AIRCRAFT TECHNOLOGY ASSESSMENT
STATUS OF THE GAS TURBINE PROGRAM
December, 1976
Prepared By:
Richard Munt, Eugene Danielson
U.S. ENVIRONMENTAL
PROTECTION AGENCY
-------
I UNITED STATES ENVIRONMENTAL PROTECTION AGENCY
' WASHINGTON, D.C. 20460
Office of
Air and Waste Management
December 15, 1976
Preface
This is a report which presents an assessment by EPA technical
staff of the status of development and application, and of effort made,
by aircraft engine manufacturers to comply with the EPA standards appli-
cable to aircraft gas turbine engines. The report is based on analysis
of information provided to EPA during public hearings on turbine engine
emission control technology held on January 27 and 28, 1976, and on
information provided to EPA subsequently through December 1, 1976.
The report is intended to serve as a basis for staff advice to
the EPA Administrator on the feasibility of industry compliance with
currently promulgated emission standards for subsonic turbine aircraft.
Some reviewers of the report may conclude that data in the report are
incomplete or obsolete. We solicit from readers of this report any new
or additional information bearing on either the accuracy or the complete-
ness of the information presented.
Comments should be sent to the Director, Division of Emission
Control Technology, Environmental Protection Agency, 2565 Plymouth Road,
Ann Arbor, Michigan, 48105.
Eric 0. Stork
Deputy Assistant: Administrator
for Mobile Source Air Pollution Control (AW-455)
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Table of Contents
Summary [[[
Section I . Foreword ........................ . ................. 1
Section II. Introduction ............................. ........ 3
Section III. Emissions Control Technology .................... 11
Section IV. Leadtime Requirements ............................ 93
Section V. Industry Status ................................... 123
Bibliography ................................................. 201
Appendices
A. Programs for Reduction of Engine Emissions ........... A-l
B. Data Bank .................... ........................ B-l
C . Glossary of Emissions Control Concepts ............... C-l
D. Derivation of NOx Technology Curve
(Production Engines) ................................. 0-1
E. NOx Calculations for Newly Certified
T Class Engines ............. . ........................ E-l
F. Newly Certified P2 Class Standards - Computation ..... F~l
G. Calculation of NOx Trend Line for
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Summary
The principal findings of the EPA aircraft gas turbine technology
assessment program, as reported herein, suggest a qualified optimism for
the eventual control of a majority of those engines which may be expected
to*be produced in significant numbers in the future. The major area of
difficulty is the group of engines for which even marginally adequate
technology has not been identified.
In general, significant technological advances have been made.
Two basic technologies of disparate complexity have evolved. The first
and relatively simple technology controls only hydrocarbons (HC) and
carbon monoxide (CO). This control is achieved through changes in the
engine operation, modest changes to the fuel injection system, and
modest changes to the airflow pattern within the combustor. Because of
the low level of complexity generally found in it, this first technology
is easily retrofitable in most applications.
The second and considerably more complex technology is capable of
significant control of the oxides of nitrogen (NOx) and when used in
conjunction with techniques of the first technology general control of
all three pollutants is realized. This control of NOx is achieved
through exotic methods of fuel preparation (e.g., premix/prevaporization)
and multiple zones of combustion (e.g., staging). Combustion systems
and the auxiliary fuel controls which incorporate these concepts are
radically different from existing conventional systems and consequently,
such new systems do not lend themselves readily to retrofit. Also, the
greater complexity of this technology necessitates longer periods for
both development and proving.
The following table summarizes the EPA technical staff's assessment
of the prospects of each engine meeting the levels specified in the 1979
standards, based on the data available to the EPA as of December 1, 1976.
The table is divided into two segments, one for compliance with the
hydrocarbon and carbon monoxide standards and the other for compliance
with the oxides of nitrogen standard. The overriding limitation of the
validity of this table is the accuracy and completeness of the data
which has been supplied to the EPA by the manufacturers. Some manufac-
turers have supplied relatively voluminous data; others very little.
Further, data that are presented may occasionally be limited only to the
more conservative or to that which tends to be favorable to the industry
position; the true potential of a low emissions concept may not be given
proper credit here as a result. Consequently, industry may well have
progressed further than is indicated in this summary.
iii
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Again, qualified by the small amount of data supplied by the industry
at large to EPA, the control of hydrocarbons (HC) and carbon monoxide
(CO) appears quite successful in most cases with techniques which can be
considered acceptable for retrofit, if necessary. This control can
generally be available by 1979-80. It is of course possible that this
judgment is conservatively tempered by the nature of the data.
In contrast, control of the oxides of nitrogen (NOx) is far more
uncertain. In fact, only among some of the largest 12 (commercial jet)
class engines has the requisite technology advanced to the stage of
being demonstrated in an engine. For the Tl and APU classes, in parti-
cular, the status figures, though optimistic, are based largely upon
projections of limited demonstrations. While effective control may be
generally available to the Tl and T2 classes by 1982 and to the APU
class by 1981, though not always to the level dictated by the standards,
(e.g. CF6), there is considerable question as to the practicality of
incorporating such mechanisms into Tl and APU class engines. The P2
class generally already complies with its NOx standard. The techniques
necessary for effective control of NOx are considerably more complex
than that required for only HC and CO control and consequently do not
lend themselves as readily to retrofit.
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Industry Statue
Engine 1979
Tl Class (small jet engines)
1. AiResearch
TFE731 good
ATF3 unknown
2. Pratt & Whitney
of Canada
JT15D
3. Avco-Lycoming
ALF502 fair
4. General Electric
CJ610 poor
CF700 poor
HC and CO prospects**
1980 1981
later
fair-
5. Pratt and Whitney
JT12A poor-
6. Rolls Royce
M45H
RB401 good
T2, T3, T4 Classes
(large jet engines)
good
1. Pratt and Whitney
JT9D-7
JT9D-70
JT10D
JT3D
JT8D
2. General Electric
CF6-6
CF6-50
CFM56
good
good
good at certification date
poor
fair
good
fair
fair
3. Rolls Royce
RB211-22B fair
RB211-524 good
SpeySll poor
Spey555 poor
P2 Class (turboprop
engines)
1. AiResearch
TPE331 good
2. Pratt and Whitney
of Canada
PT6A-27 good-
PT6A-41 good
3. Avco-Lycoming
LTP101
T5321A
PLT27
fair good
already met
already met
4. Allison
250 good
501
5. Rolls Royce
Dart fair-
Tyne unknown
APU Class (on board power)
good
1. AiResearch
GTCP85
TSCP700
GTCP36
GTCP30
GTCP660
good
already met
good
already met
good
2. Pratt & Whitney
of Canada
ST6 already met
3. Solar
Titan-39 unknown
NOx prospects**
1980 1981
unknown
1982
good
fair
good
1983
later
already met
already met
already met
good
good(1979)
good
good at certification date
poor
good
poor-
poor-
good
good (date unknown)
good (date unknown)
unknown
unknown
already met
already met
already met
good (1979)
already met
already met
already met
already met
good
unknown
poor-
unknown
unknown
good
good
* GE NOx technology available but inadequate for compliance.
** Prospects for meeting the existing levels specified by the EPA 1979 standards for newly
manufactured engines by January 1 of the indicated year.
iv
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Section I. Foreword
On July 17, 1973 regulations controlling the gaseous and smoke
emissions from aircraft engines were promulgated (reference 66). Except
for the fuel venting requirement and a few specific cases of smoke
standards, these regulations are scheduled to go into effect in 1979 for
all newly manufactured gas turbine engines (turbojets, turbofans, turboprops,
and auxiliary power units), in 1980 for newly manufactured, non-radial
type piston engines, and in 1981 for newly certified (i.e., newly designed)
turbojets or fans. The interval between the promulgation date and the
compliance dates is intended to permit the development of the necessary
technology as required by the Clean Air Act (1970).
The purpose of this report is to summarize the status of the
development of low emissions technology for aircraft gas turbines. The
basic issues of the investigation are:
(1) What is the available technology; what are the concepts
involved and how effective are they in reducing aircraft gas
turbine emissions?
(2) What efforts have the manufacturers made and with what success
in their attempts to comply with the standards and what more
could be done, if necessary, in a timely manner?
(3) How much time is necessary to fully incorporate a certain
technology into production?
t
These issues are treated from both a general perspective (Section III)
and a specific or engine-by-engine perspective (Section V).
This report attempts to address the questions through various
analyses. The major qualification of these analyses is the uncertainty
of the completeness and accuracy of the data which has been supplied to
the EPA by the manufacturers. Some manufacturers have supplied relatively
voluminous data; others, very little. Further, data that are presented
may occasionally be limited only to the more conservative or to that
which tends to be favorable to the industry position; the true potential
of a low emissions concept may not be given proper credit here as a
result.
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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," (reference 90) 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"
(reference 26), which offered the best projection at that time of the
feasibility of control with the knowledge then available, and third, by
promulgating standards limiting the emissions from aircraft engines
(reference 66).
In keeping with the spirit of the instructions to determine the
technological feasibility of control and to allow sufficient time to
permit the development and application of the technology, the EPA has
established an Aircraft Technology Assessment Program for the purpose of
monitoring the many programs for the development of the low emissions
technology for aircraft engines, both piston and gas turbine. This
program for gas turbines was begun in July 1974.
Through the cooperation of the industry and the several government
agencies working in this area, the EPA received in the past two years
information from which it has attempted to assess the status of the develop-
ment of aircraft engines having low emissions. In addition, the recent
hearing in Washington, January 27-28, 1976, which was originally intended
for response to the Notice of Proposed Rulemaking (NPRM) for the retrofit of
large T2 class engines with low emissions combustors .(reference 92) was
expanded to permit testimony relating to other aspects and issues of the
aircraft emissions regulations. This permitted the acquisition of recently
generated data from the industry and NASA. Appendix A lists many of the
programs which have bearing on the development of low emissions combustors
that were brought to EPA's attention. Data from these programs and from
the manufacturers' own in-house efforts have been digested and reduced
to more manageable proportions while at the same time proper consideration
was being given to the many caveats attached. The EPA has been assisted
-3-
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in its data gathering and reduction efforts by Arthur D. Little, Inc.
Appendix B summarizes the technical data; most of the data points of the
figures throughout this report are tabulated therein. Leadtime informa-
tion was drawn directly from its source which is referenced in the text
as presented.
The methodology used in assessing the development status and control
potential of low emissions technology is explained in the text as re-
quired. In essence the approach taken is to categorize all potential
control techniques primarily according to complexity and pollutants
controlled. Control effectiveness is then examined in general terms by
the degree of complexity. Next, engines are examined individually to
discover (1) their best developments, again by the degree of complexity,
(2) what might be expected of them if they employed the best control
concepts identified in the control effectiveness study, and (3) the
leadtimes necessary. The methodology used in the leadtime analysis
involves the development of a general hypothetical schedule for each
class of engine on which each specific engine is located in accordance
with EPA's understanding of its particular situation. This permits a
ready determination of the years to go to production for each category
of complexity.
This report is a. principal product of ;the EPA technology assessment
program and it focuses primarily on the successes and failures of the
industry in its attempts to develop the means to comply with the 1979
standards for newly manufactured gas turbine engines. Attention has
been given to both the emissions levels that may be achieved with the
technologies presently identified and under development, to the leadtimes
necessary to incorporate those technologies into engines, and to the
problems of application of these technologies. Also, the same attention
has been given to the investigation of the proposed retrofit program for
in-use engines and its special problems. In addition, the EPA has found
it worthwhile here to examine in a preliminary fashion the likelihood of
compliance with the later 1981 standards.
Contrary to the anticipation expressed in the introduction to the
Interim Report (December 16, 1975), this report does not address the
evaluation of piston engine development. Piston engines will be considered
in a separate document in the future.
This report specifically considers:
(1) The identification of emissions controlling factors;
(2) The emissions performance of individual control concepts
and categories of concepts grouped according to complexity;
(3) The availability and suitability of these control techni-
ques in general and to specific engines;
-4-
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(4) The best demonstrated performance of each engine within each
category of complexity of control technique;
(5) The predicted performance of each engine employing the best
recognized applicable control technique in each category of
complexity;
(6) The amount of effort put forth by each manufacturer as
reflected in the number of and diversity of control
techniques investigated or success in reducing emissions;
(7) The development problems and emissions problems of each
engine, especially those which appear to be in jeopardy;
(8) The leadtimes necessary to incorporate control techniques of
different complexity into production engines;
(9) The potential for compliance with the newly certified engine
standards.
