll
ii . i KJf
33.4 KN
IIC CO
4.5/* 59. 8/*
No Data
1.4/* 10U/*
— -— -~**No Tech no 1 o^y-~
14. 7/* 861/*
23. 4/* 1440/*
14.8/17.0 112.4/103
30.1/10.2 170/97.1
NOx
50. 5/*
47. 11*
20. 2/*
25. 2/*
' 28.8/33.0
37.0/33.0
Sk
/38
/38
/42
/36
/40
/32.4
12/31
Date of
Availability
Cert if ied
Jan. 1979
Comment
Kxturii.il air itthiht.
Primary/ injector modi 1 icJit ion.
Nozzle mod i f i rat ion .
Nuzzle moiliri(;il ion.
A t rb 1 ,-ist nozz 1 e/ 1 i..'ij»"r . pr it»i;tr y .
Adv.'ntced cooliiii*. (illit. liiuwti
rint;).
-------No Technology—— -
9.1/15.9 92.2/95.5
35.4/33.0
/31.2
Tnrt IQ7Q
Jun • iyl7
1979
Advanced cool ing*, (dbl . blown
rinp.)
Combustor same as 5021). Idle
at 10.73:. NOx hit-.li.
T2, T3, Production Engine*
T4
JT9D-7
JT9D-70
CF6-6D
CF6-50C
RB211-22B
RD211-524
Spey 511
Spey 555
JT3D-7
JT8D-17
New Knitincs
RI14 1 0
RB432
RB211-535
CF34
CFH56
CF6-32
CF6-80
JT10D-4
JT8D-209
JT8D-217
ti Production Engine*
OLY593-610
205.3 KN
228 KN
178 KN
224 KN
187 KN
218 KN
50.7 KN
43.8 KN
84,5 KN
Me VU
\ * J f*rt
71,2 KN
j> n c uu
68.3 KN
71.2 KN
163 KN
40 KN
107 KN
157 KN
213 KN
129 KN
82 KN
of vtt
92 KN
171 KN
4.5/6.7 24.6/36.1
2.1/6.7 30.2/36.1
4.0/6.7 20.0/36.1
1.8/6.7 28.3/36.1
1.0/6.7 37.1/36.1
2.4/6.7 49.8/36.1
4.2/6.7 28.8/36.1
3.1/6.7 22.4/36.1
23. O/ 162/
36.1/12.2 186/79.8
158/— 232/—
7.6/8.9 49.4/49.9
1.6/8.9 83.1/49.9
8.9/6.7 54.7/36.1
2.5/6.7 '67.5/36.1
12.7/14.4 ' 80.0/85.2
0.-9/6.7 42.0/36.1
2.0/6.7 29.8/36.1
2.0/6.7 /36.1
.-—No Data—--
2.2/7.5 33.6/41.2
<30.7/30.7 <237/237
47. 4/ —
26.2/33.0
4H.5/ —
65.7/~
60. 8/~
44.7/38.7
64.0/--
70. 2/—
68. 2/
55. 2/—
53. O/--
68. 4/—
41.0/33.0
, 51. 3/~
30.3/33.0
27.0/33.0
43.5/—
64. I/--
/
...._
54. 9/—
<70.8/70.
<20/19.3.
30/19.3
<10/18.8
16/20.0
/18.9
/18.9
/19.8
/18.7
/
/29.0
13.3/25.0
/25.5
27/25.5
/21.2
/21.2
20/29.7
/22.9
/21
/19.1
15/24.5
8 725
1986-87
Approocli aban-
doned, perfor-
mance not ac-
ceptable
Approach aban-
doned, perfor-
mance not ac-
ceptable
Jan. 1981
Date of Cert.
Date of Cert.
Date of Cert,
Date of Cert. •
.
Jan. 1980
Aerating nozzle/rich primary.
Low NOx combuntor (Vorhix
«t ;IK itt^) .
Aer;itinn nozy.l e/ r i ''h primary
Sect nr burninj* .
Sector btirn inj'./domi1 nnd nozzle
mod i f i cat ion.
Low NOx cninhiirit or ((>!'}. Anun-
Inr)
Sector burnini; »ml riili pri-
m.iry ( I'luiKe II).
Sec-tor burniiif' .'old ri«h pri-
m.'iry ( Ph;u;e II).
Ueflex airi>pr/iy.
Piloted airblast.
Leaner primary & nirblast;
intended for T3 smoke retro-
fit only.
Aernt ini> nnzzli'/rich primary.
Lox NOx combublor (Vorbix).
Sector burninp, and rich primnrv
(1'lll.SU 11).
Rich primary .iml qui'^ qtn.-nch.
Sector burn inc..
Sector burning at 61 idle.
Sector burnfiiK,
Sector burning,
Aerating noizle/ricti primary.
Blown rlni{.
-------
36
HYDROCARBON EMISSIONS VS. RATED THRUST -
LOW EMISSIONS ENGINES
100
k JT15D-4
£ TFE731-2
6TFE731-3
S&ALF502D •
a ALF502L
^ M45H
0 CF34
0 SPEY-555
O SPEY-511
X JT8D-9
&JT8D-17
HC
EPAP
(gms/KN)
10
h CFM56
• RB211-535
O CF6-32
D CF6-6
QRB211-22B
OJT9D-7
ORB211-524
O CF6-50C
AJT9D-70
Proposed standard
O
Q
A
O
D
0
50
100
150
200
-e-
250
300
350
Rated Thrust (KN)
Figure 7
-------
37
CARBON MONOXIDE EMISSIONS VS. RATED THRUST -
LOW EMISSIONS ENGINES
1000
JT15D-4
TFE731-2
TFE731-3
&ALF502D
HALF502L
0 CF34
0 SPEY-555
Ci SPEY-511
X JT8D-9 .
&JT8D-17
CO
E?AP
(gts/KN)
100
0
h CFM56
• RB21 1-535
O CF6-32
D CF6-6
QRB211-22B
O-JT9D-7
QRB211-524
O CF6-50C
AJT9D-70
Q
Proposed standard
O
O
A
10
0
50 100 150 200
Rated Thrust (KN)
Figure 8
250
300
350
-------
38
OXIDES OF NITROGEN EMISSIONS VS.
RATED THRUST-LOW EMISSIONS ENGINES
70
60
50
40
NOx
EPAP
(gns/KN)
30
20
10
0
50
L JT15D-4
$ TFE731-2
&TFE731-3
^ CF700
SkALF502D •
03 ALF502J.
DkJT8D-17
0 CF34
0 SPEY-555
& SPEY-511
X JT8D-9
b CFM56
0RB211-535
OCF6-32
D CF6-6
QRB211-22B
OJT9D-7
QRB211-524
O CF6-50C
AJT9D-70
Proposed standard
O
O
100 150
Rated Thrust (KN)
Figure 9
200
250
-------
39
*
Section IV
INDUSTRY STATUS
The discussion in this section is limited to those engines which
will be affected by the standards, namely those in commercial service
with a rated thrust of 27KN or greater. Engines in this category are
limited to those made by only four manufacturers, General Electric,
Pratt and Whitney Aircraft, and Rolls Royce (certain engines involve
joint ventures with other manufacturers) and possibly Avco Lycoming.
Each manufacturer and its products will be treated separately.
1. General Electric
General Electric is a large diverse manufacturing company in the
United States. Its commercial aircraft engine operations are located in
Cincinnati, Ohio (CF6, CFM56) and in Lynn, Mass. (CF34, CF700, CJ610).
The CFM56 is a joint venture with SNECMA of France, the core of which is
based upon the military F101 engine designed for the B-l bomber. In " '
addition to these civil engines, GE makes a number of military vari- f'
eties, some of which are essentially the same as .the civil engines. A
summary of the company's civil engines is presented in Table IV-1.
Suianary of Research and Development Effort CF6, CFM56
CF6
General Electric's NOx control effort has centered largely around
-------
Table IV-1
Engine Class Thrust
CF6-50
CF6-6
CF6-32
CF6-80
CFM56
CF34 .
CF700
CJ610
BPR
PR
T2
T2
T2
T2
T2
T2
Tl
Tl
224 KN
175 KN
145 KN
213 KN
107 KN
40 KN
20 KN
13 KN
4.4
5.9
4.8
5.9
6.0
1.9
0
29.8
24.5
24.4
32.0
25.6
19.5
6.6
6.8
General Electric Engines
Combustor Application
A DC-10,B747,A300
A DC-10
A Potentially B757
A B767, A310
A Possibly B707.A300B11
A None
A Falcon 20,Sableliner75A
A Learjet 24/25
Cert.
Date
1981
A
Number
Delivered
0
0
0
0
Product ion
Category
III
III
IV
IV
IV
IV
II
I
PROSPECTUS
Prospects of Meeting;
Engine
CF6-50
CF6-6
CF6-32
CF6-80
CFM56
CF34 .
1981 1984
poor*
poor
poor*
marginal
poor*
marginal
marginal
marginal
marginal
fair
good
good
' Emissions
HC CO
X X
X
X X
X
X X
X
X
X
X
X
*- -• : •
Performance
NOx Sk
X
X
X
X
Operational
Performance
X
X
X
X
Mechanical
Performance
X
X
X
X
•Time
X ,
X
X
X
X
X
X
X
* Assuming sector burning is not used,
-------
41
•
the NASA Experimental Clean Combustor Program (ECCP). This portion of
the jointly funded program with NASA utilized the CF&-50 engine as a
testbed, but it should be considered as a technology demonstration
program having application to annular combustors in general, within
geometric limitations. The ECCP was divided into three phases, the
first being a screening of several concepts involving prefixing, air-
blast, lean primary burning, and fuel staging incorporating some or all
of the previous concepts. The second phase involved refinement of
selected concepts and the third phase was an engine test of the best.
The program is now completed (except for documentation of the final
engine testing). The final concept developed by GE is termed the
double annular combustor and is shown in Figure 10. It is a radial fuel
staging concept and is particularly well suited to GE engines which have
short annular combustors and are, therefore,.less amenable to the
physically, longer axial staging concept (e.g., the Pratt and Whitney
JT9D Vorbix). A detailed description of how such staging reduces
emissions given in Section III and Appendix A.
The engine test in phase III of the ECCP served as both a proof of
concept demonstration (with partial success) and, in fact, went one step
further by attempting to provide flightworthy quality hardware in the
demonstration. Unfortunately, but perhaps predictably, this was at
least partially responsible for the failure of the system to live up to
the expectations of phase II. The specific revisions to the engine
hardware for phase III included a modified inner liner, revised liner
cooling, additional allowance for thermal expansion, and changes in the
-------
DOUBLE ANNULAR COMlHUiTOR CONFIGURED FOR Till' CFO-^O
\J PILOT STAGE
2) MAIN STAGE
ts3
Figure 10
-------
43
•
connections between liner parts. Together these changes led to a more
durable combustor, but one which also suffered from an inferior fuel
injection pattern and degraded mixing which manifested itself through a
deteriorated emissions performance. However, a realistic development
program must eventually address the durability problem and as the phase
II results suggested that the emissions were successfully under control,
turning attention to durability was a reasonable next step.
Current problems with the double annular combustor as presently
configured in the -50 are: (1) high CO - due largely to the hardware
changes incorporated in the engine test to aid durability (the Phase II
rig version had acceptable CO); (2) high NOx - a definite problem
although GE sees some room for improvement; (3) high smoke — due to
inadequate mixing in the robust version xvhich if eliminated would bring
the NOx down, too (the Phase II rig version had acceptable smoke); (4)
durability - mostly an.unknown but the uncooled centerbody is definitely
subjected to a severe environment and the lean front end (primary)
leaves less cooling air for the liner (see Table IV-2); (5) temperature
•/'*" "•
profile - the high air demand of the double burner dome (as seen in
Table IV-2) to run lean in the main burner plus have a separate pilot
burner plus the cooling air requirement for the liner leaves only 2% of
the air left to trim T, instead of the usual 20%-30%, and impacts the
turbine durability); (6) flow control - an altitude compensating control
is necessary to distinquish at equal flow rates between high altitude
cruise (both annuli burning) and approach (pilot only). Any change in
the configuration to improve upon these problems must retain the other
-------
44
expected operational performance levels such as acceleration, relight at
altitude and so forth.
On the positive side of the ledger, however, ground start, altitude
relight, lean blowout, pressure loss, carbon deposition, cruise com-
bustion efficiency, liner wall temperature, and engine acceleration all
meet or appear to meet engine requirements. The exit temperature pat-
tern, although out of specification, could be at least partially due to
the high fuel-air ratio required by the "worn" test engine. This is
presumed because temperature pattern exhibited a hot hub consistent with
an over fueling (by 17%) of the main stage inner annulus. Sufficient
refinement of the pattern may then be possible despite the shortage of
dilution air (Table IV-2) with which to work (the temperature profile
can also be affected by manipulation of the airflow pattern within the
primary zone although this may be anticipated to be deleterious to
emissions). Finally, it should be observed that despite the excess fuel
no liner hot spots or carbon deposition was observed.
Table IV-2
CF6 Airflow Distribution
Dome:
Cooling:
Dilution:
CF6-50 Standard
35%
32%
33%*
CF6-50 Double Annular
Primary 25%"\
Main 51% J
22%
2%
76%
*Not all of this is required for temperature trimming. The CFM56,
for instance, uses about 20%.
-------
45
•
Double annular staging apparently requires single stage operation
at approach as well as at ground idle (both stages at high power) in
order to lower the CO (the use of only one stage creates better at-
onization and nore concentrated burning). This creates the additional
need to ascertain in detail the staging behavior in flight, both in
normal operation and with malfunctioning in the fuel control logic or in
the valving. The single stage operation at approach does contribute to
the high KOx level which is of concern (about 18% of the cycle NOx
arises in the approach mode).
The present design allows no room for significant simplification in
teras of liner configuration, number of fuel nozzles, and manifolding.
Hence, it may be expected that cost estimates based upon the present
configuration are realistic.
A double annular configuration has been designed for the CF6-6
engine (Figure 11) and although differing in detail from the CF6-50
configuration, it is in essence the same, working on the same principl'e
and with similar design parameters. Certain problems have been iden-'
tified with the use of the double annular concept in the CF6-6, however.
Presently, the -6 has a smooth diffuser which is incompatible with the
flow pattern needs of the wide dome of the double annular combustor;
hence, a step or dump diffuser, similar to that employed in the -50 is
needed (scaled to size, of course). Because the diffuser and casing are
integral (compressor rear frame), this change would necessitate a major
redesign of that high pressure shell, at a cost .estimated by.;GE in the
neighborhood of 40 million dollars, if pursued. The ensuing changes in .
-------
46
•
the external dimensions, although not large, would impact the nacelle
packaging of the components external to the engine casing, such as the
fuel manifold, compressor bleed, etc. Another problem is that the large
fuel nozzles required by the double annular concept do not fit readily
in the -6 which now uses smaller nozzles. Beyond requiring larger holes
in the casing, the larger nozzles have a significant effect on the. flow
pattern and pressure drop across the combustor. For these reasons,
direct scaling of the -50 design is not considered feasible, and addi-
tional development work specific to the -6 would therefore be necessary,
even if all the deficiencies identified in the -50 program were remedied.
The CF6-32 is a clipped fan CF6-6 having the same core and hence
the same combustor, and although operating at different conditions, it
is expected that the -6 design will be adequate in the -32. GE has not
yet committed to testing the -6/32 configuration even in combustor rigs.
GE is actively continuing its double annular technology develop-
ment, though, in the NASA Energy Efficient Engine (E ) Program. This
;."*i "
is also a technology demonstration program (not a prototype development
effort) in which it is hoped to demonstrate more efficient components
and engine cycles on an engine assembled from these new components and
sized to the 110-130KN range. Thus, this development does not directly
relate to the CF6 development and, in fact, the GE combustor configur-
ation and size are more related to the CFM56. Nonetheless, the tech-
nology improvements should be relevant. Principal advances to be ex-
plored are (1) single fuel injector stem per nozzle pair, (2) cooled
-------
47
COMPARISON BETWEEN PRODUCTION AND
LOW NOx COMBUSTORS IN THE CF6-6
Production
Double Annular
Figure 11
-------
centerbody (separating the stages), (3) advanced cooling, and (4)
better NOx and smoke control. The best airflow pattern refinements for
emissions and operational performance will be sought.
GE has also pursued independent investigation of simpler concepts
through their IR&D funding. These concepts include compressor bleed,
advanced idle, sector burning, nozzle modification, and liner redesign
(airflow redistribution). These approaches are directed at HC and CO
control only and despite their relative simplicity, they can be quite
effective. Much of the preliminary investigation was done on the
developmental CFM56 (F101) as the early CF6 effort was devoted prin-
cipally to the NASA ECCP.
GE elected to continue development of the sector burning concept in
the CF6 for the proposed 1981 requirement. The governing principles
behind the sector burning concept are described in Section III and
Appendix A. Although sector burning creates a fuel control problem with
its staging at idle, a very desirable feature is that during proper .
operation, there is no effect of the emissions control on the combustor
-."*« ;
in flight so concerns about operational and mechanical performance are
considerably reduced. Used alone, this method is sufficient only in the
CF6-6 and probably CF6-32 engines. The CF6-50 and -80 would still
suffer from high CO because of their very short combustors. CO, in
fact, tends to be a problem in all of the GE engines due to their short
combustor designs which allow inadequate time for its oxidation to C0_.
A short combustor is pursued because it.requires less liner cooling air
(hence more is available for radial, temperature distribution trimming
and for turbine cooling) and it creates a shorter and hence lighter engine.
-------
The CF6-50, in addition to the sector burning, requires new fuel nozzles
to improve the mixing and local stoichiometry in the primary. The new
nozzles insure that at idle all the fuel is being injected through only
the primary orifice in each of the pressure-atomizing duplex nozzles,
thus providing greater atomization. From existing data this solution
gives the CF6-50 a 15% margin in CO emissions, but because of the antici-
pated engine-to-engine variability, there may still be compliance problems.