The aircraft emission standards (CFR, part 87, described in FR Vol.
38, N. 136, July 17, 1973, p. 19088), for which this report and the
aircraft technology assessment program have been undertaken, describe
standards for the emission of unburnt hydrocarbons (HC), carbon monoxide
(CO), oxides of nitrogen (NOx), and visible smoke from aircraft gas
turbine engines, newly manufactured and newly certified. The standards
for newly manufactured gas turbine engines are presently scheduled to go
into effect January 1, 1979, a date selected to allow time for proper
development of the requisite technology, while the more stringent stan-
dards for newly certified gas turbine engines are scheduled for two
years later, January 1, 1981 (as a newly certified engine would allow
more latitude to the designer to incorporate additional low emission
features, standards for such engines are set commensurately lower). In
addition, there are smoke standards applicable to certain in-use engines
at earlier dates. The standards are directed at the reduction of aircraft
emissions including smoke in the vicinity of major airports which are in
or adjacent to major urban areas wherein aircraft emissions will contri-
bute to the general degradation of the air quality.
The standards themselves apply not to the aircraft, but to the
engines. The approach to regulate the engines directly instead of the
aircraft was taken to minimize the enforcement effort as the same engine
is often used in several aircraft applications. Furthermore, all air-
craft in which a given engine might be used have generally similar
flight profiles and hence have similar emissions signatures in the
airport environs.
In accordance with the latitude permitted by the Clean Air Act
(1970), EPA found it desirable to establish classes of aircraft and
corresponding engine classes to which different sets of standards would
apply as determined by the technical, economic, and safety constraints
relevant to each class. The classes established are as follows:
-5-
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Table II-l Summary of Aircraft Classes
Type Aircraft Application
Piston engines Light general
(excluding radials) aviation
P2 Turboprop engines Medium to heavy general
aviation; some commercial
air transport
Tl Small turbojet/fan General aviation
engines jet aircraft
T2 Large turbojet/fan Commercial subsonic
engines intended transports
for subsonic flight
T3, T4 Special classes Commercial subsonic
applying to specific transports
engines for the purpose
of instituting early
smoke standards
T5 Large turbojet/fan SST
engines intended for
supersonic flight
APU Gas turbine auxiliary Many turbojet/turboprop
power units
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 norma-
lized with respect to the total power output of the engine over that
cycle (units given in Table II-2). As such, larger engines performing
greater useful work are permitted proportionally larger amounts of
total emissions over smaller engines.
The standards, promulgated in July, 1973 for all classes but T5 and
in July, 1976 for that class, are summarized below:
-6-
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SMOKE STANDARD
CLASS T2 ENGINES
50
40
0)
2
-------
oo
I
0)
2
0)
1
W
50
40
30
20
10
SMOKE STANDARD
CLASS P2 ENGINES
1234
Engine Rated Power (1000 shaft Horsepower)
Figure 2
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Table II-2
Newly Manufactured Engines
Class
Tl
T2
T3
T4
T5
P2**
APU***
HC
1.6
0.8
3.9
4.9
0.4
EPAP*
CO
9.4
4.3
30.1
26.8
5.0
NOx
3.7
3.0
9.0
12.9
3.0
Compliance Date
January 1, 1979
January 1, 1980
January.1, 1979
Smoke is also regulated. For the T1-T5 classes, the acceptable smoke
level is given by Figure 1 and for the P2 class by Figure 2. The
standards for advanced engines are:
Newly Certified Engines
Class
T2
T5
HC
0.4
1.0
Table II-3
EPAP*
CO
3.0
7.8
NOx Compliance Date
3.0 January 1, 1981
5.0 January 1, 1984
In addition, there has been proposed (FR Vol. 38, N. 136, July 17,
1973, p. 19050) a regulation which, if promulgated, would require all
(including those already in service as of January 1, 1979) large (i.e.,
thrust > 29,000 Ibs.) in-use engines of the T2 class to comply with the
T2 class standards of 1979 for HC, CO, NOx, and smoke. As this would
effectively require a retrofit program for the older engines (pre-1979),
the compliance date was proposed to be January 1, 1983, thus allowing
four years for that retrofit to be accomplished.
*pounds of pollutant per 1000 pounds (thrust)-hours over the LTO cycle,
except as noted.
**pounds of pollutant per 1000 horsepower-hours over the LTO cycle.
***pounds of pollutant per 1000 horsepower-hours at maximum power.
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Section III. Emissions Control Technology
A. Introduction
This section will first describe the present aircraft emissions
situation in order to give the reader some perspective of the demand
that the regulations impose on the industry. At the same time, certain
characteristic trends in emissions performance will be highlighted and
explanations given or attempted. The need for work on control develop-
ment will be obvious and the section secondly will deal with the means
and effectiveness of control. Newly certified engines will be addressed
along with the basic issue of newly manufactured engines and retrofit
engines. In a later section (V) engines are reviewed individually with
emphasis on (1) extent of exploration, (2) accomplishment, (3) estimated
best performance, and (4) leadtime requirements.
B. Production Engines
The emissions performance of current production engines is pre-
sented first in order to properly focus on the problem. The data is
exhibited in a format which suggests a certain pattern or trend in
emissions. As it is not obvious that this should be the case, a brief
analysis is performed which investigated possible governing factors.
Table III-l presents a list of all production engines and for each
engine, data on the applicable class, the standards to which it must
comply, the manufacturer, and an estimate of the engine's production
potential which crudely measures the likelihood that the manufacturer
will attempt to comply.
Production potential is not usually available in hard figures.
Generally, though, the production of all engines can be grouped into
four categories for EPA purposes.
Production
Category Situation
I Engines already out of production; engines
certain to be out of production by the
compliance date for newly manufactured engines.
II Engines at or near the end of their production
run by the compliance date. The few, if any,
units produced after that would not be suffi-
cient to amortize the development and certi-
fication cost of a low emissions combustor.
Ill Engines in the broad middle of their production
run. It is possible to amortize the necessary
development and certification costs for emissions
control over the remaining production. It is
equally possible to consider a cost-effective
retrofit of the units produced prior to the
compliance date; there are sufficient units to
amortize that development and certification
costs and to realize significant air quality
gains.
-11-
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Class
Tl
T2 '
1979 Standard
HC CO NOx
1.6 9.4 3.7
(pounds of pollutant/
1000 pound-thrust-hours/
cycle)
0.8
4.3
3.0
Table III-l
Emission Performance - Current Engines
Baseline Emissions
Engine
TFE 731-2
TFE 731-3
JT15D-4
JT12A-8
CF700
CJ610
M45H
Viper 600
JT9D-7
JT9D-70
CF6-6D
Size
3,500 Ibs
3,700 Ibs
2,500 Ibs
3,300 Ibs
4,315 Ibs
2,950 Ibs
7,600 Ibs
3,750 Ibs
46,150 Ibs
51,153 Ibs
38,900 Ibs
Manufacturer*
AiResearch
AiResearch
PWAC
P&WA
GE
GE
Rolls-Royce/
Snecma
Rolls Royce/Fiat
P&WA
P&WA
GE
HC
6.6
3.6
12.4
5.2
9.1
14.7
6.0
17.2
5.3
1.2
3.4
CO
17.5
14.8
34.6
85
94
155
40.7
102
14.3
5.9
10.0
NOx
5.0
6.5
3.8
3.2
2.2
2.7
4.2
r.s
4.9
5.8
7.2
Sk/Std**
33/40
' 45/40
14.1/43
24/38
44/35
/34
10/20
5/19
13/21
Production ***
after 1/1/79
III
III
III
I
I
I
III- IV
I
III
III
CF6-50C
49,900 Ibs GE
4.3 10.8
7.7
13/19
III
1
T3
T4
X5 ****
API!
Same as T2
Same as 12
3.9 30.1 9.0
0.4 5.0 3.0
(pounds of Pollu-
tant/1000 horse-
power-hour of
power output)
RB211-22B
RB211-524
Spey 511
Spey 555
JT3D-1
JT8D-17
Olympus
GTCP85-98CK
GTCP30-92
GTCP36
GTCP660
TSCP700
ST6
Titan (T-39)
42,000 Ibs
50,000 Ibs
11,400 Ibs
9,850 Ibs
17,000 Ibs
16,000 Ibs
38,000 Ibs
290 eshp
100 shp
192 shp
1100 eshp
910 eshp
720 shp
40 shp
Rolls Royce
Rolls Royce
Rolls Royce
Rolls Royce
*****
P&WA (Smokey)
P&WA
Rolls Royce/
Snecma
AiResearch
AiResearch.
AiResearch
AiResearch
AiResearch
PWAC
Solar
14.0
8.3
35.7
46.2
34.2
1.1
16.2
0.2
0.1
0.3
0.2
0.2
0.3
0.6
20.0
11.9
35.7
63.2
40.8
12.8
67
7.5
4.8
19.8
7.9
0.8
2.6
21
7.3
8.9
7.8
4.3
3.8
6.8
8.9
6.4
3.4
5.2
4.9
6.0
4.8
4.4
<21/21
<20/20
50/30
41/27
53/25
28/30
28/25
N.A.
it
it
11
ii
ii
ii
*Abbreviations
PWAC - Pratt & Whitney Aircraft of Canada Limited
P&WA - Pratt & Whitney Aircraft
GE - General Electric
** SK/STD - Smoke level/Smoke Standard
*** Refer to text for definition of categories
**** i960 Compliance
***** Due for retrofit in 1980 with smokeless version
III
II- III
I
II - III
III
III
III
III
III
III
III
III
-------
Table III-l continued
Class
P2
1979 Standard
HC CO NOx
Baseline Emissions
4.9
26.8
3.0
(pounds of pollu-
tant/1000 horse-
power-hours/cycle)
Engine
PT6A-27
PT6A-41
TPE331-1
TPE331-2
TPE331-3
TPE331-5
TPE331-6
250
501-D22A
T5321A
PLT27
Dart
Tyne
Size
715 eshp
903 eshp
705 eshp
755 eshp
904 eshp
776 eshp
776 eshp
400 hp
4680 hp
1800 eshp
2000 shp
2280 eshp
5500 eshp
Manufacturer
PWAC
PWAC
AiResearch
AiResearch
AiResearch
AiResearch
AiResearch
Allison
Allison
Avco-Lycoming
Avco-Lycoming
Rolls-Royce
Rolls-Royce
HC
52
54
86
68
55
80
68
25
14
2
0.3
100
CO
66
81
48
44
44
52
49
128
30
17
3
125
No
NOx
7.0
7.2
6.4
6.2
8.0
8.9
8.4
5.3
6.2
6.0
12.8
12.9
data
Sk/Std
17/48
21/48
18/48
19/48
19/48
19/47
20/47
34/50
60/29
17/37
0/36
36/36
Production
after 1/1/79
III
III
III
III
III
III
III
III
II -HI
IV
IV
II
III
-------
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 dead-
lines to warrant a retrofit program.
There are many formats by which to exhibit and to examine the
emissions characteristics of a group of engines. Some are more desirable
than others because of their simplicity and relevancy to the problem at
hand. One format that readily meets these conditions is the display of
engines emissions (in terms of the EPA Parameter, EPAP, defined con-
ceptually in equation (1) below) as a function of engine rated thrust
(Discussion here is limited to jet thrust type engines as most anticipated
compliance problems will arise from these engines; turboprop engines as
a whole will be largely successful in complying with the standards and
hence do not require a detailed analysis). This format is both simple
(EPAPs and rated thrust are both known for almost all engines) and
relevant (it displays the standard explicitly as an EPAP value and shows
its thrust dependence which is fundamental to the standards through both
the thrust-based nature of the parameter and the T1-T2 class demarcation
at 8,000 pounds thrust). Combustor performance based formats (e.g.,
combustion inefficiency vs. air loading parameter), on the other hand,
are very useful for diagnostic purposes although they cannot in and of
themselves show the emissions performance of an engine compared with the
standards. Therefore, the emissions data of Table III-l are also presented
graphically on Figures 3, 4, and 5 for jet type engines (T classes).
Table III-2 lists the symbols for the engines as they are used throughout
the figures except as noted.
Table III-2
Class Tl ffcTFE731-3
Ik ALF502D
TFE731-2
k. ATF3
JT15D-4
f JT12A
CJ610-2C
CF700-2D
A KB401
* M45H
k Viper 600
Classes T2, T3, T4 k CFM56
O CF6-50C
D CF6-6D
O JT9D-7'
A JT9D-70
L. JT3D-7
D JT8D-17
X JT10D
Q RB211-22B
O RB211-524
0 Spey 555
C* Spey 511
O Olympus 593
0 TF34
-14-
-------
A cursory examination of these curves shows that all production
engines fail to meet either two or three of the three gaseous pollutant
(HC, CO, NOx) standards. None meet the HC and CO standards, while only
a few meet the NOx standard. Further examination reveals an apparent
trend with engine size (thrust), primarily for the pollutants which rise
from incomplete combustion, HC and CO. The curves labeled "production
technology" represent that trend and xrere derived by the procedures
employed below to determine the cause of that phenomenon.
a. HC and CO Behavior
The parameter describing the aircraft standards (EPAP) can be re-
presented by two multiplicative terms, one representative of the combustor
performance (in terms of mass of pollutant j?er mass of fuel) and the
other descriptive of the fuel consumption efficiency of the engine (in terms
of mass of fuel per level of useful output), both terms being weighted
over the prescribed landing - takeoff cycle:
EPAP = (combustor performance) x (fuel consumption)
or
EPAP « mass of pollutant = mass of pollutant x mass of fuel '
useful output mass of fuel useful output
= El x SFC (1)
Since the bulk of the HC and CO emissions are produced in the idle mode,
the EPAP for these two pollutants can be approximated as follows:
EPAP - El (idle) x Mf (idle)/I (2)
where El (idle) = emission index at idle; Mf = mass of fuel consumed at
idle; and I = total impulse (thrust x hours; of the engine over the cycle.
From this, two important facts are ascertained: (1) The emissions
performance of an engine (in terms of EPAP, the parameter upon which the
standards are based) can be affected either by the combustor emissions
performance (El) or by the engine fuel performance (SFC); and (2) for HC
and CO emissions, the El and SFC at idle are paramount (Mf (idle)/I <*
SFC (idle)).
It therefore appears worthwhile to investigate the apparent EPAP
trend with rated thrust by first breaking the EPAP into its two compo-
nents and examining their behavior with rated thrust.
Figure 6 shows the idle SFC behavior of the T classes as a function
of the rated thrust.* A curve has been faired through the lower range
*
All engines are considered here to operate at idle at 6% of rated
thrust. Although this is not the specification of idle called for
in the regulations, nor is it even desirable from a regulatory point
of view, it is used here in order to reduce data scatter by putting
the engines on completely the same LTO cycle ( an "apples to apples"
comparison). , _, c_
-------
of SFCs which represent typical modern engines. The several outliers
(JT3D, JT8D, CF700, and the Speys) lie well above the curve because,
being older engines, their engine cycles (primarily the bypass ratio)
are inferior to the newer engines and consequently the SFCs are higheri.
The JT15D, shown also to be above the other curve, does possess an
engine cycle comparable to that of other modern engines of similar size;
however, it also appears to suffer very high inefficiencies at low power
which increases the idle SFC some 30% above what might be expected.
The significance of Figure 6 appears to be that there is a charac-
teristic trend of idle SFC with size (ignoring the older engines now)
such that smaller engines (those below 8-10,000 Ibs. thrust) can be
expected to have a higher idle SFC which by equation (2) would be expected
to exert an upward tendency on the EPAPs of HC and CO for those engines.
It is of some interest to determine the basis for this, primarily with
regard to newly certified engines (as modification of the SFC trend may
be considered an emissions control technique). For newly manufactured
or retrofit engines, there is no SFC control so that the character of
this SFC trend, fundamental or fortuitous, is a moot point. A good
point at which to begin is the investigation of the effect of the
engine cycle on the SFC. This is demonstrated in Figure 7 wherein a
theoretical calculation of SFC as a function of the engine parameters
pressure ratio and bypass ratio (overall temperature ratio is assumed to
be directly related to the pressure ratio) is presented and compared
with typical engines. Figures 8 and 9 then show how the principle cycle
parameters, the bypass ratio and pressure ratio, vary with size (rated
thrust) of the engine.* Figures 7, 8, and 9 support Figure 6; that is,
the faired SFC curve on Figure 6 can be closely approximated by cross
plots from Figures 7, 8, and 9. For instance, at 10,000 Ibs thrust,
Figure 8 gives BR = 5.3 and Figure 9 gives IPR =2.5 which on Figure 7
(interpolating) gives idle SFC - 0.63 while Figure 6 gives idle
SFC = 0.68. Thus, it is fair to conclude that the trend seen in Figure
6 is amply explained by the engine cycle trends shown in Figures 8 and
9. The reason for these trends seen in Figures 8 and 9 is found in
several technical and economic factors which are discussed below. There
is nothing fundamentally rigorous about the relationship and as technology
advances and economics change so also will the shape of the curves of
these figures. The historical background given below will also emphasize
the time dependent character of curves.
*ln Figure 8, the anomolous decrease in the bypass ratio for the over
40,000 Ibs thrust range is due simply to the fact that these largest of
engines are not totally new engines but are updated versions of slightly
smaller thrust originals (e. g., JT9D-7 -» JT9D-70) wherein the extra
thrust is obtained by increasing the core pressure ratio without decreasing
the physical size of the core. This effectively decreases the bypass
ratio because now more air is drawn into the core Vat the expense of the
surrounding bypass airflow. The additional core airflow permits more
fuel to be added (keeping the stoichiometry the same) and this additional
energy input is converted to additional thrust output. On the other
hand, a totally new smaller core would accept less air so that the
bypass ratio would be the same, or perhaps, even increased to obtain yet
better fuel economy.
-16-
-------
From a historical perspective, the first jet thrust type of turbine
engines were low pressure ratio, zero bypass ratio (pure jet) engines
with high SFCs. Although such engine, cycles were very undesirable they
were nevertheless unavoidable as the technology for higher pressure
ratios was not available. Fans (nonzero bypass) had not yet come into
existence and the advantages of them were not fully appreciated.
Later, as fuel economy and noise control became considerations in addi-
tion to power and thrust-to-weight ratio, low bypass fans (BR £ 1.5)
such as the JT3D and JT8D were developed as growth versions of original
pure jet (BR = 0) engines. There are a number of reasons why such
engines were limited to low bypass. A close examination of figure 10
shows that the largest SFC improvements are made in the initial stages of
bypass (say, BR < 2). For example, a PR = 10 engine reduces its TSFC
by .136 in going from a pure jet to a BR = 2 fan engine, but a doubling
of the bypass (to BR = 4) will yield a further reduction of only .083.
The smaller SFC improvement (from BR = 2 to 4) is coupled with an even
larger weight and drag penalty (from the bigger engine associated with
the larger bypass) which is significant for the whole aircraft system.
The only way to keep this penalty down is to have a high pressure ratio
core which is capable of generating more power for the fan in a smaller
system. As the original fan engines such as the JT3D were based upon
cores with only moderate pressure ratio (PR - 13), weight and size
restrictions as well as the technological infancy of the field dictated
low bypass ratios (BR £ 1.5). Later, as higher and higher pressure
ratios were made feasible by technological advances, bypass ratios rose
commensurately (Figure 11) to convert the higher thermodynamic efficiencies
into higher propulsive efficiencies.
This background then leads to the basic question: Why do the
smaller engines typically have smaller pressure ratios and lower bypass
ratios than the larger engines (Figures 8 and 9)? There are a number of
reasons for this. The first is economic: Large fan engines have higher
initial costs and although there is a fuel penalty with the low bypass
engines, the fuel savings of a large bypass engine may not pay for that
larger initial cost especially if the aircraft usage is low enough (ie.,
depreciation cost dominates). The second is technical: High bypass
engines, being heavier and more bulky, do not fit as easily into the
smaller business jets and their weight creates a fuel penalty which
partially offsets the fuel savings expected from the superior SFC. The
third is also technical and relates to the second: It is very difficult
to obtain high component efficiencies in small engines with a high
pressure ratio cycle because of the very small size of the components
(component size inversely proportional to pressure ratio).
Moving from the investigation of the idle SFC effect on the HC and
CO EPAPs, the next component of the EPAP to explore is the idle El.
Figures 12 and 13 present the idle Els for HC and CO of production
engines as a function of thrust. An exponential regression (the solid
line) through the data indicates that for conventional combustors, a
roughly 50% reduction in the El for either HC or CO can be expected as
the rated thrust of the engine increases from 0 to 50,000 Ibs. of the
-17-
-------
thrust.* Like the SFC behavior this will exert an upward tendency on
the EPAPs for smaller engines. Again, the question which must be
answered is why are these Els dependent upon the size of the engine?
There are three scale effects which are likely to have bearing on the El
behavior: (1) pressure ratio, (2) surface-to-volume ratio, and (3) air
loading.
The pressure ratio at idle has a significant effect on the combus-
tion efficiency** because, first, combustion at higher pressure enhances
in particular the CO -* CO reaction rate and also many of the hydrocarbon
reactions, second, the higher combustor inlet temperatures associated
with the higher pressure*** increases the fuel droplet evaporation rate,
thereby enhancing hydrocarbon reactions, and third, the higher pressure
into the combustor permits a greater pressure drop across the fuel
injection which can be used to shatter the fuel drops (airblast). The
significance of the pressure ratio to good combustion efficiency is
reflected in rig-to-engine pressure correction formulas used by NASA
(reference 59):
HC (eng) = HC (rig) x
CO (eng) - CO (rig) x
Figure 14 plots the combustion inefficiency (1-n ) vs. the idle
pressure ratio for production engines. Although there is considerable
scatter indicating the effect of the other factors on the combustion
process, the trend line (an exponential regression) shows roughly 30%
decrease in inefficiency between the lowest idle pressure ratio (1.5)
and the highest (4). Figures 15 and 16 show the El performance of the
individual pollutants vs. the idle pressure ratio.
*These regressions ignored the JT3D, the Spey, the CF700 and the
CJ610 engines. The emissions from these engines are not repre~
sentative of what is achievable by modern production engines, as
they represent old technology. The KB 211's and the M45H were
also ignored because of unusually high emissions, especially HC.
These are modern production engines, but their emissions are not
representative. The uniformly high emissions level of the Rolls
Royce engines may be due to the tendency of these engines to have
large idle air loading parameters (see discussion later in this
segment); however, this cannot explain the high HC, CO emissions
of the RB211-524 whose air loading parameter at idle is quite
modest.
**Combustion efficiency = n^ = 1 - [.875 x El (HC) + .2322 x El (CO)] x 10~3
***Ideally,
c
.286
-18-
-------
The second potentially significant parameter is the surface-to-
volume ratio. This is merely a measure of the relative amount of
surface area to which a given flow is exposed. The significance to
combustion arises because the liner surface must be cooled by a film
flow of air heated only to the compressor exit temperature and this
cooling air is capable of quenching certain reactions which demand a
higher temperature to proceed, in particular the CO -» C0_ reaction.
Thus, a combustor with a high surface-to-volume ratio would expose a
greater portion of its flow to the vicinity of the liner surface,
thereby quenching a greater fraction of the CO -> C0« reaction than would
a low surface-to-volume combustor which would then leave more CO in the
exhaust. A large surface-to-volume combustor would not normally affect
HC emissions too much unless the fuel nozzle were spraying fuel directly
onto the liner wall in which case the problem is not so much the surface-
to-volume ratio as it is the fuel nozzle. Figure 17 shows the effect of
S/V on combustion inefficiency. The trendline, again an exponential
regression, shows an effect comparable in magnitude to the pressure
ratio variation. Figures 18 and 19 demonstrate that the S/V effect is
almost exclusively on the CO emissions and not the HC emissions, as
discussed above. There is a point of confusion, however. Figure 20
shows that there is a general trend between higher idle pressure ratios
and lower S/V so that in either Figures 14 or 17 the underlying physical
effect may not be the parameter labeled on the abscissa, but rather the
other one.