Further reductions could be achieved by increasing the idle power, but
as this is an engine already in use, such a procedure would run afoul
commitments to the airframe manufacturers and would particularly be
difficult to implement in a retrofit program (proposed in-use compliance
requirement by 1985).
A major concern with sector burning is the effect of the asymmetric
thermal loading at idle when the fuel is sectored in the annulus. The
most favorable fuel distribution with regard to emissions is to have a
single large sector off and the remaining sector on (here 180 on and
180 off). This, however, is the most adverse for the turbine stator
and guide vanes. Furthermore, the frame distortion may cause increased
wear on the rotor blades with subsequent efficiency losses at all power
nodes. In addition, the asymmetric heat imput reduces the mechanical
efficiency of the turbine which leads to a fuel economy penalty during
the sectoring mode.
The fuel control and delivery system also has additional complexity
(see Figure 12), but this is outside the hot section and is thus easier
-------
50
FUEL CONTROL FOR SECTOR BURNING
Staging Valve
L,
".
x 1
IA'
n
ii
ITT
Fuel
o ooi
Manifolds
ir
j{ Logic
h Signal
Figure 12
-------
' 51
•
to handle. The fuel control must be able to sector when required, must
distinguish between in-flight idle and ground idle, and must be fail-
safe (any failure will cause the system to revert to full annular
operation at the proper flow rate). The failsafe mechanism is crucial
to the safety of the system as inadvertent sectoring in flight might
lead to engine damage or inability to accelerate, depending on the power
level at the time.
Sector burning development ceased in 1979, however, when GE was
informed by its customers that, in their opinion, sector burning's
potential hazards and idle fuel penalty rendered it unacceptable. GE
has since reverted to a selective azimuthal burning arrangement. This
involves selective nozzle firing at idle, but instead of a single, large
sector being turned off (eg, 180 sequential degrees), this concept might
turn off, say, 120 degrees distributed over five individual sectors [for
instance, if there are 30 fuel nozzles, at idle, a pattern of 4 on and 2
off would be repeated 5 times]. This arrangement permits acceleration
in flight and avoids the hazards of asymmetrical thermal loading on the
guide vanes and stators which could damage the engine. Thus, there is
no potential in-flight safety problem and the system is used for both "
flight and ground idle. However, the emissions performance is degraded
somewhat.
A new dome is to be incorporated, partly to help emissions, but
largely to improve pre-existing deficiencies in the mechanical performance
of the original combustor. This dome is a version of the new design
that was developed for the new CF6-80 engine and hence represents little
-------
52
additional development effort or expense.
CF6-80
The CF6-80 is a new engine family based upon the best technology of
the CF6-50, but incorporating many new features. It constitutes a new
fanily because in its conception total design flexibility was permitted.
As a consequence, major changes in the hardware are:
(1) Aerodynamically superior fan blades (same diameter),
(2) 15 cm reduction in length of diffuser,
(3) 8 cm reduction in combustor length and replacement of the
conventional brazed ring type with a new machined rolled-ring
type,
(4) Elimination of the turbine midframe (18 cm), i •
•* •
(5) New low pressure turbine.
The overall length is shortened by 4 cm, the engine lightened by 130-
230 kg, and cruise specific fuel consumption is improved by 6% over the
CF6-50.
The new combustor incorporates a revised airflow pattern to improve
burner life and performance. Less cooling air is admitted at the dome
-------
53
which permits (1) a richer front end (and hence better relight) and (2)
more cooling air for the aft liner (and hence a cooler lir.a) . A longer
version of this improved combustor will be used in nev CF&-59s and may
be also available as a retrofit option. In response to airline desires,
GE will make every effort to avoid the use of sector burning as a control
technique, but the shorter length will work to the disadvantage of CO
control; on the other hand, it will work to the advantage of NOx control
and may help to mitigate the adverse effect of the very high pressure
ratio.
The anticipated control scheme is selective azimuthal burning which
was discussed in the CF6-50 section. The necessity for in-flight operation
requires a minimum of 5 sectors in order to preserve sufficient symmetry. .
With 30 fuel nozzles, this means 4 on - 2 off, 5 times around- With
this operation, altitude relight is marginal; however, a 2 on - 1 off
arrangement, 10 times around (ie, 10 sectors), resolves this difficulty,
but at a cost of further reduced emissions effectiveness. Table IV-3
summarizes the known abilities of the different concepts as applied to
the CF6-80. "" '
Table IV-3
performance of Control Techniques
Control HC EPAP
Sector Burning
15 on - 15 off, once around 1
Sective azimuthal burning
4 on -2 off, 5 times around* 6
2 on - 1 off, 10 times around- 12
* possible for in-flight use
-------
54
Little information about NOx control.is known at this time. However,
it is apparent that the small available volume will make the incorporation
of staged systems quite difficult. Yet, because of the small size,
=
cooling air requirements are reduced and more air is therefore available
to the two stages of a double annular combustor with, perhaps, sufficient
air left for temperature profile tailoring: this may make the staged
combustor more viable in this application, if it can be fitted in.
CFM56
This engine, designed jointly with SNECMA of France, has a core
which is derived from the military F101 engine. The cycle of the hot
core is roughly that of the military engine, but the initial combustor
intended for the military was not suitable for commercial use and con-
sequently a program for the development of a proper combustor was
established. Much of the emissions development was done on this later
model combustor (the PV combustor) which is shown in Figure 13.
In addition to a design effort to configure the double annular
•-•v •
concept to this engine (Figure 14), GE has pursued other avenues of NOx
control during the development of this engine through IR&D funding.
This approach attempted to make use of the short configuration of the
combustor which tended to lox^er NOx levels anyway (short residence
time). Because of the short design, special effort had been made to
achieve proper combustion in a very short distance by excellent fuel
preparation (atomization and evaporation) and mixing with the air. This
feature, aided if necessary by sector burning at idle, could be suf-
ficient to permit the application of a quick quench approach to NOx
-------
i'V C
(CFMr>6)
Figure 13
-------
UOUIJ1..K ANNULAR CUNK l.CUKATI.UN RM CKM!)6
(SHOWN HERE, ENERGY EFFICIENT ENGINE PROPOSAL)
C^
Figure 14
-------
57
control (see Section III and Appendix A) and yet have acceptable low
power emissions. The advantage of such an approach is first mechanical
simplicity versus, e.g., staging, and second, .inherent flame stability
compared with lean flame NOx control. . As explained in Appendix A,
however, the drawback to quick quenching is that while it quenches the
^2 -*• NO reaction (a benefit), it also at idle quenches the CO -»•
CO. reaction and possibly also the oxidation of HC (a detriment). Very
quick combustion as occurs in an advanced combustor may bypass that
difficulty. Rig testing in an F101 testbed with the original combustor
demonstrated a 30% reduction in the NOx El. Later testing in.the new
combustor, however, was not as successful, and further exploration was
shelved. Development of a double annular combustor for this engine has
not proceeded beyond the design study because of the need to resolve the
developmental problems of the concept on the parent engine (CF6-50)
HC, CO control by sector burning and selective azimuthal burning
(SAB) was investigated early in the F101 combustor rig. The new PV
conbustor with SAB had better emissions than the original combustor with
SAB; however, its operational performance was degraded considerably and
the CO was still too high. Specifically, the pattern factor and altitude
relight were deficient and the CO was twice the standard. In addition,
the liner cooling requirement was not met. A subsequent program to
remedy these deficiencies through variations in the venturi configuration
of the airblast nozzles, the fuel spray angle, the primary stoichiometry,
and the manner of dilution air entry (the degree of penetration) was
undertaken. There developed' a tradeoff situation between relight capa-
bility on one hand and CO and smoke on the other. CO and relight remain.
-------
58
as problems and exploration to resolve the relight deficiency is continu-
ing, but presently any improvement in relight is made at the expense of
CO which is not too sensitive to the burning arrangement in this case
because the origin of the CO problem is not in the primary zone (where
selective burning helps), but rather in the secondary which is too short
to permit oxidation of the CO. CO remains an unresolved problem if the
final standard is equal to the proposed value.
CF34
The CF34, a civil version of the military TF34 may be regulated if
used on an airframe finding commercial application. This engine has
received the least work especially now in light of its likely exclusion
through the general aviation exclusion. In anticipation of commercial
use, selective azimuthal burning has been investigated on a prototype
engine and a modified combustor simulating sector burning has been rig
tested (sector burning alone left the CO too high). Because of its high
bypass (6) leading to a low takeoff SFC, its moderate pressure ratioi(20),
and its short combustor leading to short residence times, the baseline,1,
engine already meets the proposed 1984 NOx requirement.
CF700, CJ610 i
As the CF700 and CJ610 would not be controlled under the proposed
requirement and as the information in the December, 1976 report is still
essentially correct and current, these engines will not be considered
here. .
-------
59
Table IV-4 presents a summary of emissions performance of the GE
engines. Rig data are identified. A projection of the performance of
the double annular combustor in the CF6-6 and CF6-32 is made, but none
is made for the CFM56 because of the scaling uncertainties. Expected
availability of the technology is also given based upon the manufac-
turer's current position and existing or anticipated problems.
-------
Table VI-4
General Electric Performance
Engine
CF6-5.0
CF6-6
CF6-32
CF34
CFM56
CF6-80
Concept
Proposed Std.
Production
Sector Burn w/
Nozzle Mod.
Dbl. Annular
Selective
burning (SAB)
Proposed Std.
Product ion
Sector Burn
Dbl. Annular
SAI5
Proposed Std.
Production
Sector Burn
Dbl. Annular
SAB
Proposed Std.
Development
SAB
Proposed Std.
Mod. PFRT
Sector Burn.
SB + adv. idle
SAB
Proposed Std,
Sector Burn
SAB
Technology
Category
2
3
1
2
3
1
2
3
1
2
1
2
2
2
1
2
1
HC
6.7
63.0
1.0
2.4
12.0
6.7
43.3
1.8
2.8
11.0
6.7
48.1
2.0
3.2
12.0
14.4
53.1
12.7
6.7
12.0
1.5
0.9
4.0
6.7
2.0
6,0
EPAP
CO
36.1
119.5
37.1
49.8
36.1
96.5
28.3
61.5
36.1
102.1
29.8
72.6
85.2
205.0
80.0
36.1
79.5
51.7
42.0
36.1
NOx
38.1
60.8
60.8
44.7
33.0
65.7
65.7
35.2
33.0
64.1
64.1
35.6
33.0
24.9
27.0
33.0
42.8
42.8
43.5
45.2
Sk
19
13
20
16
16
21
30
20
20
22.9
Development
Status*
IS
SE
R
ID
IS
SE
ID
IS
SE
ID
ID
ID
IS
ID
Projected
Implcmen- Origin
tation of
Date Data*
ET
ET
1986-7 Rig
1983 ET
BT
ET
1986-7 Proj.
1983 ET
Proj.
Proj.
1986-7 Proj.
ET
Cert. Date
ET
ET
ET
ET
Rig
Cert. Date Rig
o
* TC 53 -
TV ice
SE « Service Evaluation
ID « In Development
-------
61
«
2. Pratt and Whitney Aircraft
>
Pratt and Whitney Aircraft .(P&WA) is a ^division of the United
Technologies Corporation (UTC). P&WA is the major producer of jet
engines for commercial aviation, its most popular being the ubiquitous
JT8D (B727, B737, DC-9) . It also manufactures the JT3D and the JT9D as
well as several models of military engines. Another division, of UTC is
Pratt and Whitney Aircraft of Canada, a manufacturer of small jets and
turboprops for business aircraft. A summary of the company's engines is
presented in Table IV- 5.
of Research and Development Effort
JT9D
NOx control for annular combustor engines originated around the
NASA Experimental Clean Combustor Program (EGCP). This portion of the
jointly funded program with NASA utilized the JT9D-7 engine as the ~ *
testbed, but, as in the GE case, this effort should be considered a ~:
technology demonstration program, applicable generally to annular
combustors such as are also found in the JT9D-70, JT10D, and other
engines of similar geometry, specifically those capable of housing the
relatively long vorbix configuration when it is properly sized for its
operational performance requirements.
The ECCP was divided into three phases: (I) preliminary screening
of several concepts, (II) refinement of the best, and (III) engine
-------
Table IV-5
Pratt & Whitney Aircraft Engines
Engine
JT9D-70
JT9D-7
JT10D
JT8D-209
JT8D-17
JT8D-9
JT3D
Engine
Class
T2
T2
T2
T2
T2
Tl .
Tl
1981 .
Thrust
228 KN
205 KN
71.2KN '
64 . 5KN
84 . 5KN
1984
BPR PR Combustor
Cert. Number
Application Date Delivered
4.9 24 A . B747 1974
5.2 21.4 A B747 1971
A None 1979 0
Cn-A DC-9-580 0
1.0 17.6 Cn-A B727, B737, DC-9
1.0 15.9 Cn-A B727, B737, DC-9
1.4 13.5 Cn-A B707, DC-8
PROSPECTUS
Prospects of Meeting:
Emissions Performance
HC CO NOx Sk
Operational Mechanical
Performance Performance
Production
Category
III
III
IV
IV
III
II
I
-Time
JT9D-70
JT9D-7
JT10D
JT8D-209
JT8D-17
JT8D-9
JT3D
good
good
— not
good
good
good
no
poor
poor
known —
poor
v« f"
poor
poor
no
X X
X X
X X
XXX
XXX
x ; x . x
XXX
X X
X X
•
X X
X X
X X
X
X
X
X
X
NJ
-------
63
•
demonstration. In phase I, three concepts were explored, (1)'modifi-
cation of a conventional combustor (carbureted lean burning combustor),
(2) a radial/axial fuel staged combustor with preiaix/prevap fuel, prepar-
ation, and (3) an axial staged combustor with conventional injection and
mixing (swirl) called the vorbix (VORtex Burning and mixing), shown in
Figure 15. The philosophy behind axial staging is explained in Section
III and Appendix A. The vorbix was continued into phase II and ex-
tensively optimized so that one version (S27E) was eventually tested in
an engine (phase III). The concept performed well in the engine demon-
stration and showed the viability of the system (see Table IV-6 below).
Additional development work is needed to resolve deficiencies in
the concept in order to bring the vorbix to "state-of-the-art" per-
formance, at which time detailed development for specific hardware
application could be undertaken. The first deficiency was that while
the gaseous emissions were acceptable, the smoke levels were consider-
ably in excess of the standard (30 vs. 19). This was totally unexpected
from the results of the rig tests in phase II. Subsequent investigation
revealed that the probable cause was the main zone fuel injectors which
differed from those used in the rig tests. Presumably, therefore, this
problem can be eliminated. The second problem was that several of the
operational and mechanical performance criteria were out of specific-
ation. In particular, the temperature profile was slightly beyond
tolerance and because of the shortage of dilution air (characteristic of
lean, staged combustors), control would be more difficult. Also,
coking was observed in the main stage fuel lines and carbon was de-
posited locally within the combustor. Durability was also called into
-------
VOlUilX U)W EMISSIONS COMBUSTOK (JT9D-7)
Pilot zone
fuel injector
Pilot zone
swirler
Main zone
swirlers
Igniter
Main zone
fuel injector
I:
Figure 15
-------
65
question because of the appearance of localized hot spots on the liner.
Finally, the engine performance criteria of ground starting and ac-
celeration were deficient. The latter was marginally acceptable-(i.e.,
met standards) in most cases, but quite inferior to that of the pro-
duction engine. This was due principally to the time required to fill
the main stage fuel manifolds above idle power. Barring a technical
solution, it would seem that the definition of a flight idle with both
stages fueled would improve the required acceleration sufficiently.
This would require a squat switch so that ground idle could be identi-
fied and single stage operation used. The ground start problem has been
identified as the primary stage fuel injectors. Yet any change to
correct this problem would likely influence the altitude relight which
at this point has not been well investigated anyway.
In addition to these emissions and performance problems, the ECCP
configured vorbix suffers from certain mechanical complexities, the
consequences of which would lead to an expensive and probably difficult—
to-maintain piece of hardware. Features like the throat and the 90 " '
nozzles and swirlers in the liner and dome make fabrication difficult''
and expensive. Also, the axial staging leads to the requirement for
either two axially located rings of nozzle holes in the outer pressure
casing (costly) or long cantilevered nozzle supports in the interior
subjected to the high temperature (coking and structural problems) as
shown in Figure 16. The throat section between the two stages is not
only difficult to fabricate, but is particularly susceptible to failure
because of the:high heat transfer and difficulty in cooling the region.
-------
66
•
This, of course, would lead to high maintenance expense.
It was found that, at least in the JT9D-7 application, both stages
can operate in all flight modes and yet provide acceptable emissions
levels. This minimizes the fuel control logic, enhances reliability (by
not having the staging cycling on and off repeatedly while on approach),
and lessens the coking tendency in the long secondary stage nozzle
supports (because fuel is always flowing in flight), or alternatively,
improves the engine acceleration (because the secondary fuel manifold,
which might otherwise be drained to prevent coking, need not be refilled
prior to fuel flowing into the secondary during acceleration). This
situation is in contrast to the GE double annular staging system wherein
CO control dictated pilot stage operation only up through approach power
(30%). Nonetheless, coking in the fuel passage was observed in the main
stage (which sees a hotter environment) after'the ECCP phase III test-
ing.
Since the conclusion of the NASA ECCP, vorbix work has concentrated
on the development of a new simplified and improved vorbix system rather
than upon the refinement of the one developed during the NASA work. A
simplified system would include, if at all possible, a reduction or
elimination of the throat and reduction in the number of fuel nozzles
(the latter may badly affect the low power emissions as too few nozzles
would lead to very lean zones between the nozzles and subsequent quench-
ing of the reactions). Improvements would include lower smoke levels
and improved operational performance obtained through better air dis-
tribution and fuel control.
-------
67
COMPARISON BETWEEN ADVANCED AND ORIGINAL VORBIX CONCEPTS
. Advanced E Program Vorbix
Original ECCP Vorbix
Figure 16
-------
68
The post-ECCP vorbix work has been supported by 1R&D funding and
3 3
the NASA Efficient Energy Engine (E ) program. The E program is just
beginning and is directed towards the demonstration of a lightweight,
lov specific fuel consumption engine in the 140KN thrust size range.