The third potentially significant parameter is the air loading
parameter (52) a semi-empirical factor defined by
M
n = a
P 1>8 x f(T) x V
where
f(T) = exp T /540
Being an empirical type of expression, there may be some variation in
the definition among various users. The terms in this expression are:
M = combustor air mass flow rate (Ib/sec)
a
P_ = combustor inlet pressure (atmospheres)
3
V = combustor liner volume (ft )
T_ = combustor inlet temperature (°R)
The expression represents a rough approximation of the extent to
which the fuel combustion should have proceeded. Specifically,
(reaction rate) x (residence time)
-19-
-------
Thus, larger values of ft means that the product of the rate times the
time is small, indicating that either the reaction rate or the residence
time or both are small and that, in any case, the reaction cannot have
proceeded very far before the fuel exits the combustor. High values of
combustion inefficiency would then be expected. On the other side of
the coin, smaller values of ft indicate that the reaction should have
proceeded quite far before the fuel, what is left of it, exits. Low
values of combustion inefficiency would be expected.
A combustor for any given engine will experience a range of fts
depending upon the power level of the engine, higher power means lower
ft. The actual range of ft from the high power setting to the low will
vary between engines and is dependent upon such design variables as the
compressor pressure ratio, the size of the combustor, and the idle power
setting level.
The emissions performance of different control techniques (for HC
and CO emissions) can then be investigated systematically by comparing
the combustion inefficiencies at equal values of the air loading para-
meter.
Figure 21 presents the combustion inefficiency (1 - n ) vs. the air
loading parameter (ft) for idle only. The regression line through the
data verifies the trend expected and the correlation coefficient of 0.43
indicates that this variable alone is better than either IPR or S/V.
However, there is again a point of confusion: Just as there was found
to be a correlation between IPR and S/V, similarly there are found to be
correlations between ft and both IPR and S/V as shown in Figure 22 and
23. Thus, the effects of these three parameters are not totally indepen-
dent. The air loading parameter in particular obviously accounts for
the pressure effect on the combustion process and to that extent acts as
a substitute for IPR; however, it does independently account for the
residence time effect which neither of the other parameters do.
There are certain drawbacks to ft and S/V as correlation parameters.
What is needed here is a correlation parameter that is both relevant
(i.e., has a bearing upon the combustor emissions) and also characteris-
tic of the engine so that it remains constant for a given engine regard-
less of the combustor modifications employed to reduce emissions. All
three of the parameters discussed here are relevant, but two are combus-
tor oriented (S/V and ft through the residence time effect). The values
of these parameters are likely to change, perhaps appreciably, in a low
emissions design, making it difficult to compare the old and the new.
On the other hand, such parameters are excellent design tools for they
suggest what modifications should be made in order to achieve the desired
goal (e.g., increase V to reduce ft at idle to improve the combustion
efficiency). The third parameter (IPR) is, of course, an engine oriented
parameter and consequently it meets both of the requirements above. Of
course any engine condition parameter has the drawback that for a few
category 1 concepts (defined in part C) such as airbleed, that parameter
will change in going from the production version to the low emissions
version, whereas a combustor oriented parameter (e.g., S/V) would remain
-20-
-------
constant, as desired. Therefore, an engine based parameter is not
perfect.
In summary then, the idle pressure ratio (IPR) will be used to
compare the HC and CO emissions performance of the various control
concepts with that of production type engines. Figures 15 and 16 re-
present the baseline combustor performance against which the comparisons
are to be made. It is necessary also to demonstrate that the scale
effect represented by Figures 15 and 16 does indeed explain quantita-
tively the El trend alluded to in the discussion of Figures 12 and 13.
By combining Figures 15 and 16 with Figure 9, one obtains the dashed
lines in Figures 12 and 13. As can be seen, they approximate reasonably
well the regression curves. Furthermore, using the approximation of
equation 2, Figures 6, 12, and 13 can be combined to derive the "produc-
tion trend lines" shown in Figures 3 and 4. Following the several steps
leading up to these production trend lines, one can see that the root of
these trends is the characteristic tendency towards larger engines
having higher bypass and pressure ratios (which affect the SFC) and the
fact that combustion efficiency improves with higher operating pressures
(i.e. pressure ratio).
b. NOx Behavior
Oxides of nitrogen, or NOx, arise from the oxidation of atmospheric
and possibly fuel-bound nitrogen in lean, hot combustor conditions
wherein there is both oxygen and activation energy available. Thus, NOx
is primarily a product of high power operation although over the EPA LTO cycle,
only 50-60% of the total NOx is produced at climbout and takeoff: At
idle and approach the longer times-in-mode compensate for the lower NOx
formation rate. It is, therefore, impossible to make the kind of single
mode approximation represented by equation (2): All four modes must be
considered.
Figure 5 presented the NOx EPAPs for production engines. As present-
ed, two facts become apparent. First, all but a few Tl class engines
fail the standard. Second, throughout most of the thrust range the
average EPAP is relatively constant (- 6-7) with a scatter of + 2, while
at the lowest thrust levels (Tl class), a precipitous drop in EPAP
occurs. Once again equation 1 can shed light on this behavior although
the simplicity of equation 2 is not available. For reference,
EPAP = El x SFC (1)
EPAP * El (idle) x Mf (idle)I (2)
-21-
-------
Figure 24 shows the rated SFC trend of the T classes with size
(thrust). Here, rated SFC, not idle SFC, is considered as NOX is pre-
dominantly a high power pollutant, although as indicated above, not
exclusively. The trend to lower SFCs for larger engines would, by
equation 1, contribute to a reduction in the NOx EPAP unless the El
trend is such that it counteracts the SFC trend. That such is the case
is readily seen in Figure 25 showing El (NOx) increasing with rated
thrust. Once again the question must be asked: Are these trends with
size a fundamental characteristic or a circumstantial coincidence that
can be overridden as necessary to achieve a more favorable climate for
low emissions?
The SFC trend has been explored before in the HC and CO discussion
and the understanding obtained then is applicable in this instance with,
however, reference to the rated (full power) condition rather than the
idle. Specifically, the theoretical calculation of Figure 10 can be
combined with Figures 8 and 28 specifying the trend in the cycle para-
meters with size to generate approximately the faired curve on Figure
24. This then shows how the SFC trend with size (Figure 24) is caused,
once again, by the engine cycle trend with size (Figures 8 and 28). In
any event, SFC is not a variable to deal with for newly manufactured
engines (NME); however, it certainly is for newly certified engines
(NCE) for which there is a choice in the engine cycle.
To explore the El(NOx) trend, a proper , the equivalence ratio of the primary combustion
zone (the equivalence ratio is the ratio of the actual
fuel to air mixture to the stoichiometric, hence a
measure of both the excess 0 and the flame temperature
Tf, among other things.
Because of the relationship between these parameters and other combustor
parameters, these parameters can be replaced by the other, related para-
meters, as convenient. The most notable substitute parameters are:
*See Appendix C for the definition of the combustor stations.
-22-
-------
(1) Tf, the flame temperature,
(2) T,, the combustor exit temperature (or turbine inlet
temperature),
(3) V , the reference velocity,
(4) «S, the overall combustor equivalence ratio, and
(5) T_, the combustor inlet temperature (or compressor
discharge temperature).
Despite the many relevant variables, it has been discovered (by Lipfert)
that the NOx formation (as described by EI(NOx)) can be correlated quite
accurately with only one variable, T», as shown in Figure 26« This
empirical relationship holds very well for existing engines, despite its
simplicity, because existing engines are designed so similarly that for
all practical purposes T (or its equivalent, P_) is the only independent
parameter. Specify that parameter and all the other parameters are
reasonably well determined. Because T and P_ are related by the
adiabatic relation,
(Where T and P are atmospheric conditions and n is the compressor
efficiency.), trie Lipfert correlation can also be°presented as in Figure
27. Also, the rated compressor pressure ratio (defining P_ at takeoff)
tends to be higher for the larger engines as is evidenced by Figure 28.
There is nothing fundamental about this and in fact the same can be said
of its significance as was said earlier of Figures 8 and 9: "The reason
for this (the. le£a£iom>hi.p W-ith A^ize.) is found in several technical and
economics factors... There is nothing fundamentally rigorous about the
relationship and as technology advances and economics change so also
will the shape of the curves of these figures."
From the faired curve of Figure 28, it is possible to backtrack
using the Lipfert correlation (Figure 27) to arrive independently at an
approximation (the dashed line) to the regression (the solid line) drawn
on Figure 25. From the Lipfert curve (Figure 27), the rated pressure
ratio curve (Figure 28), the rated SFC curve (Figure 24), and the approxi-
mation explained in Appendix D with sample calculation, the production
-23-
-------
technology trend line in Figure 5 is arrived at showing a self consis-
tent set of curves with the data.
Up to now the examination of the NOx El data using either T~ or P~ as
a correlating parameter has been satisfactory in view of the accuracy of
the correlation (Figure 26). However, in examining low NOx technology
this proves to be inadequate as, in that case, the combustor design and
operating environment will differ greatly between combustors so that
one can anticipate a considerable increase in the scatter. These com-
bustor differences should be treated as independent variables leading
then to more complicated correlations than have been successful in the
past. This shall be demonstrated as the low emissions data are presented
and analyzed. The difficulty of the T. correlation in adequately extrapo-
lating NOx data to new conditions suggests its replacement by the
correction formula offered by NASA in reference 59 (ECCP summary to
EPA). This correlation, of course, also has its limits of applicability
which may be exceeded in the course of analysis. In the next part, the
investigation of low emissions techniques will begin.
C. Emissions Control for Existing Engine Types
As discussed in part B, the emissions performance parameter used by
the EPA (EPAP) can be affected by both changes in the specific fuel
consumption (SFC) and in the combustor emissions performance (Els) (see
equation (1)). Furthermore as evidenced in Figures 7 and 10 the SFC is
controlled by engine cycle parameters (pressure ratio, bypass ratio, and
turbine inlet temperature) over which there is little or no chance of
control in existing engines for to change any of these parameters sig-
nificantly would imply a new engine, different from the original. It
follows then that for newly manufactured engines (NME) and retrofit
engines effective emissions control can be achieved only through combus-
tor control. This part addresses the concepts used in combustor emissions
control, their problems and performance. Some of these schemes are more
esoteric than others and, as such, are far from proven. Others are more
modest in approach and are more readily adapted into present-day flight
hardware. Unfortunately these latter techniques, while often very
effective against the low power pollutants (HC and CO), usually cannot
appreciably improve NOx emissions. Furthermore, while many of these
control techniques are simple in concept, many are nonetheless difficult
to apply in practice because they are often detrimental to the combustor
performance in non-emissions related areas such as altitude relight,
durability, etc. Lastly, some concepts are better suited to one class
of engines than another for various reasons, technical and economic.
Because of this diversity among the potential control concepts, it
becomes advantageous to evaluate them with respect to several factors.
From the remarks above, the primary factors should be:
-24-
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(1) development status
(2) control capacity
(3) complexity of the development process
(4) application
For each factor, a scale of one to four is taken to describe the rating
or relative position of a control technique. The factors and the meaning
of the numerical assignments are given in Table III-3. In addition, for
the emissions performance category the particular pollutants that are
controlled by a given technique are noted while for the applications
factor, those specific classes of engines for which a given limited
control technique is apropos are also noted.
Table 1II-4 tabulates the various control schemes and gives their
ratings for each of the four factors.
A description of each of the emissions control techniques listed in
Table III-4 is given in Appendix C. The descriptions briefly tell what
it is, how it works, what pollutants it controls and by what physical
mechanisms, and what the problems or drawbacks with it are.
The study of the emissions performance of these concepts will be
conducted by considering the performance in terms of grouping by cate-
gory of complexity (the third factor). Thus all known data for the
performance of control schemes of category 1 (trivial complexity), say,
will be displayed, then that of category 2, etc. The displays will be
of the emissions indices (El) as a function of pressure ratio (idle for
HC and CO, rated for NOx) in accordance with the conclusions reached in
the investigation of the behavior of production engines earlier in this
section.
A summary of the emissions performance of existing engine types
utilizing the best control techniques of technology complexity cate-
gories 1-3 demonstrated in those engines is presented in Table III-5 and
in Figures 29-33. As brought out earlier in the report, concepts of
complexity categories 1 and 2 essentially control HC and CO while cate-
gory 3 technology adds control of NOx (the glaring exception being water
injection, category 2). The appearance of HC and CO control by category
3 technology in Table III-5 arose because category 1 or 2 concepts for
HC and CO control have been used in tandem with the category 3 concept
in order to effect control of all three pollutants. Figures 29-33
reflect this fact; there is no consideration of the NOx performance of
categories 1 or 2, nor is there consideration of the HC and CO perfor-
mance of category 3. Generally a number of concepts of a given category
have been examined in particular engine types but only the best appear in
the table and in the figures. The demonstration has either been engine
testing or rig testing simulating engine operation. In the figures,
trend lines have been drawn following the procedure set forth in this
part. It is worth repeating here that these trend lines do not represent
-25-
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Table III-3
Factor
Ratings &f Control Techniques
Rating
Development Status
i
IVJ
OS
i
Control Capacity
Immediately available for
production in most
engines; little or no
difficulty in the certif-
fication process.
Effective in reducing
emissions(for which the
technique is intended)to
regulation levels in most
engines, including those
for which the operating
conditions and/or the SFC
characteristic may be
deleterious to compliance.
Governing principles
thoroughly under-
stood; application
to a specific engine
requires only tuning
of the concept to
the engine consistent
with the performance
requirements of the
engine; no major
difficulties expect-
ed; little diffi-
culty in the certif-
ication process.
Effective in reducing
emissions (for which
the control is in-
tended) to the regula-
tory levels in re-
presentative, modern
engines (good SFC
characteristics and
for HC, CO, at least,
favorable operating
conditions).
Basic concept is under-
stood, but a fair
amount of work is
involved in developing
flight-worthy hard-
ware in any applica-
tion; a number of
design points are not
adequately worked out;
Certification will be
quite thorough.
Somewhat effective
in reducing emissions
[for which the control
is intended), but not
necessarily to the
regulatory levels.
Concept is understood
in general physical
and chemical terms,
but little or no
actual testing in
engine type hardware
has been done; concept
is essentially in the
research stage; a
number of basic problems
are unresolved.
Effective in reducing
emissions only in
selected engines,
either because it
relies upon conditions
unique to that engine
or because it is capable
of controlling only
the dirtiest engines
(See Application below).
-------
Table III-3 (cont.)
Category
Rating
Complexity of the
Development Process
Trivial complexity; little
or no difficulties en-
countered in development
of the concept on most
engines for which it
would be intended.
Application
Generally appropriate for
all classes of engines (Tl
- T2,* P2).
TO
-J
I
*T3 and T4 engines are
included here in the T2
group since the standards
as promulgated, are iden-
tical and the class dis-
tinction exists only for
the purpose of setting
early smoke standards.
Only some minor to mod-
erate difficulty anti-
cipated in the devel-
opment of the concept
or an engine, usually
associated with the
combustor durability
or performance; devel-
opment straightforward,
although some time is
required.
Appropriate for all
classes of engines,
but with some dif-
ficulty in develop-
ment in some engines
of one or more classes
due to the character-
istics of that class
or special characteris-
tics of those particu-
lar engines.
A number of difficulties
must be overcome in
developing flightworthy
hardware; preservation
of other design criteria
requires a compromise
with the full potential
of the emissions control
device.
Appropriate generally
only for one class
although it might be
incorporated into one
or more engines of
another class.
A number of basic
problems are unre-
solved and are ex-
pected to yield to
solution only with
difficulty; in
fact, perhaps not
all of the develop-
ment problems have
even been identi-
fied.
Appropriate only for
a specific engine
or two within a
given class.
-------
Table III-4
Ratings of Emissions Control Techniques
Control
Techniques
Fuel atomization
Idle speed
Airbleed
Air-assist (external)
Fuel-sectoring
Nozzle design
Air blast (Aeration)
Combustor length
Water injection
Rich primary
Lean primary (1)
Delayed dilution
Premix (1)
Lean Primary (2)
Fuel Staging
Premix (2)
Variable Geometry
(Air staging)
Catalysis
Development
Status
1
1
1
1
2
1
2
1
1-2
2
2
2
2
3
3
3-4
4
4+
Control
Capacity
4(HC, CO, NOx)
3(HC, CO)
3(HC, CO)
1(HC, CO)
2(HC, CO)
4(HC, CO)
1(HC, CO)
4(HC, CO)
l(NOx)
3(HC, CO)
3(HC, CO)
3(HC, CO)
2(HC, CO)
2 (NOx)
2 (NOx)
3(HC, CO, NOx)
1-2 (HC, CO, NOx)
1(HC, CO, NOx)
Category of
Complexity
1
1
1
1
1
1-2
2
2
2
2
2
2
2
2-3
2-3
3-4
3-4
4
Application*
4
1
1
1
2
4
3***
4
3(T2)
2
2
2-3****
2
2
1
2
1
* classes T2, T3, T4 are considered here as a single group
** Divided chamber fuel staging; staging by divided nozzles within a single
combustor may be appropriate for small engines.
*** Higher pressure ratio engines only
**** Application easiest in those limited cases in which combustor length is
not constrained (e.g., Allison 250)
-28-
-------
Standard
CF6-50C
CF6-6D
CFM56
JT9D-7
JT9D-70
JT10D
JT3D-7
JT8D-17
Manufacturer
GE
GE
GE/Snecma
P&WA
P&WA
P&WA
P&WA
P&WA
Size
49,900 Ibs.
38,900 Ibs.
22,000 Ibs.
46,150 Ibs.
51,150 Ibs.
^23,000 Ibs.
19,000 Ibs.
16,000 Ibs.
Table III-5
Technology Advances
Sector Burning
Double Annular Corn-
bus tor (Fuel Staging)
Compressor Bleed
Baseline**
6% Idle
Sector Burning
Modified Combustor with
Sector Burning
Air Assist
Aerating Nozzle
(Air Blast)
Hybrid Combustor
(Fuel Staging)
Air Assist
(Aerating Nozzle)
(Air Blast)
With Water Injection
Fuel Staging/Premix
Baseline**
Aerating Nozzle***
(Airblast)
Air Assist
Aerating Nozzle
(Air Blast)
Vorbix (Fuel Staging)
Complexity
Category
1
2
1
2
1
2
2
3
HC(EPAP)
0.8
0.3
0.3
1.7
1.7
1.3
0.9
0.2
1.4
0 = 2
0.7
0.3
0.2
0.2
0.7
18.0
1.2
0.8
Level Achieved
CO(EPAP)
4.3
5.3
3.0
8.0
12.8
11.1
10.9
9.7
9.4
1.3
3.3
3.9
2.3
2.3
3.5
No Data
26.2
10.5
5.0
NOx(EPAP)
3.0
7.8
4.3
7.4
4.7
3.7
4.8
4.8
4.9
4.6
3.5
5.8
4.8
1.9
3.8
5.6
7.1
7.1.
0.2
9.0
* Includes special classes T3 and T4
** New engine; baseline represents the first design against
which to compare later modificatlcms.
*** Designed for smoke standard only; no attempt to
comply with 1979 gaseous pollutant standards.
4.3
-------
Table III-5 continued
Engine
Class T2*
RB211-22B
RB211-524
Spey 511
Spey 555
Class Tl
Standard
TFE731-2
ATF3
JT15D-4
CJ610
CF700
M45H
RB401
Manufacturer
Rolls Royce
Rolls Royce
Rolls Royce
Rolls Royce
AiResearch
AiResearch
PWAC
GE
GE
Rolls Royce
Rolls Royce
Lycoming
Size
42,000 Ibs.
50,000 Ibs.
11,400 Ibs.
9,850
3,500 Ibs.
^5,000 Ibs.
2,500 Ibs.
2,950 Ibs.
4,300 Ibs.
7,600 Ibs.
5,400 Ibs.
6,500 Ibs.
Unknown
Unknown
Unknown
Unknown
Air Assist (270 psig)
Piloted Airblast
Premix/Mod 3
Baseline**
Compressor Bleed (Idle +
Approach - 10%)
Airflow Distribution,
Injector Mod
Nozzle Mod
Airflow Distribution
Nozzle Mod
Airflow Distribution
Unknown
Baseline**
Baseline**
Air Blast****
Air Blast/Combustor Mod
Flightworthy hardware
Complexity
Category
1-2
1-2
1-2
1-2
1
2
3
1
2
1
2
1
2
1-2
JT12A P6|WA 3>300 lbg
****Nozzle modification, no airflow redistribution
HC (EPAP)
0.7
0.5
5.0
6.0
1.6
0.5
0.7
0.6
5.8
5.2
0.14
2.5
1.6
0.5
0.2
3.3
2.3
1.5
2.6
Level A'chieved
CO (EPAP)
5.6
3.7
20.7
23.4
9.4
6.7
. 11.6
9.2
16.9
30.0
10.9
154; 5
135.7
94
82
5.3
10.2
23.4
11.9
11.9
17.8
No data
NOx(EPAP)
6.8
7.1
7.4
4.2
3.4
5.6
5.4.
277
4.1
4.5
4.8
2.7
2.7
2.2
2.2
3.3
3.6
3.5
4.1
3.0
3.2
-------
Table III-5 continued
Level Achieved
Engine
Class P2
Standard
TPE331-2
TPE331-3
T5321A
LTP101
PLT27
250
501D-22A
i
)-
i
PT6A-27
PT6A-41
Dart
Tyne
Manufacturer
AiResearch
AiResearch
Lycoming
Lycoming
Lycoming
Allison
Allison
PWAC
PWAC
Rolls Royce
Rolls Royce
Size
755 eshp
900 eshp
1800 shp
620 eshp
2000 shp
400 shp
4680 shp
680 shp
900 shp
2280 shp
5500 shp
Primary Orifice Only
at Idle (Injector Mod)
Primary Orifice Only
at Idle (Injector Mod)
Injector Mod
Airflow Distribution
Premix Combustor
Reverse Flow Combustor
(Airflow.Distribution)
Fuel Staging
Improved Primary Mixing/
Airflow Distribution
Airflow Distribution
Duplex nozzles
Complexity
Category
1
1
1
2
2
2
3
2
HC(EPAP)
4.9
4.3
3.6
8.3
3.6
1.2
0.3
0.6
4.8
CO(EPAP)
26.8
18.8
12.8
No Data
25.6
17.5
No Data
13.8
4.6
5.7
31.6
NOx(EPAP
. 12.9
7.7
8.9
6.1
6.2
9.3
7.3
7.2
_ .
1.8
66
16.6
95
No Data
8.3
1.4
-------
Table III-5 continued
Engine
Class APU"^
Standard
CTCP85-98
GTCP30-92
GTCP36
TSCP700
Manufacturer
AiResearch
AiResearch
AiResearch
AiResearch
Size
290 eshp
100 shp
192 eshp
910 eshp
Complexity
Category
GTCP660 AiResearch
Titan (T39) Solar
ST6 PWAC
1100 eshp
720 shp
Change in Spray Cone
Angle (Nozzle Mod)
Production Combustor
Airblast
Production Combustor
Water Injection
Production Combustor
Airflow Distribution
HC(EPAP)
0.4
0.1
0.1
0.04+
0.2
0.2
0.02
Level Achieved
CO(EPAP)
5.0
3.7
4.8
5.0+
0.8
1.6
7.9
No Data
1.0
NOx(EPAP)
3.0
6.9
3.4
5.4
6.0
3.0
4.9
5.2
+ Estimated from incomplete data.
++ A number of production engines have been quoted here because while no improvements
have been reported, those engines comply with the HC standard and perhaps the CO
standards and thus are at least partially successful.