This engine is to be a technology demonstrator only and is not intended
by 1IASA to be a prototype. Hence, the combustor design is not directly
usable in the JT9D-7; however, the technology is transferable. The P&WA
3
cosbustor configuration for the E program will be a throatless vorbix
with 24 aerating nozzles in the primary (vs. 30 for the JT9D ECCP) and
48 pressure-atomized carbureted nozzles in the main zone with radial
inflow swirlers to provide fuel-air mixing and flame stabilization in
the main stage. The fewer nozzles compared with the JT9D-7 should not
necessarily be construed as a simplification as it merely reflects the
snaller size of the engine (air flow rate: 65 kg/sec vs. 95 kg/sec).
The combustor also features a single plane entry of the primary and
and main stage fuel nozzle supports which are then cantilevered fore and
aft to their respective locations (Figure 16).
The IR&D vorbix study first investigated simplifications of the
ECC? configuration which included a reduced number of primary nozzles
and variations in the throat size. The rig work was done simulating the
JT9D-7 cycle. However, despite variations in the primary stoichiometry,
the emissions performance with these simplifications was found to be
inadequate. Additional investigations of advanced nozzle concepts
including carbureted nozzles and aerating nozzles on reduced length
burners (for low NOx) were also carried out. In 1977, the carbureted
nozzle work was continued and concepts refined. A new vorbix configur-
-------
69
•
ation (Vorbix II) was designed and adapted to a can burner (JT8D). This
concept employed the tested high power stage (NOx control) with a new
primary stage offering potentially improved .low power emissions.. This
stage utilized the new carbureted nozzles and preheated air for.better
vaporization and mixing. Testing continues, but no data are available
by which to judge the potential of the new configuration.
The status of development work for compliance with the proposed
1981 HC, CO and smoke standards is slightly uncertain at this point
because the more explicit idle definition in the March 1978 proposal
differs from P&WA's earlier usage on which the bulk of their effort and
data are based. This new definition has been found to have an impact on
the EPAP values of up to 50% which would vastly reduce or even eliminate
their margin for variability in most cases. The information presented
here is based on P&WA's interpretation of EPA's original idle defini-
tion. The D-7 and D-70 have separate development programs and are
discussed separately.
(JT9D-70)
The D-70 has a bulkhead type burner (Figure 17). P&WA began with a
rich primary aerating nozzle configuration and met with success after
numerpus revisions to the liner and nozzle configurations. Pressure
atomizing nozzles were also evaluated and found to have inferior emis-
sions performance, however. Development was done with both engine
testing for emissions (rig data being considered unreliable) and rig
testing for relight, coking, and durability and, in addition, a nozzle
-------
70
support program was also conducted to examine nozzle durability, coking,
etc. Despite a persistent tradeoff between .smoke and NOx (P&WA was
attempting to keep NOx at its January 1976 recommended level), one
configuration finally yielded acceptable HC,CO, smoke, and NOx, and
acceptably small penalties in pattern factor and combustor pressure
drop (Figure 18). Initially altitude relight was deficient, but minor
alterations remedied that difficulty, improving relight to beyond that
of the production combustor.
Durability remains a concern because the low emissions air dis-
tribution is considerably different from that dictated by conventional
design. The aerating nozzles, in particular, suffer early distress.
Durability assessment of alternate construction techniques, alternate
materials, and redesigned nozzles, along with cyclic endurance testing
continue. Radial temperature profile tailoring began in late 1977 and
should be concluded in time for service evaluation in 1979. Effects on
HC, CO emissions are expected to be minor as the dilution air is added
too far downstream to impact on the reactions.
(JT9D-7)
P&WA experienced more difficulty with the D-7 combustor which
differs from the D-70 (see Figures 17 and 19) by having a short cone (20
of them) burner rather than a bulkhead burner. On the basis of the
early (Phase I) ECCP data, P&WA began with a lean primary short cone
burner with aerating nozzles, which was compatible with the existing D-7
geometry. After extensive experimentation, it"became evident that it
-------
71
JT9D-70 PRODUCTION COMBUSTOR
Figure 17
JT9D-70 LOW EMISSIONS COMBUSTOR
Figure 18
-------
72
*
was impossible to satisfy all the emissions requirements (including the
P&WA NOx goal of about 60 gms/KN). The short cone burner was deficient
because it restricted the fuel distribution in the dome, in fact, wash-
ing the walls with fuel and, secondarily, its inherently lower pressure
drop limited the turbulent mixing in the primary. Both of these effects
promoted the existence of rich pockets, resulting in high smoke. Attempts
to reduce the extent of the rich pockets by leaning the overall mixture
then resulted in lean pockets elsewhere giving rise to excessive HC, CO;
hence a tradeoff existed between HC, CO on one hand and smoke on the
other.
P&WA finally abandoned the aerated nozzle short cone burner for an
aerated nozzle bulkhead type burner (Figure 20) similar to that used in
the D-70. This represented a major change, requiring lengthened nozzle
supports, new combustor supports, and a reevaluation of the effective-
ness of the diffuser. This change resolved the emissions problem when
it was found that smoke:could be controlled independently by the amount
of air admitted through the swirler surrounding the nozzle. Initial'
configurations gave unacceptable pattern factor, temperature distribu-
tion, relight or pressure drop, but eventually most of these operational
parameters have been improved to within acceptable limits. Development
'r
is expected to be completed by the end of 1978 and endurance and per-
formance testing will continue through 1979. Service evaluation should
begin in 1979.
Table IV- 7 summarizes the emissions performance of the important
low emissions configurations.
-------
73
JT9D-7 PRODUCTION COMBUSTOR
Figure 19
JT9D-7 LOW EMISSIONS COMBUSTOR
Figure 20
-------
74
JT3D
The NOx control effort for can-annular engines such as the JT8D has
been limited compared with the annular combustor effort. The effort
began with the joint P&WA/NASA Pollution Reduction Technology Program, a
program which was designed after the ECCP of the JT9D. The program
utilized the JT8D-17 and was initially intended to have three consecu-
tive phases, parralleling those of the ECCP. However, the NASA sponsor-
ship was terminated after the first phase, apparently because NASA felt
that continued support of low NOx technology for can-annular combustors
was a benefit only to P&WA and the JT8D (an older engine to begin with)
and hence not of sufficiently general interest to warrant public funding.
The program, however, was successful as far as it went. In addition to
the NASA work, P&WA has carried on some IR&D supported work.
The NASA work had three elements, each representing a different
degree of complexity. The elements are outlined in Table IV-6. The
first involved a continuation of some earlier in-house work on airbl'ast
and carbureted nozzles with airflow redistribution to affect the primary
zone stoichiometry. In general, such an approach is not expected to
have much positive influence on NOx. The second element involved the
adaptation of the vorbix concept to a can-annular combustor such as in
the JT8D. The vorbix in this configuration had an airblast primary in
each can and two pressure-atomizing simplex nozzles in the same axial
plane injecting into two carburetor tubes which carried the fuel and
inducted air downstream until past the throat at which point they
-------
75
•
entered the can through swirler orifices forming the secondary or main
burning zone (Figure 21). The stoichiometry in the carburetor tubes was
rich beyond the flammability limit to avoid flashback into the tubes.
/
This configuration minimized the extent to which the internal (to the
casing) pressurized fuel manifolds were subjected to high temperature
(equal to the compressor discharge temperature 715°K, or higher). The
third element relied upon staging again, but with prevaporizing and
premixing fuel preparation. For this system to work safely and prop-
erly, variable geometry features possibly would be required to control
the local stoichiometry at the various power settings; such features
were not explored in this program, however. Furthermore, without such
features, the total NOx level over the LTO cycle did not improve over
that of the vorbix and the CO was worse due to the very lean stoichi—
ometry.
Element I
Element II
Element III
Table IV-6
NASA/Pratt and Whitney JT8D Program Elements
Minor modifications to the existing JT8D coinbustor
and fuel system; including fuel nozzle modifications
and replacement and airflow redistribution.
Advanced versions of the Vorbix, including carbureted
fuel induction into the main stage.
Premix/prevaporization combustor schemes which employ
the vaporized state of the fuel to control flame
stoichiometry for emissions control. Variable geo-
metry may be a necessity to achieve acceptable emissions
and stable burning, but this was not explored in the
study.
The control concepts were tested in a single can (40° sector) rig
-------
JT8D VORBIX (NASA - P&WA PRT PROGRAM)
Secondary Fuel Nozzle /Crossover Tube
Secondary Swirler
Prevaporizer Tube
Figure 21
-------
77
•
at actual engine conditions which tended to give higher absolute emis-
sions levels than an engine test (based on production corabustor data).
VThile a number of deficiencies in operational and mechanical performance
could be and were identified in the program, transient phenomena were
not even investigated as this requires engine testing for evaluation.
Hence, a major portion of the required performance characteristics is
not yet known, making evaluation of the vorbix's potential still more
speculative.
The emissions performance of the vorbix was good considering the
preliminary nature of the experiments. The best configuration, however,
met only the HC standard while NOx and smoke were 10% higher than the
proposed 1984 standards (higher if a margin for variability were con-
sidered) and CO, although improved considerably, was 25% higher. While
smoke may be improved by continued development, it is a fact that the
corabustor performance in terms of the corabustor inefficiency and the NOx
emissions index (El) matched that of the JT9D vorbix which has undergone
much more refinement (Figures 22-23). It is, therefore, difficult to
anticipate significant CO and NOx improvements with this scheme. In
fact, the CO level achieved was accomplished by operation at pilot only
during approach (as opposed to the JT9D scheme which operated both
stages in the air). As this may be considered a detriment (due to the
need tto cycle staging while in flight) , the CO. level may instead in-
crease rather than decrease with any further development that would
operate only with both stages above ground idle. Performance problems
that have been identified already are carbon deposition in the main
stage carburetor tubes (due to pyrolysis in the ultra-rich mixtures),
-------
78
NOx EMISSIONS PERFORMANCE COMPARISON
BETWEEN JT8D AND JT9D VORBIX
COMBUSTORS
100
NOx
El
/gms N0x\
kg fuel
10
JT9D Production
JT8D Production
JT9D Vorbix
JT8D Vorbix
300 400 500 600 700
Combustor Inlet Temperature (°K)
Figure 22
800
900
-------
79
COMBUSTION EFFICIENCY COMPARISON BETWEEN
JT8D AND JT9D VORBIX COMBUSTORS
0.10
O
c
O
•H
O
•H
QJ
C
H
C
O
o
O
JT9D-7 Production
0.01
JT8D-17
Production
(Low Smoke)
JT8D-17
Vorbix
JT9D-7 Vorbix
(Pilot Only)
0.001
0
10
12
Pressure Ratio
Figure 23
-------
80
overheating of the liner wall, especially in the throat, pattern factor
deficiency and altitude relight (minimally examined). The radial exit
I
temperature profile (important for turbine durability) was not inves-
tigated. Temperature profile adjustment with any vorbix can be dif-
ficult because the high air demands of the two stages and the cooling
requirement leave little left for dilution near the exit. The other
operational and mechanical performance deficiencies: appear to have no
special problems that could not be resolved with normal development.
On the positive side, though, the combustor possesses a lower than
normal pressure drop which can be converted to fuel savings or exchanged
for increased mixing and possibly reduced CO. Another favorable feature
of this vorbix combustor in the JT8D is that no major changes to the
engine are required (diffuser, casing, or transition duct). Only a
proper fuel control must be designed.
Further NOx control investigation was conducted by P&WA in 1976
with 1R&D funding after the conclusion of the NASA program. The effort
centered around achieving a simplification of the vorbix concept by
reducing or eliminating the throat and wrap-around carburetor tubes.
These proprietary configurations led to compromises in the location of
~r
the main stage injection, the degree of swirl (mixing and stabilizing),
and' the amount of premixing in the main stage (carburetion). The emis-
sions performance of these configurations was degraded, some pollutants
substantially (HC, smoke); the operational performance is not known to
the EPA. This work is being continued under a general advanced vorbix
development program whose emphasis is the improvement in low power
-------
81
•
emissions, largely through advanced nozzle concepts. Although not
considered by P&WA a part of the JT8D low emissions program (possibly
because P&WA was convinced EPA would drop the NOx requirement) , the .
experimentation is being performed on a JT8D sized burner can.
The HC and CO control program originated out of the earlier smoke
control program and element I of the NASA program, and funding via IR&D
has carried the program on. With the control concept selected in 1976
(airblast nozzles with proper airflow redistribution - a richer pri-
mary) , engine testing began in 1977 for the development of durability
and temperature pattern. While the testing has yielded a configuration
with less margin than hoped for, a number of operational performance
criteria appear to have met or exceeded that of the production version
now in use: (1) Better altitude relight, (2) better cold start, and (3)
no appearance as yet of durability problems. Two separate, but similar,
burners are normally required for the JT8D, one for the D-17 and another
for the D-9 model; however, P&WA has concluded that only minor changes
will be needed for the D-9 to achieve its proper temperature profile*
(the D-9 has uncooled first stage turbine valves and hence requires a '
different temperature profile from the D-17 which has cooled vanes).
Table IV-7 suininarizes the emissions performance of several of the
important low emissions configurations.
JT10D
This is a totally new engine designed to the anticipated needs of
the next generation of commercial aircraft. As it was not selected
-------
82
*
initially to be used on the new Boeing 757 or 767/777 families, its
ultimate utilization is in question. It is intended to be in the 110-
160 KN thrust range although its final configuration has not yet-been
established. If built, it presumably would be certified prior to 1984
and hence its HC, CO emissions levels would be dictated by the proposed
1981 NME standards and not the more severe 1984.newly certified engine
(NCE) standards. Like all new larger engines, it employs an annular
combustor.
In anticipation of a low NOx requirement, the JT10D casing and cora-
bustor housing was designed to accept a vorbix type combustor, patterned
after that which would go into the JT9D. However, pending the develop-
ment of an acceptable vorbix configuration, the 10D would use a con-
ventional, single stage combustor employing only the HC, CO controls
used on the JT9D (airblast nozzles and rich primary zone). The success
of a vorbix type combustor in this application is, of course, uncertain
inasmuch as considerable development work is still required to refine
the vorbix into a state-of-the-art concept. Further hardware develop-
ment would then be required to apply the concept to the JT10D configur-
ation. On the other hand, the P&WA work for the NASA Energy Efficient
Engine Program (providing a demonstrator engine in the 100 KN class)
'f
should be very helpful as they are continuing development of the vorbix
type burner in that program.
Despite P&WA's funding of JT10D emissions since 1973, little about
the combustor geometry, performance, or status is known to the EPA at
this time.
-------
83
*
Table IV-7 presents a summary of the emissions performance of the
?&WA engines. Rig data are identified. A projection of the performance
of the vorbix combustor, as presently configured, is made for the JT9D-
70 for vhich no testing has been done. Expected availability of the
technology is also given based upon the manufacturer's current position
and existing or anticipated problems.
-------
Table IV-7
Pratt & Whitney Performance
;ine
JD-70
.'b-7
Concept
Proposed Std.
Production
Aerating Nozzle
w/ Rich PZ.
Vorbix
Proposed Std.
Production
Aerating Nozzle
w/ Rich PZ
Vorbix
:D-209 Proposed Std.
;D-17
:D-9
IS -
R »
Baseline
Aerating Nozzle
w/ Rich PZ
Vorbix (NASA)
Proposed Std.
Production
Aerating Nozzle
w/ Rich PZ
Vorbix (NASA)
Proposed Std.
Production
. Aerating Nozzle
w/ Rich PZ
Vorbix (NASA)
In Service
Research
*
Technology
Category HC
6.7
31.5
2 3.9
3 2.0
6.7
61.0
2 9.5
3 2.1
7.5
2 2.2
3 1.4
8.9
37.3
2 5.6
3 1.6
9.9
. i i
35.1
2 6.7
3 1..6
SE » Service
ET * Engine
EPAP
CO
36.1
87.5
24.4
26.3
36.1
150.0
28.0
30.2
41.2
33.6
67.4
49.9
112.7
46.5
83.1
55.9
124.5
48.5
88.0
Evaluation
Test
NOx
33.0
54.3
48.5
35.2
33.0
61.8
47.4
26.2
33.0 .
54.9
40.7
33.0
60.1
68.4 .
41.0
33.0
52.2
59.1
36.0
Sk
19
8
10
19
8
20
30
25
15
26
24
14
27
26
23
11
ID »
FT =
Projected
Implemen-
Development tat ion
Status* Date
ID, FT 1982 .
R 1986-8
IS
ID, FT 1982
R 1986-8
ID Cert. Date
R ?
IS
ID, FT . 1982
R ?
ID . 1982
R ?
In Development
Flight Test
Origin
of
Data*
ET
Proj .
ET
ET
• ET
Proj .
Proj.
ET
ET
Rig
ET
Proj.
Proj.
oj a Projected
Rig
-------
85
3. Rolls Royce
Rolls Royce is a large British manufacturing firm, occasionally
owned by the British government. Its two major divisions, Bristol and
Derby, manufacture a variety of civil and military gas turbine engines
of their own design as well as of cooperative design. The civil engines
are, in descending order of size, the RB211 family, Olympus 593, RB432
(in development), the Spey family, M45H, RB401 (in development) and
Viper for the jets, and the Tyne and Dart for the turboprops. A summary
of the company's engines is presented in Table IV-8.
Summary of Research and Development Effort
RB211
The RB211 consists of the original -22 model, the larger -524
model, and the proposed -535 which is smaller than the -22, but has its
entire high pressure core intact, including the identical combustor.
The -524 was developed with a different combustor designed to alleviate
some of the operational problems experienced by the early -22 combustors,
such as durability and smoke. That combustor, called the stage I com-
bustor, has since been incorporated virtually unchanged into the -22.
Inasmuch as all models of the RB211 presently utilize virtually the same
combustor, the discussion will generally consider all models together.