-------
any particular fundamental constraint upon emissions performance. For
existing engines, though, the engine cycle is fixed and therefore so is
the SFC characteristic (Figures 7, 8, 9, 10, and 24). By equation 1,
then, the only remaining mechanism by which to influence the EPAP value
is the combustor emissions performance (Els). To this extend there is a
fundamental constraint on the EPAPs achievable for existing engines:
Only the combustor performance can be controlled.
a. HC arid CO Emissions
For this part concepts with technology categories 1 and 2 will be
addressed as they only control HC and CO production. First the perfor-
mance of category 1 will be examined, then that of category 2. Category
1, from Table III-3, involves only trival complexity to implement and
thus may be considered to include the ideas that are the least expensive
and most quickly implemented.
Figures 34, 35, and 36 present the HC and CO emissions performance
of the best category 1 techniques applied in current engine combustors
(rig or engine). The specific techniques employed in each engine com-
bustor are those listed in Table III-5 and in Appendix B, the data bank.
Regressions of the data are shown and represent the "average" performance
of the group; also shown are the regressions from the production engine
curves, Figures 14, 15, and 16.
It should be understood that control concepts with data points
above that "average" do not necessarily constitute inferior concepts. In
the discussion on the selection of an appropriate parameter with which
to compare emissions performance (IPR was chosen, S/V and ti are not
used), it was shown that there are other effects besides IPR which
impact on the final emissions. Some of the effects can be modified
through selective design (e.g., V changes to change ft) but only to a
point, for there are other competing design criteria (e.g. V changes
affect altitude relight). Thus, it is possible that despite all attempts
to optimize design, the combustor of a specific engine may nonetheless
exhibit relatively poor emissions characteristics because of inherent
features in the combustor that cannot be completely eliminated.
For example, the CFM56 has a higher than average El(CO) in its
category 1 application considering the IPR of the engine. The reason is
due largely to its short (low volume) combustor which leads to a fairly
short residence time (reflected partially in the idle J2 = 0.66) which in
turn causes high CO emissions because there is inadequate time for the
CO -» C02 reaction to approach equilibrium. There is little the designer
can do to increase appreciably the volume of this combustor (category 2
work) without incurring considerable problems with engine envelope (cost
and weight penalty) and liner cooling or turbine parts (durability
penalty as cooling is adversely affected by the requirement that now a
larger surface area must be cooled by the same amount of air - an increase
in the cooling air would mean less air for temperature pattern adjustment
before the turbine creating a durability problem there).
-33-
-------
Figures 37, 38, and 39 present the HC and CO emissions performance
of the best technology rating 2 technology. The regressions for that
data, the category 1 data, and the baseline data are labeled. In these
figures, the JT3D data have not been included although the low smoke
work now undergoing service evaluation constitutes category 2 work. The
reason is basically that work was not directed towards compliance with
the 1979 standards and does not necessarily reflect the best that can be
done with that combustor. Again the specific concepts which are involved
are those listed in Table III-5.
b. NOx Emissions
The discussion here shall be generally limited to category 3
concepts. The only category 2 technique that is capable of controlling
NOx is water injection which is recognized to be extremely effective.
However, as its use would normally be considered a stop-gap measure,
only a little space will be devoted here to a study of its performance;
the remainder will be devoted to category 3 technology.
Category 3 technology, and for that matter, water injection, in-
herently control mainly NOx, leaving HC and CO control to the schemes of
category 1 and 2. In fact, NOx control concepts frequently encourage a
compromise between NOx and CO emissions, letting CO increase in order to
further reduce NOx. While not ideal, this is at least acceptable so
long as CO control techniques are independently available to keep the
levels tolerable.
Figure 40 shows the NOx emissions performance of the known category
3 concept studies presented on the modified Lipfert plot. It is impor-
tant to realize that there is nothing absolute about these curves. The
existing techniques are generally quite capable of lower NOx levels
simply by retuning the air patterns or fuel flow schedules. However,
this would result in exponentially increasing CO levels which, at the
present performance (and likely the future performance) of category 1
and 2 technology, would yield unacceptably high CO levels (i.e., in
excess of the standard). Thus, the performance level of this NOx tech-
nology is based in large measure upon the CO standard and the limits of
available CO control.
The distortions of the curves from a line with a monotonically
varying radius of curvature are due to the fuel staging arrangements
wherein at approach power, in particular, there are a variety of options
available for the fuel scheduling which lead to radically different
emissions levels. At the two power limits at which optimum conditions
are better defined, it is seen that the low power NOx is comparable to
production combustors while the high power NOx is lower, a result of the
use of lean burning. At approach power, in particular, the JT9D-70 is
seen to have high NOx emissions, higher, in fact, than production engines
(Lipfert). This is a result of trying to keep the CO down and to preserve
stable operation at that power. The variability between rich and lean
operation and the stability of the latter is achieved by the fuel staging.
-34-
-------
The capability of water injection is presented in Figure 41. The
Spread reflects the fact that almost any.level of NOx is achievable
if desired short of putting the flame out altogether. The penalty paid
for high water injection rates is excessive CO production which is
demonstrated in Figure 42.
With this information of the "average" performance of the different
categories as represented by the regression lines, it is possible to
establish the EPAP trends. This serves to give an indication of the
potential of a given level of technology (complexity) to achieve compli-
ance with the standards when installed in existing engine types. This
is only an indication, however. First, a number of concepts tried on
various engines do not represent necessarily the best concepts of that
category and yet those data affect the location of the regression or
"average" line. If the few best techniques had been tried on all engines,
the overall performance levels would have been better. Second, a given
control concept, although potentially very good, may not have been fully
optimized for a particular application before being abandoned. Hence,
that too will raise the average performance level above that which is
possible.
Using the combustor performance curves on Figures 35, 36, 38, and
39, combining with the engine data on Figures 6 and 9, and using the
approximation of equation (2), one can establish the trend lines shown
in Figures 29 through 32. One could adopt the same procedure in the
derivation of a trend line for NOx (using Figures 40, 24, and 28).
However, this would not be particularly accurate in view of the depen-
dence of the El(NOx) value upon the power level as well as upon the
pressure ratio. For example, the data point obtained at the approach
power of the JT9D-7 engine (eg., PR = 10) would be totally invalid if
applied to, say, the climb power point of a low pressure ratio engine
also operating at PR - 10. The two engines, being at different power
modes, would have quite different temperature patterns due to fuel
staging differences which would result in different El(NOx) despite the
same combustor pressure. One procedure that improves upon the method is
presented in Appendix G. The resulting trend is drawn on Figure 33.
D. Newly Certified Engines
1. Introduction
In this part a review of the technology applicable to newly certi-
fied engines is presented. The analysis conducted is similar to that
conducted for the T classes of newly manufactured engines (NME) and in
fact much of the data is drawn from those figures. The major extension
of the analysis is the requirement to specify a new SFC curve. This
extension is necessary because of the very fact that newly certified
engines (NCE), which imply newly designed engines, are under considera-
tion. These engines are more capable of low emissions because of the
greater degree of freedom available (1) in the selection of the engine
-35-
-------
cycle (affecting the SFC) and (2) in the design of the combustor and
ancillary systems (permitting more advanced technology to be used,
giving lower Els). From equation (1), it is evident that this freedom
of control of both the SFC and the El offers the maximum potential for
effective control. In addition an analysis is conducted on the so-
called "prop-fan" engine concept (Reference 61 and 62) which appears to
be an attractive alternative to the turbofan engine for. fuel conserva-
tive aircraft of the future.
2. T1-T2 standards
As remarked in part C, category 1 and 2 technology offer effective
HC and CO control, but little or no NOx control, whereas category 3
technology contains concepts which are capable of achieving a fair
degree of NOx control although sometimes at the sacrifice of some HC and
CO control (requiring, therefore, category 1 and 2 techniques to be used
in conjunction so as to retain low HC, CO characteristics). Category 4
technology should be capable of optimum control of all three pollutants,
but as said at the beginning of the report, so little has been done
that accurate reporting is not possible. Furthermore, the infancy of
the development work and the long leadtime required render category 4
useless insofar as early compliance is concerned.
The category 1 and 2 technology that is available to newly manu-
factured engines (NME) for HC and CO control is already adequate to
allow future engines to comply with the presently promulgated HC and CO
standards for T2 class newly certified engines (NCE). It cannot then be
expected that there will be much further development effort. The best
demonstrated El performance for HC and CO is represented by the air-
assist (category 1) and airblast (category 2) points from Figures 35,
36, 38, and 39 and is shown again in simplified form in Figures 43 and
44.
Without the availability of category 4 technology (which should be
capable of ultra-low NOx levels) only category 3 technology remains to
achieve NOx control. Water injection, a category 2 concept, is not con-
sidered here because it is an add-on type of solution with a number of
operational drawbacks such that if an alternative is available, that
will be preferred; water injection certainly is available in any case.
Category 3 potential is known, as with categories 1 and 2, from the
development work done in pursuit of the newly manufactured engine (NME')"
standards. Unfortunately, it is also known (part C) that its performance
is inadequate to allow the large, modern T2 class engines to comply with
the existing NME standards, so that it is dubious whether it can do
better for newly certified engines (NCE) in which the NOx environment is
expected to be much worse (owing to the higher pressures, see Figure
27). On the other hand, some SFC improvement is expected so that the
net effect may not be too bad.
-36-
-------
The other major technology change affecting the levels of the NCE
standards is modification of the thermodynamic cycle to improve fuel
economy (SFC). Figure 45 shows both the rated TSFC performance and the
idle SFC performance projected for newly certified engines. The rated
TSFC curve is based upon the performance of the following hypothetical
engines:
(1) Thrust = 3500 Ib. (TFE 731 size)
Bypass ratio (3) = 4
Pressure ratio (PR) = 20
Turbine inlet temperature (TIT) = 2680°R
(2) Thrust = 47,000 Ib. (JT9D siise)
g = 8
PR = 40
TIT = 2860°R
The cycle parameters for each engine have been selected on the basis of
current trends in engine design. Actual designs in the future, however,
may differ considerably from these estimates because of the existence or
non-existence of a certain technology and because of redirected economic
priorities which may reorient the trends now perceived. Specifically,
the parameters for the small engine are based upon the following points:
(1) Pressure ratio is limited by the physical size of the high
pressure components. A very high pressure ratio implies very
small dimensions in small engines as the air is dense once
pressurized. This leads to high component inefficiencies
arising from various, mostly aerodynamic, reasons. At present,
engines of this size have pressure ratios up to 13-15.
(2) Along with the increase in pressure ratio is the concomitant
increase in the turbine inlet temperature. This occurs be-
cause of the need to maintain a proper margin between the
compressor discharge temperature (which rises with higher
pressure) and the turbine inlet temperature. An inadequate
margin results in too large an engine (insufficient work per
pound of air). At present, engines of this size (pressure
ratio ^ 15) have turbine inlet temperatures of up to 2300°R.
(3) Bypass ratio is limited by the core operating parameters
(temperature and pressure) as the fan cannot extract more
power than is being produced by the core. This limit is
not reached in general practice, nor is it here, because of
possible SFC degradation at cruise conditions and because of
the weight and cost of a very large fan. Current engines of
this size have bypass ratios of less than 3, although large Tl
type engines (^ 6000 Ibs. thrust) have bypasses up to 6.
For the large engine, the following points were considered in the selection
of the engine cycle:
-37-
-------
(1) The pressure ratio is generally made as large as is prac-
ticable as small size is not a limiting factor. Factors which
do limit the pressure ratio are the maximum engine temperature
limit and the mechanical efficiency achievable. Pressure
ratios of 30 are now in service and compressors capable of
ratios of 40 to 50 are being considered. The lower level was
selected to be conservative.
(2) The remarks made with regard to the turbine inlet temperature
of the small engine are also relevant here: A higher pressure
ratio causes a higher combustor inlet temperature. On the
other hand, the trend to very high temperatures leads to
increased wear on the hot section parts and higher maintenance
costs, as well as higher initial cost (arising from the need
for better materials and turbine cooling arrangements). At
present engines of this size have turbine inlet temperatures
up to 2800°R while operating at pressure ratios up to 30. The
effect of higher maintenance cost was judged to be the dominant
influence here instead of the engine size as shown by the
continued use of the overbuilt JT4A (despite its fuel penalty
as a pure jet) and the trend towards variable thrust take-off
(if the runway length and aircraft load permit, less than 100%
power is used at take-off so that less stress is put on the
engine). This suggests that future engines, while perhaps
built to operate at high temperature, will not generally be
operated there. For this reason, it is assumed here that
while the pressure ratio will rise some, the overall operating
temperature will largely remain the same.