-------
Table IV-8
Rolls Royce Engines
Eng ine
Class
RB211-524 T2
RB211-22B T2
RB211-535 T2
Olympus 593 T5
RB432 T2
Spey 511 T2
Spey 555 T2
M45H Tl
RB401 Tl
Viper Tl
Engine 1981
RB211-524
RB211-22B
RB211-535
Olympus593
RB432
Spey 511
Spey 555
M45H
poor*
poor*
poor*
good
— not
no
no
Thrust
218 KN
187 KN
163 KN
50.7KN
43 . 8KN
32.4KN
1984
poor
poor
fair
no"Std,
known —
no
no
marginal
poor
BPR
4.5
5.0
0
0.64
1.0
3.0
4.2
0
HC
X
?
X
?
X
7
i
X
X
X
X
X
PR
27.2
25.0
19.3
16.1
16.9
16
Emissions
CO
X
7
X
7
X
7
X
X
•'• • .
X
X
X
Combustor
A
A
A
A
A
Cn-A
Cn-A
A
'A
PROSPECTUS
Performance
NOx Sk
X
X
X
X
X X
X
X X
X
Cert.
Application Date
B747.L-1011 1975
L-1011 1973
B757
Concorde
None
GS-II 1963
F-28 1963
VFW-614 1974
None
HS-125
Prospects of Meeting:
Operational
Performance
X
X
(Limited Data)
Number
Delivered
540+
c.80
0
1560+
1560+
38
0
Mechanical
Performance
X
X
X
X
X
X
Production
Category
III
III
IV
I
IV
II-III
II-III
III-IV
IV
I
Time
X
X
X
X
X
X
* Assuming sector burning is not used.
-------
87
•
Rolls Royce appears to be somewhat behind the U.S. manufacturers in
the exploration of fuel staging as a means of ^Ox control due largely,
perhaps, to their lack of participation in the NASA Experimental Clean
Combustor Program. However, since then Rolls Royce has begun its own
investigation with a goal of a 50% reduction in the NOx EPAP while main-
taining acceptable idle emissions; their effort has been financed partly
by British government funds. Due to the short combustor design of the
RB211, similar to that of the General Electric CF6 family, Rolls has
elected to pursue the radial staging approach.
Two alternative designs were chosen for evaluation, with the
better of the two slated for a proof of concept demonstration (similar
to ECCP phase III) test in March 1979. The selected design (Figure 24)
is a double annular combustor with the pilot and main stage nozzles
housed within short cones and surrounded by air swirlers to enhance the
fuel-air mixing. The approach is similar to GE's inasmuch as both have
double annuli of nozzles surrounded by swirlers, forming two stages
separated by a centerbody; both require a dump type diffuser to provide
airflow acceptable to the dome. However, the differences are also
considerable.
The GE combustor has the nozzles piercing the flat dome.directly
into the burning volume common to all nozzles in the stage (see Figure
f
10) r the Rolls nozzles enter into individual cones wherein mixing,
vaporization and some combustion of the fuel occurs prior to passage
into the annular burning zone. The Rolls conbustor provides for cooling
-------
LOW NOx POUBLE ANNULAR COMBUSTOR (ROLLS ROYCE)
PILOT BURNING
ZONE
MAIN BURNING
ZONE
oo
oo
Figure 24
-------
89
of the centerbody which is not found on the original GE version of the
double annular combustor, although a later version as planned for in the
E program with NASA has centerbody cooling. The Rolls arrangement
requires, unfortunately, a new combustor casing and a new diffuser (the
original RB211 diffuser being a smooth type), although not evidently a
diffuser casing. The GE combustor was designed to fit into the existing
CF6-50 envelope although application in the CF6-6 vould require a new
diffuser and casing (the -6, like the RB211, has a smooth type diffuser).
Also, it is of significance that the Rolls double annular combustor has
72 nozzles fed through 18 bosses. It thus has four times the nozzles of
the production engine (18), but manages to minimize the impact on the
casing by utilizing the existing boss arrangement. In contrast, the GE
double annular combustor only doubles the number of nozzles from 30 to
60, using the same bosses also.
Rig demonstration and development of the concept has been ongoing
since mid-1977 and is expected to continue through 1979 in an effort to
resolve a number of difficulties which include ignition and temperature
profile shortcomings and, especially, inadequate emissions reductions.
The actual performance demonstrated to date in the rig is not known to
the EPA nor are other important operational points such as the need for
in-flight staging (as in the GE combustor). New ideas are still being
investigated, but the program is continuing with an engine demonstration
(equivalent to the NASA ECCP phase III testing) scheduled early in 1979.
Design and development of production quality hardware will begin in 1979
and will involve separate, but similar, combustors for the three vari-
-------
.90
ants of the RB211. Full production is possible in 1986, if no major
development difficulties arise.
In addition to fuel staging, Rolls Royce has investigated the
potential of NOx control via quick quenching used in conjunction with an
extended rich primary zone that is swirl driven. This approach was very
similar to that explored by GE in the CFM56 engine. Their program goal
was a 25% reduction in NOx. This combustor has been tested on a -524
engine and while the excessive CO and smoke demonstrated that this was
not a solution in its present developmental state, the potential does
exist. The actual EPAP figures for the -524 test are not known, but a
Rolls Royce extrapolation of the data to the -535 operating conditions
(lower pressure ratio, in particular) predicts that the -535 would meet
the 1984 NOx requirement. This is called the stage III combustor and is
discussed again in the HC, CO section.
Acceptable HC, CO emissions in the RB211 have followed a long path
since the original 1967 design. The original combustor, although
possessing a slightly lean primary zone and airblast nozzles (simplex,
however) suffered a high degree of non-uniformity (inadequate mixing)
which resulted in excessive smoke emanating from rich pockets. Com-
pounding the problem was the fact that the combustor had only 18 nozzles
(each 20 ) despite the engine being equal in size to the CF6 (30 nozzles)
and the JT9D (20). Correction of this problem (smoke being considered a
nuisance even before the 1973 regulations) led to a redistribution of,
as well as an overall leaner, stoichiometry in the primary, both conditions
-------
91
of which would lead to less smoke production. However, as this led to
very lean conditions at idle, the combustion efficiency there suffered,
giving the -22 the worst idle emissions among the new high pressure
^
ratio engines. This combustor was referred to as the stage I combustor
(Figure 25) and as it entered service in 1975, it permitted the RB211 to
comply with the 1976 large engine smoke standard.
A parallel development, the stage II combustor (Figure 26), was
initiated at the same time (1973) in an effort to improve the HC, CO
emissions and a number of operational and mechanical deficiencies of the
stage I burner. This combustor operates with a richer primary (by
airflow redistribution and new airblast simplex injectors) , and yet
provides sufficient uniformity to keep the smoke within limits. The
stage II is now entering production in the -22B and -524.
The operational and mechanical performance of the stage II burner
is indeed superior to that of the stage I burner; hoxv'ever, it does not
provide sufficient emissions control. The control improves in the higher
pressure applications. The performance of the -524 (PR = 27), for instance,
is roughly equal to that the best non-sector burning CF6-50 technology.
On the other end of the spectrum, the smaller -535 (PR = 19) with the
stage II burner is little better than the baseline JT9D-7. Hence, Rolls
Royce has explored other avenues to supplement or replace the stage II
*
combustor. Operational control by compressor bleed and sector burning,
in particular, have been investigated to supplement the stage II combustor.
Sector burning by firing 12 of the 18 nozzles (a 240 sector) at idle
permits the -524 version to approach HC, CO standards. However, the
data available to the EPA show that the -22B, operating at a lower idle
-------
STAGE I COMBUSTOR
Figure 25
-------
STACK 11. COMM.U;TUK
U)
Figure 26
-------
94
pressure and with a loxver rated output, is still well above the stan-
dards despite the sector burning. The -535 fares even worse. Additional
improvement is not expected by firing fewer than 12 nozzles during
dector burning (GE fires only a 180 sector), as the stoichiometry is
optimized at 12 firing.
Rolls Royce is committed to the stage II combustor in all applications
at the present, largely because of their contractual guarantees for long
combustor life, but also because of the economy involved in having a
single combustor for all engines. This is true even for the -535 model,
which being new, might normally be expected to have a shorter initial
combustor life. .
As insurance against the failure of the emissions performance of
the stage II burner and out of expectation of failure for the -535,
which operates at yet a lower idle pressure and with less rated output
than the -22B, Rolls began preliminary work in 1977 on the stage III
burner, which employs an extended rich primary zone, stabilized by i ,
substantial swirl from around the nozzles, and followed by a quench to
stop the NOx reactions (Figures 27 and 28). This burner is presently in
the research stage and with a commitment to proceed, it would be available
in 1985 or 1986 (service evaluation included therein). The only emissions
performance data for the stage III combustor which is available to EPA
is that froni an early version of the combustor which did not have the
strong swirl aerodynamics in the primary driven by the swirl cups around
the nozzles (see Figure 28). That data shows the emissions performance
to be still insufficient.
-------
95
Rolls Royce is predicting that a fully developed combustor with
acceptable operational and mechanical performance could have an HC EPAP
of less than 20 in an RB211-535 application. This is a considerable
f • •
improvement beyond the stage II (EPAP = 35), but it is still little
better than the baseline JT9D-70 (It must be remembered, however, that
the JT9D-70 combustor was designed initially with emissions in mind and
is, therefore, much cleaner than other presently produced engines).
Significant CO improvements are unlikely because the quick quench design
would promote the freezing of the CO ->• CO oxidation outside the primary.
This represents a prime example of CO-NOx tradeoff inasmuch as this
quick quench feature reduces the NOx from an EPAP of 51 for the stage II
combustor to possible 30 for the stage III (marginally below the 1984
standard). Nonetheless, the addition of the swirl should have some
beneficial effect on the CO level, in particular, although a prediction
cannot be made. Again, it is possible that the addition of sector
burning to the stage III combustor might provide a sufficiently favorable
environment at idle to promote faster CO oxidation and offer additional
HC control. However, this cannot be relied upon without demonstration
because with the primary already redesigned to provide a hotter, richer
flane at idle, further richening may be excessive and in fact increase
the CO emissions (see Figure 33).
*t
£
Table IV-10 presents a summary of the emissions performance of the
^
RB211 family.
Olympus 593
This is the only T5 class engine in use. It is an outgrowth of an
older family of Olympus engines, but was considered best suited to the
-------
96
COMPARISON BETWEEN LOW EMISSIONS COMBUSTORS
Stage II
Stage III
Figure 27
-------
STACK I'll: COHIHJSTOK
Figure 28
-------
98
cask because it was sized right and possessed the proper thermody-iaraic
cycle for supersonic flight (moderate pressure ratio and no bypass).
This engine is a collaborative project, with SNECMA responsible for the
development of the afterburner, a feature not found on engines intended
for subsonic use. Because of the vintage of parent engine, the 593
began with a can-annular type of combustor (like the P&WA JT8D and RR
Spey) with pressure atomizing nozzles. Smoke problems early on precipi-
tated a conversion to an airspray type of nozzle (an airblast nozzle
similar to some of GE's with a low pressure orifice surrounded by a
swirl ring). This improved mixing and leaned out the primary zone.
However, this too proved inadequate, largely because the requirements of
coast-down from supersonic flight forced the use of a very rich primary
zone. This led to a total redesign of the combustor resulting in the
enployment of a modern annular combustor (like, e.g., that in the JT9D
or RB211) and vaporizer injectors.
Vaporizer injectors are not used by any U.S. manufacturers, but are
found in several Rolls Royce combustors of various vintages (see Ap-
pendix A, /-'I). It is basically a premix/prevaporizing concept wherein
low pressure fuel is injected into a tube which also contains a portion
of the compressor air entering the burner dome. The heat of the com-
v
bustion within the primary zone into which the vaporizer tube is in-
serted vaporizes the fuel stream before it flows out of the tube into
the primary. The usual configuration has a reverse flow at the exit,
-------
• 99
making the vaporizer resemble either a "l" or a walking cane ("J") as
the case may be. Both the cooling needs of the vaporizer (done by the
fuel) and the prevention of flashbacks require very rich stoichiorr.etry
r
in the tube.
Because of the relative leniency of the T5 standards compared with
the T2 standards, Rolls Royce is able to comply with the 1980 require-
ment (HC, CO only) with only the application of their "blown ring".
advanced cooling technology. This technology, which is also employed in
the M45H (discussed below) controls the amount and direction of cooling
air into the dome so that premature quenching of the reactions near the
wall is avoided. While this advanced cooling scheme does result in
marked HC, CO reductions (to the levels required by the T5 standards),
nevertheless, the combustor still has a greater combustion inefficiency
at idle than would be expected of a new high pressure fan engine (e.g.,
the JT9D) run at those operating conditions and using the best HC, CO
control (Figure 29), despite the fact that the high pressure engines
need and are designed for high liner cooling flows capable of signif-,
icantly quenching the reaction. -.-.
Additional combustion inefficiency (HC, CO) is found in the after-
burner employed at takeoff. This is to be expected in light of the low
pressure (though high temperature) and short residence time. Methods to
^
improve the combustion efficiency have apparently been identified by
SNECMA, the responsible partner, which would raise the efficiency from
95% to 99% (downstream from the exhaust at the completion of reaction).
-------
.100
Such methods would presumably include better and more rapid mixing of
the fuel-air mixture, probably at the expense of a slight increase in
the pressure drop across the afterburner (and therefore poorer fuel
economy).
The standards for T5 newly manufactured engines do not include any
significant requirement for NOx control, the standard in fact permitting
a minor increase in NOx in exchange for CO control. Consequently, no
technology for NOx control has been investigated. .
RB432
The RB432 is a new engine now under development which is sized to
compete directly with the existing JT8D. The engine was begun as a
successor to the Spey although its size has now grown somewhat beyond
that. It is essentially a straightforward scale-up of the 25KN RB401
engine which is also under development for the business jet market.
Very little is known about the engine at this time in view of its early
•-.•** !
stage of development. The smaller RB401 has been designed by Rolls to
satisfy the presently promulated Tl class emissions standards and con-
sequently, it may be expected that the larger RB432 would perform as
i~
well. The combustor is annular with vaporizer nozzles.
Spey
The Spey family consists of a large number of members which orig-
inated in 1963 (date of first certification). Being of older vintage,
-------
101
COMPARISON OF EMISSIONS PERFORMANCE,
JT9D AND OLYMPUS 593
0.1
0.01
.001
Production
Low Emissions
14
-------
.102
the Spey uses a can-annular burner (Figure 30), with duplex high pres-
sure fuel nozzles. Beyond that, it suffers three additional disad-
vantages, the first being a low bypass ratio (0.6-1.0) giving it a high
sea level SFC compared with modern engines; the second being a larger
nunber (10) of highly loaded short cans, and the third being a burner
fabrication technique which uses "wiggle strip" cooling (Figure 31).
This approach to supplying cooling air to the burner is simple, but
excessive, yielding a can with exceptional durability (16,000 hrs.).
The excessive cooling air, coupled with the small size of the cans
(implying short residence time and larger surface to volume ratio which
enhances the importance of quenching of the reactions by the cooling
air) together create an environment conducive to incomplete combustion,
especially at idle. Hence, the HC, CO emissions are very high (idle
cotabustion efficiency is only 90%) and extraordinary effort must be made
to reduce them. Smoke also is a problem due to poor mixing in the
primary zone, a result in part of the small pressure drop across the
combustor head which is, in turn, partially a result of the large cool-
ing air flow (low resistance).
Low emissions work first began in 1969 when Rolls was first con-
tracted by the USAF to produce a low smoke combustor for the TF41 (a
military Spey). Both a piloted airblast nozzle and a revised combustor
head configuration (the conical head) were investigated (Figure 32).
These approaches reduced smoke (to SAE No. 14) through improved local
stoichioinetry, but without any concurrent HC, CO reduction. Addition-
ally, the conical head scheme, which was the better, suffered persistent
-------
Spey Production Burner
o
CO
Figure 30
-------
104
COMBUSTOR LINER COOLING METHODS
Wiggle Strip Cooling
Splash Plate Cooling
Figure 31
-------
105
•
and severe carbon deposition and durability problems. A low HC, CO
emissions (as well as smoke) investigation was begun in 1972. This
program, was quite extensive, involving half a dozen different approaches
and nearly 400 rig and engine tests through 1976, at which time Rolls
concluded that while substantial reductions to the HC and CO could be
made, compliance with the existing HC and CO standards (and for that
matter, with the proposed regulations) was impossible. The basic con-
cepts that were investigated are listed in Table IV-9.
TABLE IV-9
Low Emissions Investigation - Spey
1. Improved atomization
2. Airblast nozzles
3. Sector burning
A. Advanced cooling
5. Vaporizer nozzles
6. Reflex airspary burner (RAB)
There was a considerable variation in the degree of success and'the
difficulties encountered among these six concepts. Improved atomization
was possible by rescheduling the duplex nozzle fuel flow to run on the
primary only at idle. The primary, being sized for ignition, gave good
atomization at low power and, if used alone, afforded some reductions.
Apparently, though, this was not the principal source of the emissions
as the amount of control was quite modest. Piloted airblast nozzles
reduced smoke by leaning the primary zone some and improving mixing at
-------
CONICAL HEAD BURNER (LOW SMOKE)
Figure 32
-------
107
•
high power, but had even less effect on the emissions at low power, both
because poor atomization was not the principal source and because the
airblast feature was least effective at the low pressure drops expe-
f
rienced at idle.
Sector burning provided considerable reduction, especially of HC,
by richening the primary (the Spey runs slightly lean at full power, and
leaner yet at idle in normal operation) and improving atomization. This
large improvement would not necessarily be witnessed with combustor
configurations other than the production (its success is very dependent
upon the sensitivity of the combustion efficiency to the fuel/air ratio
in the primary as is seen in Figure 33), but inasmuch as sector burning
alone cannot achieve.compliance, then compounding of this with other
schemes for emissions reduction would be mandatory. However, it is
understood that this sector burning was achieved by the elimination of
only 3 of the 10 cans (leaving a 252 sector on). GE fires only a 180
sector to achieve its reductions and while this adds to the potential
operational and mechanical problems, it is likely to be more effective.
An engine test of sector burning with 3 cans out resulted in the burnout
of some nozzle guide vanes so the mechanical problems are indeed a
reality. Also, sector burning in a can-annular system results in
peculiar crossover flow patterns resulting from a difference in pressure
drop between the lit and unlit cans. This possibly would affect the
liner durability. On the other hand, the primary zone stoichiometry is
not the only contribution to the emissions problem as was discovered
during the investigation of advanced cooling.
-------
Impact of Combustor Design on
Effectivenss of Sector Burning
o
o>
•H
O
•H
W
c
o
•H
4J
W
3
,0
O
Improvement in
Low Emissions
Burner by Sector
Burning
Low Emissions Burner
Sector burning
operating line
Improvement in
Production Burner
by Sector Burning
Production Burner
o
CO
vFull annular
operating line
Fuel/Air Ratio (Primary)
Figure 33
-------
109
Advanced cooling schemes are based upon recognition that the
primary source of HC, CO emissions in the Spey is the reaction quenching
that occurs adjacent to the lean primary at'idle. Sector burning, in
contrast, richens the primary to reduce the quenching effect, but
advanced cooling attacks the problem directly by designing the combustor
to survive without the excessive film cooling that is provided (without
choice) by the "wiggle strip" construction. The basis of the advanced
cooling is the use of a new composite sheet material (unknown to EPA)
which is more durable in the thermal environment than what is now used
in production. This requires less cooling air to provide the same
excellent life and, therefore, permits the redesign of the cooling air
patterns accordingly. Fabrication problems have been solved and the
operational performance has not degraded. The major mechanical problem
seems to be the interface between with the composite and the conven-
tional materials in the latter half of the liner. Experiments have
shown that about half of the quenching was at the head and the other
half in the front portion of the liner, so all of this must be made of
^ «.
the new material. Idle combustion efficiency increased from 90% to
96.9% using production nozzles which is, unfortunately, still insuf-
ficient. Continued development would necessarily consider the effect of
compounding schemes, in particular, the mating of the advanced cooling
.'
burner with such as cross stream aerated injectors, piloted airblast
noz2les, sector burning, blown rings (see Olympus or MASH), and enhanced
mixing domes ("potted head").
Vaporizer nozzles were investigated early, being a standard feature
-------
110
•
on other Rolls Royce engines (e.g., M45H). Although emissions were re-
duced some (smoke especially) and most operational criteria were satis-
j
factory, relight was much degraded, evidently due to the airfloxy dis-
turbance resulting from the blockage caused by the vaporizers themselves
(5 per can). This in turn led to the development of the reflex airspray
burner (RAB). .
The RAB as developed is not merely a fuel injector concept, al-
though that is an important facet of it. The RAB includes a radical
change in the primary zone aerodynamics, specifically a dual reversal
flow which acts as an aerodynamic staging system (see Figure 34) . The
first reversal is sized for idle and the second for takeoff. At take-
off, the first zone burns excessively rich (without fuel staging, it
still accepts all the fuel) and acts as pilot zone for the second re-
versal. A simplex injector is employed and located within a carburetor
tube which acts as a vaporizer. Evaluation of numerous variants has led
to a best version attaining a 98.4% combustion efficiency at idle with
an improved pattern factor. Additional work resolved earlier head
cooling problems, but the deteriorated ignition and relight has remained
intractable. In theory, the concept also has the potential for NOx :
reduction by controlling the residence time in the second reversal zone.
Although two attempts at control gave 30% and 50% reductions ^ they could
be achieved without deterioration of the still high idle emissions.
The RAB and the advanced cooling approaches gave the best emissions
(Table IV-9), but because of the constraint upon resources, Rolls was
-------
KF.FLKX AIKSPRAY BURNER
Figure 34
-------
112
forced to select only one for continued development. In 1977, Rolls
chose to abandon the RAB for although it gave the best emissions, the
relight and ignition problems appeared insurmountable while, on .the
other hand, advanced cooling seemed to have the potential to at least
match the RAB emissions performance with better operational performance.
/
Table IV-10 summarizes the.Spey performance to date.
M45H
The M45H was originally a collaborative project with SNECMA and was
developed for application in the short haul airlines, the VFW-614. The
airliner, however, had few buyers and production ceased in 1978 after
only 16 were built, thus leaving the M45H without purpose. Since 1976
Rolls Royce has taken over full responsibility for the engine which is
now its own. The engine is quiet and fuel efficient which is likely to
be favored in future aircraft for which its size is appropriate. A
large refanned version using the same core, called the M45SD (RB410)* has
been demonstrated and should expand its potential for future applica-
tion. In fact, the -SD at 64KW is sized slightly larger"than Spey and
so is a potential replacement. The -SD remains a collaborative venture
with SNECMA and Dowty Rotol. .
The engine utilizes an annular combustor with vaporizer nozzles
(walking cane configuration). Its production version is quite high in
HC and CO due to reaction quenching at the wall by entrainment of the
-------
113
fuel into the film cooling air. However, employment of the "blown ring
advanced cooling technique, used also in the Olympus 593, brings the HC
and CO emissions to within the proposed standards. The NOx emissions
are already below the proposed.level because of the moderate pressure
ratio and low sea level SFC which occurs as a result of the moderately
high bypass ratio. It remains to be seen, however, if there is suf-
ficient margin (for CO and NOx especially) for variability, but if not,
slight modifications to the liner to provide a better airflow distri-
bution may prove sufficient. Other low emissions concepts which had
been investigated earlier on the M45H were redesign of the vaporizers,
alternative schemes to fuel the vaporizers (specifics unknown), and
leaner primary zone stoichiometry, but none of these were successful in
lowering the emissions to the requisite level.
Further low emissions development work has ceased pending the
identification of a new application of this engine.
Table IV-10 presents a summary of the emissions performances of 'the
various Rolls Royce engines. While most data was obtained by engine'*"
testing, some was from either rig testing or was derived from extra-
polation from other conditions (e.g., the RB211-535 data). In certain
cases, the data, while showing excellent emissions control, may also
reflect unflightworthy hardware or operationally defective systems.
Therefore, care should be exercised in the evaluation of the potential
success an engine may have in achieving a given emissions level.
-------
Table IV-10
Rolls Royce Performance
Concept
Technology
Category
'•211-22B Proposed Std.
Production
Phase II 2
Phase II w/ sector burn 2
Double Annular 3
•>211-524 Proposed Std.
Production
Phase II 2
Phase II w/ sector burn 2
Double Annular 3
v211-535 Proposed Std.
Phase II 2
Phase II (7% idle) 2
Phase II w/ sector burn 2
Rich PZ w/ Quick Quench 2
Iytr.pus593 Proposed Std.
Production
Blown ring
pey MK511 Proposed Std.
Production
RAB
Advance cooling
pey MK555 Proposed Std.
Production
RAB
Advance cooling
2
2
2
2
45H-01' Proposed Std.
Production
2 Bloxm. rings (7% idle) >2
EPAP
HC
6.7
135
8.3
4.2
6.7
110
6.0
. 3.1
6.7
19.1
32.4
8.9
2.5
30.7
129
<30.7
12.2
278
23.0
75.5
13.6
441
36.1
75.6
16.2
162
30.1
CO
36.1
172
49.6
28.8
36.1
145
39.0
22.4
36.1
90.0
96.6
54.7
67.5
237.0
530
<237
70.9
436
162
229
79.8
420
186
232
97.1
526
169.9
NOx
33.0
51.9
61.7
64.0
34.6
61.4
68.0
7.0.2
33.0
49.0
51.3
30.3
70.8
70.1
<70.8
33.0
68.1
68.2
58.0
33.0
49.5
55.2
54.2
33.0
31.2
37.0
Sk
20
15
18
19
18
18
21
25
26
28
66
29
IS = In Service
R a Research
SE = Service Evaluation
ET. = Engine Test
31
46
12
ID
FT
Development
Status*
IS
ID, FT
ID
R
IS
ID, FT
ID
R
ID
ID
R
R
IS
ID, FT
IS
ID
ID
IS
ID
ID
IS
ID
In Development
Projected
Implemen-
tation
Date
1982
1986+
1982
1986+
1983
1983
1986
1986+
1980
Cancelled
1982
Cancelled
1982
In abeyance
pending new
orders
ET
Rig
Rig
ET
Rig
Rig
Rig» C78)
Rig C'79)
Proj
Rig
ET
Proj
ET
ET
Rig
ET
ET
Rig
-------
.115
4. Avco Lycoming
Avco Lycoming (generally referred ,to as simply "Lycoming") is
a U.S. manufacturer of both piston engines (Williamsport, PA) and gas
turbine engines (Stratford, CT) for aircraft. It is a subsidiary of the
larger Avco Corporation. The gas turbine division produces principally
shaft power turbines for fixed wing aircraft and helicopters, largely
for military applications. Of interest here, however, is the new ALF
502 turbofan engine of the Tl class. Because of its size (29-34 KN) , it
would be subject to the proposed standards if used commercially. A
description of the ALF502 is given in Table IV- 11.
Summary of Research and Development Effort
It was recognized early that the prototype 502 would not comply
with the standards promulgated in 1973.
The basic engine core is that of the T55 turboshaft engine (2800
KW), a design which dates back to about 1960. That core was first tried
in jet applications in the F-102, a military predecessor to the ALF502
series. Although the cores are essentially the same, minor combustor
modifications are necessary in the new applications in order to provide
acceptable performance in the new environment (e.g., increased cooling
in higher pressure applications). The cotnbustor is a reverse flow
annular type and is the only such one affected by the proposed stan-
dards.
-------
Table IV-11
Lycoming Engines
Engine
ALF502D
ALF502H
ALF502L
Engine
ALF502D
ALF502K
ALF502L
Class Thrust
Tl 28.9KN
Tl 29.8KN
Tl 33.4KN
1981 1984
fair
fair
7
7
fair
marginal
BPR PR
5.8 11.1
11.4
13.3
Emissions
HC CO
X
X
X
X
Cert. Number Product io
Combustor Application Date Delivered Ciitep.ory
RFA None 1976 0
RFA HS146 — 0
RFA Challenger 1979 0
PROSPECTUS
Prospects of Meeting:
Performance Operational Mechanical
NOx Sk Performance Performance
X
X
X
X X
IV
IV
IV
Time
•„.
-------
117
The baseline engine, employing the T55 standard dual orifice
nozzles, required more than a 50% reduction in KG and CO emissions to
-eet the 1973 mandated levels. This was due largely to the fairly low
conbustor inlet temperature experienced at idle which made vaporization
difficult and the reaction speed slow. Countering this to some degree
was the excellent fuel dispersion which arose from the large number of
nozzles (28, each covering only a 13° sector).
Methods to reduce HC, CO emissions logically centered around means
to enhance the fuel vaporization. Vaporizing injectors similar to those
used by Rolls Royce on some of its engines were investigated; however,
because of the low air temperature entering the combustor, the vapor-
izers could not function adequately. This was unfortunate in that such
injectors have the potential to reduce high power NOx production suf-
ficiently (15% or more) on the 502 to achieve compliance with that
standard also.
Airblast injectors were also tried with greater success. These
injectors, supported by coinbustor airflow redistribution, provided the :
best KG, CO control although the best configuration still failed to meet
the 1973 HC and CO requirements and, in addition, suffered from degraded
operational and mechanical performance. Subsequent development to
improve upon these deficiencies increased the HC, CO emissions somewhat
(certification configuration). Variations in the primary zone stoi—
chioiaetry, cooling air flow mixing patterns, and residence times were
explored and tradeoffs between reductions in HC and CO on one hand and
radial temperature profile, pattern factor, and the lean stability limit
-------
118
*
on the other were found to exist which could not be eliminated.
Lycoming has never emphasized research on NOx control, preferring
instead to rely upon the output of the NASA Pollution Reduction Tech-
nology Program for Small Gas Turbines (PRTP-SGT), the contract for which
had been awarded to AiResearch rather than Lycoraing. This program 'will
be completed in 1979. Nonetheless, Lycoming claims to have examined NOx
control techniques such as preinixing, fuel staging, and staged air
introduction. Nothing is known about the extent of the work and pre-
sumably it has not been carried beyond preliminary rig testing. Ac-
tivity in this area was continuing in 1977, but. no further information
has been provided.
With the loss in 1976 of the contract for the U.S. Coast Guard
medium range surveillance aircraft which v/ould have used the ALF502H
model, the first use of the 502 now appears to be the Canadian Chal-
lenger (formerly Learstar 600), a new business jet which employs two
502L engines. The L version, using nearly the same combustor as the D,'
(except for minor durability changes) also fails the 1973 requirements!
In fact, at the originally specified idle of 1800 KN, it suffered a
greater combustion, inefficiency (HC and CO) than the H model, apparently
because of its even lower pressure and temperature. To achieve com-
pliance, Lycoming resorted to an advanced idle point which is an op-
•
erational control technique. Unfortunately, the approach borders on the
desperate in this case, for an idle power of 10.7% is required (3580 KN)
which has precipitated a need for thrust spoilers at idle, thus adding
weight as well as noise. In fact, while the 10.7% is ample-for an
adequate margin for variability under the 1973 standards, the proposed
-------
119
•
standards, although relaxed in principle, do not give-"credit" for a
high idle in the denominator of the control parameter and hence nearly
all of the margin is eliminated. How Lycoming plans to deal with this
is uncertain at this time.
Certification of the Challenger is planned with the high idle and
thrust spoilers and is expected next year. Under 1973 standards,
Lycoming was in an excellent position to comply in a timely fashion;
however, with the proposed standards and the eroded margin, its status
is less certain, although the two year delay should be sufficient for
adjustments to be made. The biggest advantage for Lycoming and the
Challenger is the general aviation exclusion which eliminates the ALF502
from control unless, perhaps. Federal Express decides upon the Chal-
lenger in addition to its freight fleet. The ALF502 is also scheduled
for use on the new four engined HS146 short haul airliner by British
Aerospace Corporation (formerly Hawker Siddeley). However, this air-
craft appears to be aimed principally at the European market and thus
would not be under the umbrella of the standards. Should U.S. commuter,
airlines decide to buy it, hox^ever, Lycoming would again be confronted.
with the issue whether to proceed with advanced idle in this application
for emissions control.
-•-.
Table IV-12 summarizes the emissions performance of the various 502
• _ .
configurations.
-------
Table IV-12
Lycoming Performance
Technology
Engine
ALF502D
Concept
Proposed Std.
Prototype
Category
Airflow distribution
2
HC
17.
31.
14.
0
0
8
EPAP
CO
103
183
112
NOx
33.0
38.3
28.8
Sk
32
25
Development
Status*
IS
Projected
Implemen-
tation
Date
Cert.
Orgin
of
Data*
ET
plus airblast nozzles
ALF502L
Proposed Std.
Baseline (same as
Cert, configurat
Advanced idle
502D
ion)
1
15.
28.
9.
9
6
1
95.5
136
92.2
33.0
32.3
35.4
31
FT
FT
1982
1982
DT
ET f
* IS = In Service
R = Research
Proj = Projected
SE = Service Evaluation
ET = Engine Test
Rig =
ID = In Development
FT = Flight Test
-------
121
•
References
1. Aircraft Emissions: Impact on Air Quality and Feasibility of
Control, EPA (no date). ,
2. Assessment of Aircraft Emission Control Technology, NREC, EPA
Contractor Report No. 1168-1, E. K. Bastress, et al., September
1971.
3. Control of Air Pollution from Aircraft and Aircraft Engines, 40
CFRPt87, FR 3£, N. 136, July 17, 1973, p. 19088.
4. Aircraft Technology Assessment - Interim Report on the Status of
the Gas Turbine Program, EPA, R. Hunt, et al., December, 1975.
5. Aircraft Technology Assessment - Status of the Gas Turbine Program,
EPA, R. Hunt and E. Danielson, December, 1976.
6. Control of Air Pollution from Aircraft and Aircraft Engines, NPRM,
FR 43^, N. 58, March 24, 1978, p. 12615.
7. Review of Past Studies Addressing the Potential Impact of CO, HC,
and NOx Emissions from Commercial Aircraft on Air Quality, EPA AC
78-03, P. Lorang, March 1978.
8. An Assessment of the Potential Air Quality Impact of General
Aviation Aircraft Emissions, EPA, B. C. Jordan, June, 1977.
9. Potential Impact of NOx Emissions from Commercial Aircraft on N02
Air Quality, EPA, B. C. Jordan, November, 1977.
10. Cost-Effectiveness Analysis of the Proposed Revisions in the
Exhaust Emissions Standards for New and In-Use Gas Turbine Aircraft
Engines Based on Industry Submittals, EPA AC 77-02, R. S. Kilcox ' '
and R. W. Munt, December, 1977.
11. Cost-Effectiveness Analysis of the Proposed Revisions in the
Exhaust Emission Standards for New and In-Use Gas Turbine Aircraft
Engines Based on EPA's Independent Estimates, EPA AC 78-01, R. S.
Wilcox and R. W. Munt, February, 1978.
12. Letter from D. W. Bahr (GE) to J. P. DeKany (EPA) dated April 13,
1977. Subject: Comments to reference 5.
13. "Telecom, R. Munt (EPA) to D. Bahr (GE), December 7, 1977. Subject:
Technology for the Proposed 1981 Standards.
14. Telecom, R. Munt (EPA) to D. Bahr (GE) , December 16, 1977. Subject:
Technology for the Proposed 1981 Standards.
15. Letter from D. W. Bahr (GE) to E. Danielson (EPA) dated March 21,
1977. Subject: CF6 data.
-------
122
*
References cont.
16. A Petition from Rolls Royce, Ltd. for Modification of the Emis-
sion Standards to Rolls Royce Spey Engines, Rolls-Royce DP314,
March, 1977.