(3) The increased pressure ratio permits the bypass ratio to
be increased some without degradation of the cruise SFC.
However, there are constraints. A very large bypass implies a
larger and heavier engine unless the core can be made small
enough through higher operating pressure and temperature (more
work per pound of air). In any case, an upper limit of 8-10
for the bypass ratio is generally recognized; beyond that
point, it is more economical to remove the fan shroud and
create a turboprop engine (see segment 3 of this part). As no
increase in the turbine inlet temperature (operating tempera-
ture) has been assumed, it is more reasonable here to opt for
the lower bypass, 8.
The idle SFC is determined in Figure 45 by applying to the rated SFC
curve the same ratio (at each rated thrust level) as exists between the
rated and idle SFC curves for newly manufactured engines (Figures 6 and
24). It must be recognized that these SFC levels may be a little low in
view of the relatively modest turbine inlet temperatures (TIT) involved.
While these are representative of today's engines, newly certified ones
may be able to attain higher temperatures without undue wear. The
advantage of the higher temperatures is that it permits smaller engines
to do a given amount of work (more power per mass of air) and thus saves
on weight. SFC will generally increase, however.
-38-
-------
The SFC line from Figure 45, the technology lines from Figures 43
and 44, and the approximation for EPAF given in equation (2) allow
calculation of the HC and CO EPAPs expected to be realized for newly
certified engines. These EPAPs are shown on Figures 46 and 47. Implicit
in this calculation is the assumption that the idle pressure ratios of
the two sample NCEs are the same fraction of the rated pressure ratios
as is found in existing engines (6% idle assumed). This assumption
leads to idle pressure ratios of 3.0 and 6.0 respectively for the two
hypothetical engines above. It should be cautioned, however, that this
EPAP computation procedure involves some extrapolation for the technology
lines of Figures 43 and 44 as they are based upon data only up to IPR -
4.0 whereas the larger hypothetical engine has an IPR of 6.0.
As indicated earlier, the control of NOx emissions is expected to
be achieved through use of the category 3 technology. The NOx control
performance of these techniques (as typified here by the NASA/GE double
annular combustor) and for that matter, any combustor, is sensitive to
the operating pressure and temperature. As the combustors of new engines
will be run at considerably different operating conditions from that of
the CF6-50C for which the double annular combustor performance is known,
it is necessary to adjust the El data to that which would be expected in
the hypothetical newly certified engines. This could be done by referral
to a modified Lipfert type chart wherein category 3 technology data have
been correlated against P- (combustor pressure). However, the concern
about the validity of sucn a simple correlation which was discussed in
parts B and C is equally valid here, and perhaps more so, in light of
the extrapolation of conditions involved. Hence, the NOx correction
formula from the NASA ECCP (reference 59) will be utilized although even
its application involves extrapolation. Its primary advantage is that
it represents a semi-empirical multivariate correlation taking into
account several physical mechanisms of NOx production. Appendix E
explains the calculation procedure.
Inasmuch as the demonstrated technology upon which this projection
is based comes from the 50,000 Ib. thrust class CF6-50C engines, an im-
portant assumption in this analysis is that some sort of advanced combus-
tion system will be available to the small Tl class engines that offers
low NOx performance comparable to the double annular combustor. Such an
advanced combustor might evolve from the piloted premix/prevaporization
system being investigated for Tl class engines in NASA's Pollution
Reduction Technology Program (Reference 34). It is necessary to make
this assumption because the double annular system is likely to have a
very high surface to volume (S/V) ratio in small sizes (which would
probably destroy the HC and CO gains made through application of various
control schemes - see Figure 17) and is thus probably not well suited to
application in small engines. Application of the CF6-50C double annular
emissions data as corrected and combined with the modal fuel flow rates
leads to the NCE technology curve shown on Figure 48 (computation in
Appendix E).
-39-
-------
In summary, it is evident that the NCEs can be very clean with
respect to HC and CO emissions and reasonably clean in NOx emissions
although apparently the presently promulgated number (EPAP = 3.0) cannot
be met with the category 3 technology demonstrated to date. If the NCE
cycles do have higher turbine inlet temperatures than postulated, the
projected NOx emissions will definitely rise because both the SFC and
the El levels will increase.
3. P2 Standards
There has been considerable discussion recently (References 61, 62,
63, 64) that suggests that turboprop propulsion may again find utility
in large, long haul air transports because of the excellent fuel economy
afforded. Such fuel efficient designs are becoming more attractive
despite drawbacks (i.e., propeller maintenance) because of the ever ,
rising cost of fuel.
A commonly discussed concept is the "propfan" engine designed for
cruise speeds approaching that of current subsonic jet transports (Mach
0.78 vs. 0.82). This engine represents the transition between a true
fan engine (with numerous shrouded blades) and a true propeller engine
(with a few unshrouded blades). The choice of one system or the other
is basically a choice of the optimum thermodynamic cycle which in turn
determines the SFC. The optimum cycle is very sensitive to both takeoff
power requirements and the selected speed and thrust requirements at
cruise. The trend for jets over the last two decades has been towards
larger and larger bypass ratios (initially bypasses of 0, now typically
5-6 and 8-10 proposed) in order to improve the thermodynamic cycle and
hence the SFC. The propfan is merely a continuation of this trend
(bypass, 12-25) for which the bypass has reached the point that the
weight of the surrounding shroud is excessive and hence has been removed,
producing instead a low bypass, highly loaded propeller for high subsonic
speeds (a typical low cruise speed propeller is lightly loaded in terms
of work done per unit of air and has a bypass of 50 to 100).
In view of the possible development and extensive deployment of
this type of turboprop, it appears possible that the EPA may find it
necessary to take measures to ensure that the emissions technology of
such an engine be compatible with competing engines (principally newly
certified T2 class engines).
The technology that would be available to these engines is the same
as that which will be available to newly certified T2 class engines (as
summarized in Figures 43 and 44 for HC and CO, and in reference 59 for
NOx). In fact, it is conceivable that the same basic gas generator may
be used in both applications with either a fan or a propeller/gear
reduction apparatus attached. The emissions performance of the combus-
tor, however, will differ between applications in the two classes because
of different operating conditions in each mode in the two cases, primarily
at idle wherein the turboprop is likely to run at a higher speed and
hence possess lower emissions of HC and CO, but higher emissions of NOx,
from the combustor than it would experience in a turbofan application.
-40-
-------
To project the EPAP values for the advanced large turboprop (propfan)
it is necessary to estimate the power specific fuel consumption (PSFC)
and the combustor operating conditions. The former is necessary because
equation (1) requires SFC values for calculation of EPAP.* The latter
is necessary in order to project the El values for the pollutants when
the combustor operating environment differs from that for which the
values are currently known or estimated.
To estimate the SFC, it is necessary to know the thermodynamic
cycle of the engine. Two cycles were postulated for this, one represent-
ing a smaller engine, the other a large engine. The cycles, given below
in Table III-6, are compatible with those selected to represent newly
certified T2 class engines.
Table III-6
Power 5,000 HP 50,000 HP
Bypass Ratio* (3) 12,25 12,25
Pressure Ratio 20.0 40.0
Turbine Inlet 2680°R 2860°F
Temperature
* This is a measure of the propeller size and as such is char-
acteristic of the conversion of power to thrust. This is, there-
fore, not required to estimate PSFC, but is useful in establishing
core size (H ) comparisons between P2 and T2 engines of the same
thrust. alr
From these points, the PSFC curve shown on Figure 49 established. That
curve is essentially independent of the particular bypass ratio (propeller
size) selected as mentioned in the footnote to the table.
The engine sizes were selected by the following consideration. The
Tl class (general aviation) and the T2 class (commercial) are separated
at the 8000 pound thrust mark. With a few exceptions (eg., the 6,500
Ib.-thrust M45H in commercial service and the 11,400 Ib.-thrust Spey Mk
511 on a business jet), this offers a clean distinction. Depending upon
the propeller selected (corresponding to a bypass of 12 to 25), 8000
pounds of thrust corresponds to 4100 horsepower on the low end and 5800
horsepower on the high end. It is reasonable then to select 5000 horse-
power as the lower limit of the P2 class which may be utilized for
advanced commercial application. For comparison, the largest P2 class
engine now in use in the U.S. is the Allison 501 with a 4700 horsepower
rating.
*Note that for the P2 class, the power SFC in units of pounds of
fuel per hour horsepower is selected rather than the thrust SFC
in units of pounds of fuel per hour per pound of thrust as is the
case for the T1-T2 classes; this reflects the difference in the
definitions of EPAP between the classes specified in the regulations.
-41-
-------
Utilizing the combustor technology information from Figures 43 and
44 and in reference 59 along with the fuel consumption performance in
Figure 49, technology lines for HC, CO, and NOx were made. The compu-
tation is presented in Appendix F and the results shown in Figures 50,
51, and 52. The calculation for the HC technology line resulted in very
low EPAP values similar to those recorded by the technology line for T2
class newly certified engines (Figure 53). That is consistent with the
fact that (1) for a propfan type of propeller, the thrust to power ratio
at the rated condition is of the order one (eg., EPA estimates for a $
= 12 propeller the thrust to power ratio would be about 1.4) and (2) the
two classes of engines are utilizing the same combustor technology,
albeit under slightly different operating conditions at idle. The CO
calculation shows a fairly level curve (EPAP goes from 2.9 to 3.6 over a
range of 45,000 horsepower). NOx increases for the larger engines are
due to the increasingly severe combustor environment which outweighs the
meager SFC improvements seen in the larger engines.
-42-
-------
a
H
O
100
90
80
70
60
50
40
30
20
HC EPAP AS A FUNCTION OF RATED THRUST
PRODUCTION ENGINES
10
9
8
7
6
5
4
1
.9
.8
.7
.6
.5
.4
.3
.2
10,000
20,000 30,000 40,000
Rated Thrust ~ Ibs.
Figure 3
50,000 60,000
-43-
-------
100
CO EPAP AS A FUNCTION OF RATED THRUST
PRODUCTION ENGINES
w
o
u
PRODUCTION TECHNOLOGY
Idle converted to 6%
10,000
20,000 30,000 40,000 50,000
Rated Thrust Ibs.
Figure 4
-44-
-------
NOx EPAP AS A FUNCTION OF RATED THRUST
PRODUCTION ENGINES
a,
w
X 3
o
2
PRODUCTION TECHNOLOGY
* Idle converted to 6%
10,000
20,000 30,000 40,000
Rated Thrust Ibs.
50,000 60,000
Figure 5
-45-
-------
IDLE TSFC AS A FUNCTION OF RATED THRUST
PRODUCTION ENGINES
* Idle converted to 6%
0 10,000 20,000 30,000 40,000 50,000 60,000
Rated Thrust Ibs
Figure 6
-46-
-------
IDLE SFC AS A FUNCTION OF BYPASS RATIO
2.0
1.5
n
I
1.0
u
Q
0.5
^2.5)
X
* Idle converted to 6%
Estimated
(pressure ratio)
0(2.2)
Idle Pressure Ratio = 2.0
234
Bypass Ratio
Figure 7
-47-
-------
BYPASS RATIO AS A FUNCTION OF RATED THRUST
.2 4
-p
(D
to
n
0 10,000 20,000 30,000 40,000 50,000 60,000
Rated Thrust Ibs
Figure 8
-48-
-------
IDLE PRESSURE RATIO AS A FUNCTION OF RATED THRUST
O
H
4J
ti
0)
CO
CQ
0)
M
0)
rH
o
*0
*Idle converted to 6%
10,000 20,000 30,000 40,000 50,000 60,000
Rated Thrust Ibs.
Figure 9
-49-
-------
RATED TSFC AS A FUNCTION OF BYPASS RATIO
1.0
0.75
M
X!