17. Status of Rolls-Royce Emission Reduction Technology and Pro-
grammes Applicable to Newly Manufactured Engines of Current
Models, Rolls-Royce DP305, December, 1976 (Private Data).
18. Letter from A. Gray (Rolls-Royce) to J. P. DeKany (EPA) dated
January 28, 1977. Subject: Comments to reference 5.
19. Comments by Rolls-Royce on the EPA Aircraft Technology Assess-
ment, Status of the Gas Turbine Program, December 1976, Rolls-
Royce DP311, February 1977.
20. Additional Comments Related to the EPA Aircraft Technology Assess-
ment, Status of the Gas Turbine Program, December 1976, Rolls-
Royce DP311 Addendum.1, May 1977.
21. Letter from A. Allcock (W. K. Dept. of Industry) to J. P. DeKany
(EPA) dated February 17, 1977. Subject: Comments to reference
5.
22. Telecom, R. Munt (EPA) to R. Rudey (NASA), November 17, 1977.
Subject: JT8D NOx Control.
23. Telecom, R. Munt (EPA) to R. Rudey (NASA), November 18, 1977.
Subject: NOx Control Technology.
24. NASA Comments to draft NPRM, September 20, 1977.
25. Letter from G. N. Frazier (PWA) to C. Day (LMI), December 15,
1977. Subject: Estimated Economic Impact of Proposed EPA
Emissions Regulations for Aircraft.
26. Letter from G. N. Frazier (PWA) to J. P. DeKany (EPA), April 11,
1977. Subject: Comments to reference 5.
27. Telegram from D. F. Heakes (Dept. of Transportation, Canada) to
W. Oleksak (FAA) dated April 14, 1977. Subject: Idle Speed
for certification of the Canadian Challenger.
*
28. Letter from G. Opdyke (Lycoming) to N. Krull (FAA) dated June
30, 1977. Subject: ALF502 Data.
29. Letter from G. Opdyke (Lycoming) to J. P. DeKany (EPA), April
15, 1977. Subject: Comments to reference 5.
-------
123
•
References cont,
30. Letter from D. Bahr (GE) to R. Munt (EPA), dated August 7, 1978.
Subject: CF34 Emissions Data.
31. Letter from D. Bahr (GE) to R. Munt (EPA), dated February 8, 1979.
Subject: Fuel Flowrates at Various Idle Power Settings.
32. Letter from D. Bahr (GE) to A. J. Broderick (FAA), February 24,
1978. Subject: Emissions Data for ICAO.
33. Telecom, R. Munt (EPA) to D. Bahr (GE), May 17, 1978. Subject:
Status of GE Low Emissions Programs.
34. Telecom, R. Munt (EPA) to D. Bahr (GE), October 20, 1978. Subject:
Double Annular Combustor, Performance and Problems.
35. Letter from M. Sherwood (Rolls Royce) to R. Munt (EPA) dated June
2, 1978. Subject: Emissions Control Technology for RM211 and Spey.
36. Comments by Rolls-Royce on the EPA Aircraft Technology Assess-
ment, Status of the Gas Turbine Program, December, 1976, Rolls-
Royce DP311, ISSUE 2, June, 1978.
37. Presentation by Rolls-Royce Representatives to FAA Personnel,
December, 1978. Subject: Emissions Technology, Idle Definition
for Regulations, Compliance Procedure for Regulations.
38. Memorandum to the Record, R. Munt (EPA), April 3, 1978. Subject:
Report of Visit to NASA - Lewis Research Center.
39. Results and Status of the NASA Aircraft Engine Emission Reduction
Technology Program (NASA TM 79009) October, 1978, R. E. Jones,
et al.
40. Telecom, R. Munt (EPA) to P. Goldberg (PWA), October 20, 1978.
Subject: Installation of the Vorbix and Airblast Cornbustors on
the JT9D.
41. Telecom, R. Munt (EPA) to P. Goldberg (PWA), May 26, 1978.
Subject: Technology Status of Emissions Control.. '•
42. Pratt and Whitney Aircraft Submittal to ICAO (AEESG), February
21, 1978. Subject: Emissions Data.
43. Submission by Pratt and Whitney Aircraft to EPA Docket OMSAPC
78-1, Control of Air Pollution from Aircraft and Aircraft Engines,
December 1, 1978.
44. Submission by Rolls-Royce Ltd. to EPA Docket OMSAPC 78-1, Control
of Air Pollution from Aircraft and Aircraft Engines (Rolls-Royce
Document, DP347), November 1978.
-------
124
References cont.
45. Submission by Avco Lycoraing to EPA Docket OMSAPC 78-1, Control of
Air Pollution from Aircraft and Aircraft Engines, November 30,
1978.
46. Submission by General Electric to EPA Docket OMSAPC, Control of Air
Pollution from Aircraft and Aircraft Engines, November, 1978.
47. Testimony of the EPA Public Hearing on Revised Aircraft Engine
Emission Standards, Volume 1 and 2, November 1-2, 1978.
48. Aviation Week and Space Technology, Vol. 110, No. 19, May 7, 1979,
pp. 43.
49. Aviation Week and Space Technology, Vol. No. 22, .May 28, 1979,
pp. 46.
-------
APPENDIX A
EMISSIONS CONTROL TECHNIQUES
HC and CO Techniques <
Operational Control Techniques
1. Increase in Idle Speed - As engine power is increased, HC and
CO levels generally decrease as a result of higher temperatures and
pressures at the combustor inlet. However, the NOx level may increase
because of the increased temperature in the combustor. The increased
idle speed is limited on turbofan and turbojet engines by the capability
of the aircraft brake systems as there is an increase in thrust at idle.
This problem does not exist with turboprop (class P2) engines as the
thrust can be held nearly constant by properly varying the propeller
pitch with engine speed. However, there is an attendant increase in
fuel consumption and noise with increased engine speed.
2. Airbleed - Airbleed (of compressor air) is a means of in-
creasing the work load of an engine with, hopefully, the same result as
occurs by increasing-the idle speed (another form of work load), yet
without the concomitant penalty of higher thrust and its ensuing braking
requirement. To be effective, the engine must increase its fuel con-
i
sumption with the bleed to provide the necessary energy for compressing
the extra air and to maintain the idle power. Failure to do so will
cause the engine to lose power and the emissions to rise.
-------
A-2
*
Fuel Preparation Control Techniques
3. Spray Improvement - Design changes to pressure atomizer
nozzles can lead to changes in the character of the fuel droplet size
distribution. Decreasing the flow number (equal to the fuel flow rate
divided by the square root of the injector pressure differential)
reduces the droplet size. This in turn reduces the evaporation time and
strongly influences the amount of HC left unburned. To a lesser extent,
the change in the evaporation rate affects the local fuel-air mixture
ratio and thus the local temperature which would likely affect the CO
and NOx levels. This approach is not universally profitable, however,
as at very low combustor inlet temperatures, no degree of atomization
will improve the droplet evaporation for there is simply insufficient
heat available.
Incorporation of this approach into a hardware system involves
changing both the pressure differential across the injector and the
orifice diameter as otherwise the fuel flow rate would be increased at
each power setting because of the change in pressure. Changing the '"•
pressure differential requires only a new set of valves and possibly a
pressure boost in the fuel pump.
In addition, nozzle design changes intended to optimize the fuel
•
spray cone angle, and thus the distribution of fuel in the primary zone,
are relatively easy to incorporate into a combustor. Decreasing the
angle of a wide (richening the. mixture) angle spray cone reduces wall
wetting and increases the local equivalence ratio to produce a hotter
-------
A-3
•
flame which in turn helps to reduce KG and CO. Similarly, widening the
spray angle reduces the equivalence ratio, providing a leaner mixture
which might be necessary in the case of an excessively rich mixture
(insufficient oxygen to burn the fuel). There is no impact on the
system hardware as only the fuel nozzle is changed and the new one is no
more complex than the old. The new heat release distribution, however,
may require changes in the liner cooling air.
STD. FUEL NOZZLES
NARROW ANGLE FUEL NOZZLES
GE CJ610
4. Airblast - The pressure differential that exists between the com-
pressor and the combustor is employed to achieve high velocity air through
a venturi system in the fuel nozzle. This high velocity air is directed
^
at the fuel stream as it comes off a lip. The fuel is thus sheared
off and shattered into minute droplets, conducive to dispersion and
•
complete evaporation. The addition of the airblast air into the primary-
affects the stoichiometry and consequently it proves necessary to re-
distribute the airflow throughout the liner in order to reestablish the
-------
A-A
optimal fuel-air ratio pattern.
Success in. improving the combustion efficiency by utilizing this
technique has varied among the manufacturers depending upon the extent
to which the liner flow was optimized. Also, it has been found that
NOx increases slightly at idle as a result of the better combustion ef-
ficiency. At low power and especially in low pressure ratio engines,
the pressure differential across the injectors is reduced causing the
air velocity around the fuel injectors to be reduced. Therefore, the
airblast effect on fuel atoinization tends not to be as effective at lotJ
power where the bulk of the HC and CO is formed. The concept also tends
to be less effective in reverse flow annular combustors as the nozzle is
located so as not to be able to take advantage of the dynamic component
of the pressure.
Conventional Pressure
Atomizing Nozzle
Airblast Nozzle
P&WA JT8D
-------
A-5
«
5. Air assist - In the air-assist technique compressor air is
diverted and compressed externally, and then discharged around the fuel
injectors. This high velocity air is directed through the fuel in
jectors in a manner similar to the airblast technique and achieves the
sane goals. However, the external compression provides high velocity
air at all powers so this technique may be expected to be more effective
than airblast in controlling HC and CO at idle.
The use of air assist would have a large impact on aircraft hard-
ware systems because of the requirement to bleed compressor air and
compress it externally with an auxiliary compressor. The externally
coEpressed air is more effective than the air of the airblast concept
and so less is needed. This then has a less marked impact on the
stoichiometry so that the need to redistribute the liner airflow is
considerably lessened.
Airblast
Air-Assist
Fuel
JSI
JX1
Far
Compressor
Burner
Turbine
Air-Assisted, Airblast Fuel Nozzle
AiResearch TFE731 Concept
External
Compressor
-------
A-6
•
.'-
6. Preinix(l) - The basic idea is that HC and CO emissions often .
arise because of poor mixing within the primary so that while the
average equivalence ratio in the primary nay be acceptable (roughly
stoichiometric at idle), there are zones of excessively rich or lean
mixture, both of which lead to HC and CO production. One way to prevent
this is to premix (and prevaporize) the fuel prior to the flame zone so
that the additional mixing time will lead to a more homogeneous mixture.
In order to prevent flashback of the flame, the premixing can be com-
promised a bit by keeping the local equivalence ratio above the stable
deflagration limit in the premix zone. Upon entering the flame zone,
further mixing can occur, permitting combustion. Although this can lead
to less than perfect mixing and thus reintroduce to a degree the
original problem, the HC and CO emissions are much improved because of
the partial mixing and total fuel evaporation in the premix zone. The
excessively rich premix zone can lead to carbon deposition, however.
NOx cannot be controlled by this approach unless total and lean
premixing occurs which directly leads to the flashback problem, making
premix for NOx control considerably more complex (premix(2)).
Implementation of this scheme into an existing design can be
difficult in that the rather major combustor modifications must normally
*Number refers to one of two different degrees of control generally
recognized possible with this concept.
-------
A-7
be kept within the constraints of the existing envelope. The more space
that is taken for the premix region, the less that is available for
dilution and pattern factor adjustments. Ideally, the combustor'would
be nade longer with the premix zone merely being tacked on to the
existing geometry (with some airflow adjustments).
PREMlXIfiS TU3E
11.4%
12.3% ".6%
P&WA JT8D
Air Flow Distribution Control Techniques
7. Advanced Cooling - High pressure ratio engines which operate
at high corabustor temperatures require high levels of cooling air to
control the temperature of the liner (the amount of cooling air required
\
is even higher because the high pressure ratio causes the cooling air to
be proportionally hotter and therefore less effective). This cooling
air, upon entering the liner, tends to quench the reactions of the
burning fuel near the wall, especially that of CO -»• C0?. Any
advanced cooling technique which will provide the requisite cooling
-------
A-8
*
effect while at the same time reducing the air needed may improve CO
emissions and possibly also HC emissions. The key to its effectiveness
is the degree to which combustion, principally CO oxidation, is occuring
near the wall.
The overall impact of such a revision is considerable for it
represents essentially a totally new combus.tor. While benefits beyond
simple emissions control may be accrued (combustor of greater longevity) ,
the development cost is likely to be high.
Old
New
GE CF6 Combustor liner
8. Rich Primary (1) - the term "rich" applies to the stoichio-
» - - ' —" - •—™ '
metry at the design point (high power). Such a condition leads natural-
ly to near perfect stoichiometry in the primary at idle. This results
-------
A-9
4
in a very hot flame which is most conducive for evaporating and burning
the fuel despite the quite low pressure experienced at idle. The hot
flame does leave considerable equilibrium CO which must be consumed by
proper temperature control (>1500°K) in the'intermediate zone.
The biggest difficulty with the rich primary concept is that
because it is rich (excess fuel) at the high pressure high power condi-
tion, there is a strong tendency to produce smoke in the primary which
in turn must be consumed in the intermediate and dilution zones. This
is not readily done so smoke control is usually done another way, by
running the primary with perfect or slightly lean stoichiometry, not
rich. Thus, smoke control in this instance opposes HC, CO control.
It is also possible to have excessive HC levels despite the favor-
able flame temperature if the mixing and fuel distribution is inadequ-
ate, thus leaving pockets of excessively rich mixture which cannot burn.
Proper spray characteristics and atomization are thus required.
Secondary air admission
Primary air
admission \
Svvirler ~
flow
uuvv
SC^TT
# ®
. f r
Liner cooling air
-------
A-10
•
9. Lean Primary (1) - As noted above, this approach
intrinsically produces low smoke at high power. However, at idle the
mixture is even leaner. If done properly, this is beneficial because
excessively rich pockets are avoided, thus controlling HC, and the CO
level may be lower than in the rich primary case because the lower
flame temperature will result in a lower CO equilibrium level. In any
event, the CO will have to be consumed in the intermediate zone.
Difficulties with this approach lay in the possibility of exces-
sively lean pockets wherein the reactions may be quenched (Excessive HC
and CO) and in flame stability and relight (especially at altitude).
Also, as this approach utilizes more air up front in. the primary, there
is necessarily less for use in the aft portions for cooling and tem-
perature profile control (dilution). This can have serious ramific-
ations for combustor and turbine durability.
Secondary air admission
Primary air
admission \
Swirler ^^fli
flow -<> fV
Lean
15*?=
'Combustion
2one
Fuel
nozzle
r r
Liner cooling air
-------
A-ll
*
10. Delayed Dilution - By delaying the introduction of dilution
air, a longer combustion zone at intermediate temperatures is provided.
This increases the residence time of the reactants which allows the CO
to C07 conversion to approach equilibrium and for unburnt hydrocarbons
to be consumed. The temperature in the intermediate zone should ba,
however, low enough so that NOx formation rates are slow. The dif-
ficulty lies in adjusting the air flow into the intermediate zone pro-
perly at all power settings so it is hot enough for CO consumption, yet
cold enough to prevent NOx, and still achieving flame stability, liner
durability, etc.
\000O O/
0 0 O (
c
«
> O 0 o
r
1 0 O O O 0
oc
DO
1
Conventional Liner
Delayed Dilution Liner
Allison 250
Staging Techniques
11. Sector Burning - Sector burning is a circumferential fuel
staging technique designed to combine elements of the spray improvement
and rich primary control techniques (3 and 8) and at the same time not
affect in any way the combustor at high power (e.g., smoke from a rich
primary). If the baseline combustor has a lean primary, then at idle
-------
A-12
•
when the combustor is burning quite lean and at low flame temperature, .
the combustion efficiency is poor, resulting in much HC and CO, because
of inadequate heat to vaporize the fuel and to stimulate the CO • -»• C0~
reaction. This problem is resolved by cutting off the fuel entirely to
part of the combustor (usually about half) and injecting it with the
rest of the fuel into the remaining part of the combustor. This has two
beneficial effects: (1) the pressure drop across the nozzle is neces-
sarily increased, improving atomization and (2) the.fuel/air ratio is
increased (richened) so that a hotter flame exists, improving vaporiz-
ation of the fuel and enhancing the CO -> C02 reaction.
Hardware requirements include a split manifold, proper cracking
pressure of the valving in the nozzles, an additional valve in the fuel
control system to control the sector burning itself (with an override to
avoid it entirely while in flight) , and a sensing device to determine
flight vs. ground activity. Changes in the nozzle orifices may be
necessary to provide proper fuel flow at all power, regardless of the
sector (on or off at idle).
The primary concerns with this system are the reliability of the
more complex fuel control system and the possible degradation of the
V
turbine efficiency at idle (while sector burning) with an ensuing fuel
• penalty.
K.I n
n
Compressor
1
Burner
t
Turbine
.J
Hom.il
Operation
Fuel Sectored
at Idle
GE CF6-50
-------
A-13
NOx Control Technique
Air Flow Distribution Control Techniques
12. Quick Quench Primary - The idea here is to introduce the
intermediate air as close as possible to the upstreas dome or bulkhead
so as to minimize the extent of the primary zone. This zone, operating
at high power near to the stoichiometric point, produces a very hot
flame (2600°K) well in excess of that needed to activate NO production.
The reaction time to equilibrate NO at this temperature is only a few
milliseconds so in order to avoid significant NO production, it is
necessary to introduce the intermediate air quickly to quench the NOx
producing reactions (temperature < 1800°K). However, as the quench
temperature should still be in excess of 1500°K in order not to quench
the CO oxidation, great care is required to properly tailor the airflow.
The problem is that the quick quenching occurs also at idle and if tuned
to work correctly for NOx control at high power, it tends to quench the
HC and CO consuming reactions at idle producing excessive low power
emissions.
f
0
O
o
O
o
O
o
o
o
o
o
o
o
o
o
^J
c
u
£5
n
•~/
-^\
Q
Standard
Quick Quench
GE.CFM56
-------
A-14
*
13. Rich Primary (2) - A sufficiently rich primary at high power
will provide a cooler flame and a shortage of oxygen, both of which will
discourage NO formation. As there will be further reaction in the inter-
mediate zone where more air is available to burn the excess fuel (and
create MO), it is necessary to carefully tailor the airflow in order
to maintain the cool flame throughout. This approach is not satis-
factory, however, because of the high smoke levels and the generally
poor low power emissions (HC, CO) which arise from the excessively rich
primary which occurs at idle.
14. Lean Primary (2) - The lean primary zone is achieved by intro-
ducing a larger percentage of the total combustor airflow into the
primary zone. In sufficient amount, this creates a very lean, and
therefore, cool flame which prevents the formation of NOx by lowering
the N~ -> NO reaction rate.
Several problems are created, however, by this procedure. First,
the large amount of air into the primary creates a shortage of air
downstream for use in cooling and temperature profile control (dilution).
This may adversely affect the durability of the liner (cooling) and the
turbine downstream (temperature profile). Second, HC and CO-.emissions
}-
are very much affected adversely. Too cool a flame at high power may
quench the CO oxidation as well as the NO production. More importantly,
however, is the fact that being lean at high power implies a very lean
flame at low power (e.g., idle) so that under those adverse conditions
of low temperature and pressure, the fuel may not vaporize, the HC and CO
oxidation reactions will proceed too slowly, and in the extreme case,
will cease (flame instability). This also creates altitude relight
problems.
-------
A-15
•
Fuel Preparation
15. Spray Improvement - The spray improvements discussed in (3)
also affect NO production. Better atonization eliminates droplet
burning (locally stoichiometric and hot) and the spray angle affects
the fuel distribution (local hot spots). Better atonization is univer-
sally good for all pollutants, whereas a change in the spray angle may
adversely affect one or more pollutants while favoring the others, or
it may favor all. The outcome depends on other factors in the primary
zone, specifically the initial fuel-air distribution, the airflow
pattern through the doae and the amount of cooling air.
16. Premix/Prevap - Fuel and air are mixed in a prechamber prior
to entering the primary combustion zone. This prefixing allows combus-
tion to occur at a much leaner condition where NO formation rates are
slower. This technique is most applicable to high pressure ratio engines,
which produce the high combustor inlet temperatures required to sufficently
vaporize the fuel. With the premix concept careful attention must be.--,
given to the prechanber exit conditions. Exit velocities of the fuel-air
mixture must be high enough at all power levels to prevent flashback
which is very damaging to the liner and the nozzle. Also, in creating
a uniform lean primary zone, stability may be a problem leading to alti-
tude flaiaeout and difficulties in relighting. Low power emissions(HC,
CO) may be a problem if adequate mixing and vaporization do not occur
at idle where the conditions are much less favorable.
-------
A-16
0
The prersix concept requires a significant change to the combustor
liner geometry since the premix chamber must be included in the com—'
bustor. This may lead unavoidably to a longer overall combustor, thus
co najor changes in the engine configuration. The alternative, to
exchange some of the combustor length in the dilution zone for the
prenix mechanism, also can lead to complications because there is then
less space available to tailor the temperature profile at the turbine
ip.let. This compounds the already difficult tailoring job precipitated
by the shortage of dilution air arising from the lean (excess air)
primary (see 14).
PREMIX
TUBES 13 LOCATIONS
Premix coinbustor as seen on an axially staged(18) cornbustor
P&WA JT9D-7 :-
Staging Techniques
17. and 18. Fuel staging - The combustor is divided into two
regions, each having its own fuel injection system. These are termed
the pilot stage and the main stage. At low power, fuel is supplied only
-------
A-17
*
to the pilot stage, thereby allowing a much higher local fuel/air ratio
than would be possible if the fuel were distributed throughout the
combustor. This is basically the rich primary approach (8). This
mixture is then able to burn hotter, enhancing droplet evaporation
(aiding HC burning) and CO oxidation. Some form of fuel preparation
control (3 or 4) would also be incorporated.
At high power, the bulk of the fuel is injected into the main stage
which designed to burn lean for NOx control (concept 14) and low smoke.
Flame stability is provided by the primary stage which is still burning
rich. The pollutants produced by the primary at high power (CO and
smoke) should be consumed when diluted by the much larger main stage
flow.
Staging requires two manifolds and two fuel injection locations and
adds to the complexity of the fuel supply system and the fuel control.
The combustor liner is also more complex with additional cooling and
temperature profile problems. - •
There are two basic types of fuel staging here, radial (17) and
axial (18). In the former case, the stages are in parallel and fit more
readily within a short combustor volume. In the latter case, "the stages
are in series. Its primary advantage is that the upstream primary stage
«
is.better located to act as a flame stabilizing agent on the main stage
which may then be run more lean.
-------
A-18
\
Pilot z:nt
Pilot zone swirlsr
f'Jin zont
swirlsrs
Main zon»
fuel Injector
t:
Radial Staging (17)
GE CF6-50
Axial Staging (18)
P&WA JT9D-7
19. Variable Geometry - Variable geometry (or air staging) provides
airflow control of the primary and intermediate zones such that the
stoichiometry provides stable efficient combustion with a minimum of NOx
production over the complete operating range of the engine. Air enters
the combustor through holes equipped with a mechanism (usually a sliding
ring) that meters the airflow in proportion to the fuel flow. - With this
system the primary zone fuel air ratio can be controlled to be soichio-
metric at idle power for HC and CO reductions, and to be lean (but
stable) at high power for NOx reductions. However, as there is no flame
stabilizing mechanism here, the degree of lean burning, and hence the
NOx reduction, is limited. :
This system has a number of operational drawbacks, primarily the
reliability of the mechanical system in such a severe environment which
is a safety issue. However, the notion of moving mechanical systems in
severe environments is not new to gas turbines; variable pitch com-
-------
A-19
*
pressor stators and variable turbine nozzle guide vanes do exist. Failure
of the mechanism, however, must not prevent the engine from providing
adequate operational performance over its flight regime.
Catalysis
«. *
20. Catalysis - Catalysis is a process by which a special sub-
• :*•*
stance, usually a solid substrate, causes the acceleration of a chemical
reaction while not being permanently affected itself. Catalysis is
often used on current automobiles to limit the emissions of HC and CO by
r
the placement of the catalyst in the exhaust gas so that these pol-
'lutants, which are products of incomplete reaction in the cylinders, can
be consumed.
In an aircraft engine, however, the catalyst would have to be
-------
A-20
*
placed in the combustor proper. This then permits the reaction to
proceed under uniformly lean conditions, thereby giving a cooler flame
and less NOx production while still consuming the KG and CO through the
enhancement of the reaction rate.
Primary development problems are getting the catalyst to work
quickly during the warm-up period, prevention of poisoning of the
catalyst, prevention of mechanical wear on the catalyst material,
prevention of excessive pressure loss through the catalyst bed, and
prevention of flashbacks into the premixed air-fuel upstream of the
catalyst. Recent investigations suggest that, if used for NOx, a
catalyst would probably have to be used in conjunction with variable
geometry in order to keep the stoichiometry within limits acceptable
to the catalyst.
Catalyst
Catalyst Combined with Axial Staging
General Electric
-------
A-21
Water injection
21. Water Injection - Water injected into the primary zone of the
combustor results in a lower primary zone temperature. This lower
pri-zary zone temperature in turn results in a significant reduction in
NOx as a result of the lower N_ -> NO reaction rate. However, if the
temperature is reduced too much, an increase in CO occurs due to the
quenching of the CO oxidation. Water flow rates equal to that of the
fuel flow rate are possible giving a 50% reduction in the NOx level.
The use of water injection presents a number of problems: (1) The
increased aircraft weight due to the mass of water carried may reduce
the useful payload of the aircraft. (Usually, hox^ever, water injection
results in increased thrust, and hence the payload possibly can be
increased if takeoff performance is the critical factor); (2) Higher
fuel consumption is required to maintain turbine inlet temperatures; (3)
Precautions must be taken to prevent ice formation in the water in-
jection system for operation at ambient temperatures below the freezing
point of water; and (4) water must be demineralized in order to prevent
turbine blade corrosion and pitting. The use of tap water results in
substantial turbine deterioration and .thus compromises safety and engine
reliability. Also, demineralized water can be very expensive (over
•SO.30 per gallon) depending upon the location. Logistics for deraineral-
ized vater may be a problem also, especially for those aircraft using
snaller, more remote fields.
-------
A-22
WATER
FUEL
WATER
DISCHARGE
CF6 COMBUSTOR
CF6 SPRAY NOZZLE
-------
A-23
*
New Procedures
22. Selective Azirauthal Burning (SAB) - This control technique for
HC and CO is an offshoot of the sector burning concept (11). The two
drawbacks to sector burning are recognized as (1) fuel penalty while in
operation as a consequence of the severely asymmetric loading on the
turbine, and (2) the potential hazard arising from a malfunction that
might cause sector burning in flight. At higher power, the turbine
would be damaged, and at flight idle, the engine might be unable to
accelerate. SAB removes both of these difficulties.
The essence of the method is to reduce the number fuel nozzles
in operation at idle so as to improve the atomization of those in use
and, especially, to improve the local stoichiometry. This is, of course,
precisely the procedure and rationale for the sector burning idea. In
this case, however, the distribution of those nozzles firing and those
off is more or less uniform around the annulus. For instance, in the
CF6-50, sector burning would include a solid 180 sector (15 nozzles)
off and the other 180 sector on. For selective azimuthal burning, on
the other hand, a typical arrangement might be to distribute 20 firing
nozzles into 5 segments of 4 each so that each segment will be separated
by 2 off nozzles [ie, 4 on - 2 off, 5 times around]. Many other arrange-
ments are possible and the optimum is chosen by the competing demands of
emissions control performance, operational performance (eg, relight),
and mechanical performance (eg, pattern, factor).
-------
A-24
This method eliminates the problems of sector burning because it
provides sufficient symmetry around the annulus. Thus, at idle, the
turbine efficiency and acceleration are not noticeably degraded, and
at low power above idle, the pattern factor is not so poor as to cause
significant turbine damage. It follows then that SAB can be used
successfully in flight, thus removing the need for much of the complex
plumbing and the failsafe mechanism needed to prevent a malfunction from
leading to inadvertent sector burning in flight.
There is, of course, a drawback to this procedure: it is not as
effective as sector burning in reducing HC and CO. There are two reasons
for this. First, in order to provide an acceptable temperature distribution
for the turbine, fewer nozzles can be turned off. Hence the stiochiometry
and atonization quality are comprimised. Second, quenching of the reactions
always occurs at the boundary between a reaction zone and an adjacent,
cool airflow. Sector burning has only two such boundarys, whereas SAB has
several, the number depending upon the geometry (eg, for the 4-2 case
mentioned above, there are 5 such quenching boundarys). There would be,
therefore, a motivation towards the minimum number of quenching boundaries
and towards fewer nozzles in operation at idle. This would be nothing
more than sector burning. The practical limit is the point at which the •-•
drawbacks of sector burning come into play, that is, where SAB can no
longer be used in flight because of possible turbine damage or failure
to accelerate, or where the degradation of the turbine efficiency leads
to significant fuel penalty. In fact, if the method were to be used in
flight on a normal basis, then there are other performance considerations
that weigh on the selectionof the actual geometry. The system must
-------
A-25
(1) provide stable combustion over flight envelope, (2) be able to
relight over sarae envelope, and (3) be able to accelerate front starting
power (sub-idle), as well as (4), accelerate from flight idle. .
-------
APPENDIX B
ENGINE DATA
-------
Engine: CF34. Thrust(KN) PR BR" S/V'(ra ) Idle
Baseline
Idle
Approach
Climbout
Takeoff
. 40
P T M M,
(atm) (°K) (Kpjs) (Kg/hr) f/a
3.1 447 4.0 173 .0119
430
1230
19.5 1480
19.5 6.4%
El
Comb .
Ineff. HC* CO* NOx* Sk
.0487 28.2 104.8 2.2
.0018 0.2 6.9 3.1
.006 0 ' 2.4 11.2
.0004 0 1.9 14.2 20
EPAP 53.1 205.0 24.9
Sector Burning at Idle " •
Idle
Approach
Climbout
Takeoff
Idle
Approach
Climbout
Takeoff
173**
.0146 6.7 38.1 3.3
20
EPAP 12.7 80.0 27.0
•
•
w
. .*EIs estimated from EPAP values.
**Assumed unchanged.
-------
Kiip. 1 UK : I'.
Mocl.PFRT
Idle
Approach
Climbout
Takeoff
•>!',(, Thrust- (KM) TK HU S/V(m )
IO/
!'. T M M .
(atm) (°ib (KK/S) (Kjjlir) f/a
3.6 463 360
10.6 617 1087
21.9 762 3042
25.6 779 3758
:,">.(. TJ.U
Comb .
Tncff. He: CO
.0189 8.0 51.7
.0013 0.4 4.0
.0003 0.1 1.0
.0002 0 1.0
EPAP 12.0 79.5
Sector Burning at Idle (15 of 20)
Idle
Approach
Climbout
Takeoff
360*
.0083 0.8 32.6
EPAP 1.5 51.7
Sector Burning at 6% Idle (15 of 20)
Idle
. Approach
Climbout
Takeoff
400*
.0057 0.3 23.4
U.le
•'('1 0">l I 1 i in)
F.I
NOx Sk
3.5
9.3
20.6
23.9
42.8
3.5
»
w
i
u>
42.8
4.3
EPAP
0.9 42.0 43.5
*Based on full annular performance. It is understood that with sector burning the turbine efficiency
may be degraded at idle to the point that additional fuel would be required to maintain power.
-------
Engine: CF6-32 ' Thrust (KN) PR
Idle
Approach
Climbout
Takeoff
145
P T.. M . M.- Comb.
(a Cm) (°K) (Kg'7s) (Kc/hr) f/a InefE.
2.6 427 443
9.7 599 1584
21.5 749 . 4772
23.1 766 . 5779
.0482
.0014
.0002
.0002
EPAP
Sector Burning Projection .
Idle
Approach
Climbout
Takeoff
443**
.0055
EPAP
Double Annular Projection
Idle
Approach
Climbout
Takeoff
.
.0117
.0046
.0005
.0004
BR
nc
36.1*
0.2
0.1 '
0.1
48.1
1.3*
2.0
2.0
0.6
0.1
0
S/V(m
24.
ET
CO
73.3*
5.5
0.6
0.6
102.1
18.7*
29.8
42.9 '
17.7
1.8
1.6
~A) Idle
8 4%
NOx Sk
4.3
11.0
29.0
33.2.
64.1
3.7
8.6 ' •
14.0
16.3 .
td
I
EPAP 3.2 72.6 35.6
*Estimated from EPAP values. .
**Based on full annular performance. It is understood that with sector burning the turbine efficiency
may be degraded at idle to the point that additional fuel would be required to maintain power.
-------
Engine: CF6-.6D
Thrust(KN)
PR
BR
S'/Vdn'1)
•174.8
24.5
5.9
24.8
Idle.
3.34%
El
Production (atm)
Idle
. Approach
Climbout
Takeoff
2.7
24.5
T. M
(°K) (Kg?S)
435 10.1
779 71.2
"s
(Kg/hr)
483
1728
5206
6304
Comb .
f/a Ineff.
.0132
.0244
.0489.
.0014
.0002
.0002
HC
36.0*
0.2
0.1
0.1
CO
76.9*
5.5
0.6
0.6
NOx
4.6
12.1
32.4
39.7
Sk
16
EPAP
43.3
96.5
65.7
Sector Burning at Idle
Idle
Approach
Climbout
Takeoff
483**
.0056
EPAP
Double Annular Prelection
Idle
Approach
Climbout
Takeoff
•
.0112
.0040
.0005
.0004
EPAP
1.1*
1.8
1.9
0.6
0.1
0
2.8
19.9*
26.1
41.0
14.8
1.8 .
1.6
61.5
4.6
65.5
4.0
8.7
15.6
18.1
35.2
»
w
i
Ln
*Estimated from EPAP values.
**Based on full annular performance. It is understood that with sector burning the turbine efficiency
may be degraded at idle to'''the-point that additional fuel would be required to maintain power.
-------
Engine: CF6-50C
Thrust(KN)
PR UR
S/VCnf1) Idle
224.2
29.8 4.4 27.2
3.4%
P3
Production (a tin)
Idle
Approach
Climbout
Takeoff
2.9
11.7
25.9
29.8
(O t/\
Iv/
429
630
786
820
H
(KK7s)
13.8
47.6
92.1
103.0
(Kgi'hr)
548
2394
7034
8554
Comb *
f/a Ineff.
.0110
.0140
.0214
.0231
.0765
.0010
.0001
.0001
11C
59,3*
0.2
0.1
0.1
CO
109.4*
3.9
0.7
0.7
NOx Sic
3.5* 13
12.0
29.1
33.9
EPAP
63.0 119.5
Sector Burning at Idle with Nozzle Modification
*Estiraated from EPAP values.
**Based on full annular performance. It is understood that
may be degraded at idle to the point that additional fuel
60.8
Idle
Approach
Climbout
Takeoff
548**
.0079
EPAP
Double Annular
Idle
Approach
Climbout
Takeoff
•
.0103
.0028
.0005
..0004
EPAP
0.7* 31.4*
1.0 37.1
1.8* 37.7*
0.5 .10.1
0.1 1.8
0 1.6
2.4 . 49.8
3 . 5*
60.8
4.0*
10.0
19.1
25.5
44.7
w
I
with sector burning the turbine efficiency
would be required to maintain power.
-------
Knj;ine: JT3U-7
Thrust(KN)
IVi.!.
I. !. j
J'.R
Kl.*
1'.. T M M, Comb.
Production (at:m) (°K) (Kp'/S) (Kj',/hr) f/u lm>ft.
Idle
Approach
Climbout
Takeoff
460
1400
3720
4525
.1561
.0043
.0011
.0008
IIC
149
2.6
0.8
0.6
CO
119
8.5
1.5
0.9
NOx
1.4
4.6
9.4
12.0
Sk
n
EPAP
356
294
31
Aerating Nozzle/Leaner. PZ_Combustor
Idle
Approach
Climbout
Takeoff
•0788 66.0 94.1 2.3
.0026 1..2 6.7 7.8
.0005 0.3 1.2 16.0
.0004 0.3 0.7 20.3
EPAP
m*m&*nauiHa*iL»*C*»f) ***vt*'tffi*rru****.wv**Ki*w.
158 232 53
w
i
Idle
Approach
Climbout
Takeoff
*EIs estimated from earlier data. '
-------
Engine: JTRD-9 • Thrust (KN) PR
Production
Idle
Approach
Climbout
Takeoff
64.5 16.9
p ** T** . M . M, Comb.
(atiu) (°K) (Kg'?S) (Kg7hr) f/a Ineff.
2.7 404 476
6.6 .530 1072.
14.7 673 • 3044
16.9 708 3744
.0168
.0037
.0008
.0007
EPAP
Aerating Nozzle/Rich PZ Combustor Projection
Idle
Approach
Climbout
Takeoff
•
.0053
.0010
.0003
.0002
• EPAP
Vorbix Projection
Idle
Approach
Climbout
Takeoff
.0049
.0013
.0022
.0015
EPAP
BR
1.04
HC
10.0
1.7
0.5
0.5
35.1
2.2
0.3
0.2
0.2
7.8
0.26
0.14
0.32
0.15
1.6
S/V(m
25
El
CO
34.5
9.4
1.7
~~LJL»_-
124.5
14.5
3.4
0.6
0.2
51.3
20.0 '
5.2
8.3
5.8
88.0
~1) Idle
.2 7 . 0%*
NOx Sk
2.9
5.6
14.2
17.9 23
. 52.2
3.3
6.4
16.0
20.3
59.1
2.4
• 5.3
8.6
11.2
36.0
»
w
i
CO
*Quoted value, unrealistically high.
**Estimates
-------
Engine: JT8D-17 Thrust (KN) PR
1'roductlon
Idle
/von roach
CllTnhoupn_ .
Takeoff
71.2 17.
*
P T M Mf Comb.
(a tin) (°R) (KR'/S) (Kg/hr) f/a Ineff.
531
1275
3588
4482
.0160
.0037
.0010
.0008
EPAP
Aerating Nozzle/Rich PZ Combustor
T.rllP
Approach
Climbout
Takeoff
Production
Values
.0049
.0010
.0003
.0002
EPAP
Advanced Vorbix (II-9) Rig Test
Idle
Approach
Climbout
Takeoff
2.87 412 14.2 514 .0100
6.83 535 30.9 1247 .0112
15.08 678 60.0 3553 .0164
17.40 714 67.1 4406 .0182
.0046
.0011
.0019
.0013
EPAP
BR S/Vdrf1) Idle
4 0.
HC
10.2
2.0
0.8
0.7
37.3
2.1
0.3
0.2
0.2
7.6
0.25
0.14
0.32
0.15
1.6
99
CO
31.0
8.5
1.0
0.7
112.7
13.7
3.2
0.6
0.2
49.4
18.9
4,9
7.8
5.5
83.1
25.2 7.0%*
El
NOx Sk
3.3
6.1
15.2
19.2 24
60.1
3.7
7.0
17.3
21.9
68.4
2.7
5.8
9.3
12.1 27
41.0
*5'C
»
w
1
vO
^Quoted value, unrealistically high.
**EIs estimated from EPAP values.
-------
Engine: JT
(c
Production
Idle
Approach
Climbout
Takeoff
8D-17 Thrust (KN) PR BR S/V(m ) Idle
/ 1 « Z
17.4 0.99 25.2 . 7.0%*
El
P. T M M Comb.
(atm) (°K) (Kgfs) (Kg/hr) f/a Ineff. HC CO NOx Sk
Advanced Vorbix (P&WA proprietary configuration)
Idle
Approach
Climbout
Takeoff
Production
Values
.0075 3.0 22.2 3.3
.0012 0.3 4.5 8.0
.0008 0 3.8 12.2
.0005 0 2.4 14.3
EPAP 10.0 85.9 53.3
Idle
Approach
Climbout
Takeoff
'
»
i
H-*
o
*Quoted value, unrealistically high.
-------
.TT9D-7
Thrust(KN)
PR
BR
S/VOt."1)
205.3
21.4
5.2
19. 3
El
P3
Production (a tin)
Idle
Approach
Climtjourr
Takeoff
3.6
8.8
19.1
21.4
(5l
447
588
736
M
20.6
43.7
81.0
»^LL
780
2109
6010
Comb.
f/a Ineff.
.0105
.0134
.0206
.0360
.0013
.0004
. OOQ4
HC
26.4
0.6
0.3
0.3
CO
57.0
3.3
0.4
NOx
3.1
7.4
31.6
Sk
4
4
EPAP
45.4
98.5
.61.8
Aojv\M m» ^ryyrl f/Illrh P7, CnmhnSf fT*"
Idle
Approach
Climbout
Takeoff
.0056
.0013
.0004
.0004
EPAP
Vorblx (S27K)
Idle
Approach
Climbout
Takeoff
.0038
.0029
.0004
.0003
EPAP
2.9
0.6
0.3
0.3
5.7
1.0
0.5
0.1
0.1
2.1
13.2
3.3
0.4
0.4
24.6 .
12.8
10.8
1.2.
1.0
30.2
2.9
7.4
23.1
30.8 <20
47.4
2.9
4.7
11.6 '
13.8 30
26.2
estimated from EPAP value.
-------
Engine: JT9D-70 Thrust (KN) PR
Production
Idle
Approach
Climbout
Takeoff
.228 24.2
P.* T * . M M Comb.
(a tin) (°K) (Kg?S) (Kg/hr) f/a Ineff.
4.1 465 853
10.0 612 2449
.21.6 764 ' 7199
24.2 793 8791
.0228
.0007
.0002
.0002
EPAP.
Aerating Nozzle/Rich PZ 'Combustor •
Idle
Approach
Climbout
Takeoff
•
.0044
.0007
.0002
.0002
EPAP
Vorbix Projection
Idle
Approach
Climbout
Takeoff
. •
.0031
.0026
.0003
.0003
- ... EPAP
BR
4.9
HC
12.0
0.3
0.2
0.2
20.0
2.0
0.3
0.2
0.2
3.8
• 0.9
0.4
0.1
0.1
2.0
S'/Vdif1
KI
CO
53.0
1.7
0.2
0.2
87.5
11.4
1.7
0.2
0.2
20.0
10.1 '
9.5
0.9
0.8
26.3
) Idle
n
NOx Sk
3.0
7.8
25.6
31.6 8
54.3
2.9
7.8
21.7
28.9 <10
48.5
3,3
5.8.
14 . 5
17.4
35.2
»
i
h-*
hO
*Estimated
-------
Engine: .TTRn_?nq • Thrust (KN) PR • BR • S/V(m ) Idle
Idle
Approach
Climbout
Takeoff
85.6
PY T M M
(a tin) OO (K!..,'?S) (KR/hr) I/a
19.2 1.62 25.2 7%*
El
Comb .
Incff. I1C CO NOx Sk
Aerating Nozzle/Rich PZ Combustor Projection .
Idle
Approach
Climbout
Takeoff
Idle
Approach
Climbout
Takeoff
544
1265 •
3511 '
19.2 . 4355
0.74 10.7 3.4
0.12 2.3 7.9.
0.02 0 16.7
0.01 0.9 21.3
j
EPAP 2.2 33.6 54.9 15
•
w
I
^Quoted value, unrealistically high
-------
Engine: M45H-01
Thrust (KN) PR
BR
S/VCm"1)
32.4
16.9 3.0
23.75
Idle
El
*3
Production (atm)
Idle
Approach
Climbout
Takeoff
3.0
6.5
14.6
16.9
rib
424
541
693
723
(Kg/SS)
5.24
10.8
21.2
23.8
Mf
(Kg/hr)
191
526
1498
1793
uomoi
f/a Ineff.
.0101
.0135
.0196
.0209
.0935
.0183
. 0024 •
.0021
HC
59.5
7.4
0.7
0.8
CO
178.4
51.0
7.9
6.2
NOx Sk
1.5
3.6
9.3
11.5 46.3
Double Blown Ring
EPAP
162.4 526.0 31.2
Idle
Approach
Climbout
Takeoff
200
508 .
1444
1753
.0222
.0043
.0009
.0006
EPAP
•
10.7
1.0
0.2
0.2
30.1
55.5
14.7
3.0
2.0
169.9
2.1
5.1
10.9
13.1 12
37.0
* Normal minimum idle is 7.6%
-------
F.iiK inc: Spry Mkr>!>5
Thrust(KN)
PR
BR
S/V'dif1)
• 43.8
16.1
1.0
38. 7
Id.le
El
Idle
Approach
Climbout
Takeoff
2.1 388 6.86 341
7.0 546 22.2 793
14.2 667 39.. 5 2126
16.1 698 43.3 2606
RAB
Idle
Approach
Climbout
Takeoff
301
785
CVS
Idle
Approach
Climbout
Takeoff
305
785
.0138
.0099
.0150
.0167
.141
.0118
.0005
.0048
EPAP
.0122
.0098
.0235
.0029
.0008
.0008
EPAP
.0124
.0098
.0380
.0029
.0008
.0008
EPAP
130.0
8.3
0.5.
5.1
441
11.4
1.7
0.1
0
36.1
24.3
2.0
0
0
75.6
117.7
• 20.0
0
1.1
420
57.4
4.9
3.9
4.3
186.1
72.0
5.1
3.6
3.4
232.0
0.9
5.9
14.7
19.0
49.5
3.4
7.9
14.0
16.1
55.2
3.1
8.2
13.7
15.4
54.2
»
i
i — *
01
-------
Knr,lm>: Spcy MU511 Thrust (KM) PR
Production
Idle
Approach
Climbout
Takeoff
50.7 19.9
P
P T_ M Mf Comb.
(atin) (°I() (Kg'7s) (Kg/hr) C/a Incff.
2.2 407 7.5 401 .0149
8.0 575 23.6 998 .0117
17.1 700 47.2 2619 .0154 '
19.9 734 52.4 3202 .0170
.094
.011
.0016
.0012
EPAP
RAB
Idle
Approach
Clirabout
Takeoff
370 .0137
990 .0116
.0165
.0002
.001
.0009
EPAP
CVS .
Idle
Approach
ClimbouC
Takeoff
377 .0140
900 .0116 •
.0355
.0022 •
.0008
.0005
EPAP
HR
0.64
HC
76.7
7.2
1.3
1.0
278.4
6.7.
1.4
0
0
23.0
22.9
1.6
0
0
75.5
S/V(m •*-)
38.7
El
CO
117.4
20.3
2.1
1.8
435.8
46.3
3.3
4.3
4.1
161.6
67. Z
3.6
3.4
2.3
229.0
Til IP
Min.
NOx Sk
1.1
7.9
19.2
23.3
68.1
3.8
9.1
16.2
18.8
68.2
1.1
9.4
15.5
17.5.
58.0
w
I
-------
Elaine: KB'211-22IJ
1' T M Mf
Phase I (ntiii) (°K) (Kg?S) (Kg/hr)
Idle
— Approach
Climbout
Takeoff
3.6 446 18.1 627
10.6 616 46.1 2007
22.1 752 83.2 5555
25.0 781 91.5 6716
Phase II
Idle
Approach
Climbout
Takeoff
- 571
1989
5550
6712
Phase- II with Sector Burning
Idle
Approach
Climbout
Takeoff
568
Thrust (KN) PR
187 25JD
Comb.
f/a Ineff.
.0096
.0120
.0185
.0204
.100
.010
.001
.001
EPAP
.00875
.0120
.0185
.0204
.0131
.0007
.0005
.0004
EPAP
.0087
.007
Brt
5.0
lie
86.8
5.8
0.9
0.8
134.6
5.6
0.3
0.4
0.3
8.3
2.54
S/V(m )
20.11
CO
104 . 9
21.1
1.6
1.4
172.3
35.0
.1.9
0.7
0.6
49.6
19.6
Idle .
Min.
El
NOx Sk
2.3
8.2
25.4
33.2
51.9
4.3
12.4
29.0
34.3
61.7
5.5
w
I
EPAP
4.2
28.8
64.0
-------
Engine: K.15m-53!> Thrunr(KN) PR
Phase II
Idle
Approach
Climbout
Takeoff
142 19.3
P. T M M Comb.
(atm) (°IO (Kgfs) (Kg7hr) f/a Ineff.
3.1 424 15.8 544 .00955
8.3 574 36.8 1561 .0118
17.0 701 65.0 4340 .0188
19.3 727 71.6 5305 .0206
.022
.001
.0007
.0006
EPAP
Phase II with Sector Burning
Idle
Approach
Climbout
Takeoff
540 .00947
.011
EPAP
Phase III with Quick Quench
Idle
Approach
Climbout
Takeoff
539 .00946
.009
.002
.001
.001
EPAP .
BR
HC
11.1
0.28
0.46 .
0.41
19.1
4.9
8.9
0.84
0.55
0.49
0.50
2.5
CO
52.2
3.4
1.1
1.0
90.0
30.8
54.7
37.2
5.3
2.0
2.1
67.5
S/V(m ) T<11o
20.11 M.in.
El
NOx Sk
4.3
9.3
21.4
25.0
49.0
5.4
51.3
3.9
6.3
11.8
13.9
30.3
w
I
CO
-------
Engine: RB211-524
4
P T M M
Phase I (a tin) (°il) (Kg'?S) (Kgjhr)
Idle
Approach
Climbout
Takeoff
3.7 453 19.7 661
11.3 629 50.9 2498
23.9 771 94.4 6684
27.2 801 104.1 8120
Phase II
Idle
Approach
Climbout
Takeoff
609
2477
6677
8144
Phase II with Sector Burning
Idle
Approach
Climbout
Takeoff
604
Thrust (KN)
218
Comb .
f/a IneEf
.0093
.0136
.0197
.02167
.093
.009
.0007
.001
EPAP
.00854
.0135
.01966
.02165
.011
.0007
.0004
.0002
EPAP
.0085
.0058
PR
27.2
11C
79.8
5.8
0.7
1.0
110.4
4.5
0./4
0.3
0.1
6.0
2.1
BR
4.5
CO
99.8
17.8
0.7
0.8
145.1
.30.7
1.7
0.6
0.5
39.0
16.9
S/VCm"1) Idle
20.11 Min.
I'M
NOx Sic
2.5-
9.1
30.2
40.5
61.4
4.4
13.6
32.1
38.3
68.0
5.5
EPAP
3.1
22.4 70.2
-------
Engine: Olympus 593
Thrust (KM) PR BR S/V(m~
171
15.5
0
Idle
4.7%
P. T M Mf Comb.
Production (atm ) (°R) (Kgc?S) (Kg/hr) f/a Ineff.
Idle
Descent
Approach
• Climbout
Takeoff*
Afterburner **
1140
2360
4550
. 9100
15.5 12700
10000
.0584
.0380
.0201
. 0059
.0003
.0207
11C
3f>
22
8.5
1.5
0
6.6
El
CO
118
82
55
20
. 1.1
64.5
NOx Sk
2.5
4.0
6.5
12.5
22.3
0
EPAP
129
530
Blown Ring (Projected Worst Case)
70.1
Idle
Descent
Approach
Climbout
Takeoff*
Afterburner'''*
.0166
.0162
.0110
.0039 .
.0002
.0102
EPAP
7.2
7.7
2.9
0.8
0
•3.4
30.7
44.8
41.3
36.8
13.8
0.7
31.4
237
2.5
4.0
6.5
12.5
.22.3
0
70.8
w
i
. N)
O
*Data on this row refer to main burner only,
**0n during Takeoff only. .
-------
Fnjyinc: AT.F502D Thrust (KM) PR
Cert.Confi^.*
Idle
Approach
Climbout
Takeoff
.28.9 11.1
P T M Mr Comb.
(a Lin) (°K) (KB?S) (KB7hr) f/a Ineff.
2.3 397 3.6 168 .0130
5.0 499 8.0 354 .0123
9.8 ' 614 13.1 1005 .0213
11.1 639 14.4 1205 .0232
.0139
.0028
.0002
.0002
EPAP
Dual Orifice Pressure Atomization (Baseline)
Idle
Approach
Climbout
Takoof f
•
.0256
.0042
.0002
.0002
EPAP
Idle-
Approach
Climbout
TakeoEf
• -
M
5.8
HC
5.6
0.6
0.1
0.1
14.8
11.9
1.2
0.1
0.1
31.0
S/V(m" )
3M.O
F.:I:
CO
40.7
11.0
0.5
0.5
112.4
67.6
15.0
0.5
0.5
183.4
rdl.r
NOx Sk
!J'j»<^.«tAi*o'"»
-------
Ku K.I.I u,'i Al,r.>02l,
Thrust(KM)
33.4
PR
BR
S/V(m ')
13.3
3H.
Baseline (atm)
Idle
Approach
Climbout
Takeoff
2.1
5.6
11.7
13.3
T
381
511
636
662
M
(Kg?S)
3.76
8.06
15.3
16.82
MF
ICn 7 liir )
173
478
1185
1416
Comb .
f/a Ineff.
.0128
.0165
.0215
.0234
.0240
.0020
.0001
.0001
HC
12.6
0.3
0 •
0
CO
56.0
8.6
0.7
0.7
NOx
TWr^Tgua^raug^rno*1**
2.6
4.7
10.7
12.1
Sk
5
13
25
25
EPAP
28.6
136.0
Idle at 10.7%
32.3
Idle
Approach
Climbout
. Takeoff
Idle
Approach
Climbout
Takeoff.
3.18 426 5.29 200 .0105
.0098 3.4 31.8 3.8
EPAP 9.1 92.2 35.4
•
w
I
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