M-l
XI
fe 0.5
CO
E-i
d
0)
-p
5
0.25
PR =
TIT = 2400
PR = 10
TIT = 2300°R
PR = 30
TIT = 2860°R
2 3 4
Bypass Ratio
Figure 10
-50-
-------
BYPASS RATIO AS A FUNCTION OF
RATED PRESSURE RATIO
O
H
4J
(0
ti
0
O
5 10 15 20 25
Rated Pressure Ratio
30
Figure 11
-51-
-------
HC EMISSION INDEX AT IDLE AS A FUNCTION
OF RATED THRUST
PRODUCTION ENGINES
100
90
80
70
60
50
§ 40
«H
XI
o
o
o
I
-------
-------
COMBUSTION INEFFICIENCY AS A FUNCTION OF IDLE PRESSURE RATIO
PRODUCTION ENGINES
O
c
W
4-1
0)
C
H
O
H
4J
CQ
3
O
O
,10
.09
.08
,07
06
05
.04
03
02
01
O
A
2 3
Idle Pressure Ratio
Figure 14
-54-
-------
100
90
80
70
60
50
40
30
HC EMISSIONS INDEX AT IDLE AS A FUNCTION OF
IDLE PRESSURE RATIO
PRODUCTION ENGINES
0)
3
4-4
o
o
o
20
0)
H
T)
-U
(0
X
0)
CO
C
O
H
(0
CO
H
U
a
10
9
8
7
6
5
4
2 3
Idle Pressure Ratio
Figure 15
-55-
-------
CO EMISSIONS INDEX AT IDLE AS A FUNCTION OF
IDLE PRESSURE RATIO
PRODUCTION ENGINES
1000
900
800
700
600
500
400
i
i 300
o 200
o
o
O
p
flj
X
0)
"0
c
W
C
O
H
U)
in
H
O
U
100
90
80
70
60
50
40
30
20
10
2 3
Idle Pressure Ratio
Figure 16
-56-
-------
COMBUSTION INEFFICIENCY AS A FUNCTION OF
SURFACE TO VOLUME RATIO
PRODUCTION ENGINES
.10
.09
.08
.07
,06
.05
.04
.03
.02
o
-------
HC EMISSIONS INDEX AT IDLE AS A FUNCTION OF
SURFACE TO VOLUME RATIO
PRODUCTION ENGINES
on
on
70
gn
SO
40
30
on
. H
3
^ 10
-Q £3 Q
pHH g
g 7
S 6
0)
H ..
D 4
H
* 3
X
0)
£ 0
Missions ]
H f
.9
7
g
.5
.4
3
.2
.1
e._
^
M^^^M*^^"->^M^HMI
D
^ ^-~
_
l-j
A
"
^
A
IV
^
7 8 9
Surface to Volume Ratio
Figure 18
10
11
12
-58-
-------
1000
900
800
700
600
500
400
300
CO EMISSIONS INDEX AT IDLE AS A FUNCTION OF
SURFACE TO VOLUME RATIO
PRODUCTION ENGINES
200
100
90
80
70
60
50
40
30
o
X
0)
o
M
o
H
(0
M
rH
20
10
9
8
7
6
8 9
Surface to Volume Ratio
Figure 19
-59-
ft
10
-1
11
-------
IDLE PRESSURE RATIO AS A FUNCTION OF
SURFACE TO VOLUME RATIO
O
H
-P
id
0)
M
3
w
(0
0)
789
Surface to Volume Ratio
Figure 20
10
-60-
-------
COMBUSTION INEFFICIENCY AS A FUNCTION OF
AIR LOADING PARAMETER
PRODUCTION ENGINES
(.
c
>1
O
c
0) in
*^ A O
O 09
H . 08
£ -07
a! 06
. . ft c
H . U J
c nfl
o '°4
^
4J 03
w >UJ
3
ja
6 n9
0 . 02
U
. 010
009
008
007
006
f\ f\ C
n^A
on^
on?
.001
^
o
^
. A
W
'\
^*
£
^
A
^
P
^
^
^
NT*
ri
^x*
^
^*^
*
*.
^^
^^
.01
.10
Air Loading Parameter
10
Ib
ft3 - sec
Figure 21
-61-
-------
o
0)
U)
M
(U
P
H
T3
a
o
n
-H
10
9
8
7
6
1
.9
.8
.7
.6
.5
.4
.2
AIR LOADING PARAMETER AS A FUNCTION OF
IDLE PRESSURE RATIO
PRODUCTION ENGINES
O
D
Idle Pressure Ratio
Figure 22
-62-
-------
10
9
8
7
6
AIR LOADING PARAMETER AS A FUNCTION OF
SURFACE TO VOLUME RATIO
PRODUCTION ENGINES
o
-------
RATED TSFC AS A FUNCTION OF RATED THRUST
PRODUCTION ENGINES
I
J-l
I
u
73
0)
-P
0 10,000 20,000 30,000 40,000 50,000 60,000
Rated Thrust Ibs.
Figure 24
-64-.
-------
100
90
80
70
60
50
40
30
3
IM
XI
o
o
o
X
0)
13
0)
O
-^
in
to
H
X
i
20
10
9
8
7
NOx EMISSIONS INDEX AS A FUNCTION OF RATED THRUST
PRODUCTIONS ENGINES
10,000
20,000
30,000 40,000
50,000
60,000
Rated Thrust Ibs
Figure 25
-65-
-------
NOx EMISSIONS INDEX AS A FUNCTION OF COMPRESSOR DISCHARGE TEMPERATURE
100
90
80
70
50
40
30
20
o
o
o
o
o
o
H
co
CO
X
o
2
10
9
PRODUCTION ENGINES
Lipfert A)
9
T2 Class Engines
T3 Class Engines
T4 Class Engines
Tl Class Engines
APU Engines
P2 Class Engines
600 800 1000 1200 1400
Compressor Discharge Temperature -"R
Figure 26
1600
-66-
-------
NOx EMISSIONS INDEX AS A FUNCTION OF
PRESSURE RATIO
PRODUCTION ENGINES
DCLASS Tl ENGINES
OCLASS T2 ENGINES
3CLASS T3 ENGINE
CLASS T4 ENGINE
10 15 20
Pressure Ratio
Figure 27
-67-
-------
RATED COMPRESSOR PRESSURE RATIO AS A FUNCTION OF
RATED THRUST
O
H
-P
(0
0)
w
w
O
W
CO
0)
a
§
U
a
0)
P
30
28
26
24
22
20
18
16
14
12
10
8
6
4
O
10,000 20,000 30,000
Rated Thrust* Ibs
Figure 28
40,000
50,000 60,000
-68-
-------
g
w
u
10
9
8
7
6
5
1
.9
.8
.7
.6
.5
.4
.3
.2
HC EPAP AS A FUNCTION OF RATED THRUST
TECHNOLOGY CATEGORY 1
NME STANDARD
M.
o*
Idle converted to 6%
0 10,000 20,000 30,000 40,000 50,000 60,000
Rated Thrust Ibs
Figure 29
-69-
-------
100
90
80
70
60
50
40
30
20
10
9
Si 8
cm
« 7
o
U x-
CO EPAP AS A FUNCTION OF RATED THRUST
TECHNOLOGY CATEGORY 1
NME STANDARD
0*
if Idle converted to 6%
10,000 20,000 30,000 40,000 50,000 60,000
Rated Thrust*lbs
Figure 30 . .
-70-
-------
SI
a
100
90
80
70
60
50
40
30
20
10
9
8
7
6
5
4
1
.9
.8
.7
.6
.5
.4
.3
.2
HC EPAP AS A FUNCTION OF RATED THRUST
TECHNOLOGY CATEGORY 2
if Idle converted to 6%
-NME STANDARD
o1
10,000 20,000 30,000 40,000
Rated Thrust ~ Ibs
50,000 60,000
Figure 31
-71-
-------
100
CO EPAP AS A FUNCTION OF RATED THRUST
TECHNOLOGY CATEGORY 2
wldle converted to 6%
10,000 20,000
30,000 40,000
50,000 60,000
Rated Thrust- Ibs
Figure 32
-72-
-------
a
04
w
X
§
10
9
8
7
6
5
4
NOx EPAP AS A FUNCTION OF RATED THRUST
TECHNOLOGY CATEGORY 3
Category 3
NME STANDARD
o
10,000 20,000 30,000 40,000
Rated Thrust Ibs
50,000 60,000
Figure 33
-73-
-------
COMBUSTION INEFFICIENCY AS A FUNCTION OF IDLE PRESSURE RATIO
TECHNOLOGY CATEGORY 1
001
2 3
Idle Pressure Ratio
Figure 34
-74-
-------
100
90
80
70
60
50
40
30
20
0)
3
o
o
o
I
a)
l-l
a
0)
a
e
U)
o
H
(0
01
H
O
10
9
8
7
6
1
.9
.8
.7
.6
.5
.4
.3
.2
HC EMISSIONS INDEX AT IDLE AS FUNCTION OF
IDLE PRESSURE RATIO
TECHNOLOGY CATEGORY 1
Production
Category 1-
\
2 3
Idle Pressure Ratio
Figure 35
-75-
-------
CO EMISSIONS INDEX AT IDLE AS A FUNCTION OF
IDLE PRESSURE RATIO
TECHNOLOGY CATEGORY 1
-------
001
COMBUSTION INEFFICIENCY AS A FUNCTION OF
IDLE PRESSURE RATIO
TECHNOLOGY CATEGORY 2
2 3
Idle Pressure Ratio
Figure 37
-77-
-------
100
90
80
70
60
50
40
30
20
HC EMISSIONS INDEX AT IDLE AS A FUNCTION OF
IDLE PRESSURE RATIO
TECHNOLOGY CATEGORY 2
0)
3
o
o
o
X
0)
CO
O
-rl
(0
(0
H
W
U
K
10
9
8
7
6
5
4
1
.9
.8
.7
.6
.5
.4
.3
.2
Production'
/Category
Category 2
O
1
Idle Pressure Ratio
Figure 38
-78-
-------
1000
900
800
700
600
500
400
300
200
« ioo
* 8901
a 70
o 60
o
2 50
I
d>
4J
n)
0)
T3
H
CO
§
H
(0
co
-H
8
40
H 30
20
10
9
8
7
6
5
CO EMISSIONS INDEX AT IDLE AS A FUNCTION OF
IDLE PRESSURE RATIO
TECHNOLOGY CATEGORY 2
2 3
Idle Pressure Ratio
Figure 39
-79-
-------
100
o
o
o
X
-------
100
90
80
70
60
50
40
30
20
o
o
o
X
0)
0
en
q
o
H
to
Ul
X
o
10
9
8
7
NOX EMISSIONS INDEX AS A FUNCTION OF
PRESSURE RATIO
Production Data (Lipfert)
z
z
z
With water injection
at high power
10 15 20
Pressure Ratio
Figure 41
25
30
35
-81-
-------
JT9D CLIMB POWER
0)
3
U-l
O
o
O
X
0)
a
H
(0
c
o
H
(0
(0
H
0.5 1.0
Water to Fuel Ratio
1.5
Figure 42
-82-
-------
HC EMISSIONS INDEX AT IDLE POWER AS A FUNCTION OF
IDLE PRESSURE RATIO
AIR ASSIST - AIR BLAST TECHNOLOGY
HC Emissions Index at Idle Power Ib
1000 Ib fuel
........ ,_
M NJ U> ^ Ul W -J 00 VO M K» (jj .b Ul en -J CO VO C
" !
Q
e
O
A
Idle Pressure Ratio
Figure 43
-83-
-------
100
90
80
70
60
50
40
30
<3 20
o
o
o
t
0)
iH
-O
H
X
0)
c
H
to
O
H
10
to
H
O
U
10
9
8
7
CO EMISSIONS INDEX AT IDLE AS A FUNCTION OF
IDLE PRESSURE RATIO
AIR ASSIST - AIR BLAST TECHNOLOGY
2 3
Idle Pressure Ratio
Figure 44
-84-
-------
NCE TSFC AS A FUNCTION OF RATED THRUST
.7
.6
M
f'5
CO
EH
\
NCE IDLE TSFC
NCE RATED TSFC
10,000 20,000 30,000 40,000
Rated ThrustIbs.
Figure 45
50,000 60,000
-85-
-------
1.0
.9
.8
.7
.6
Si
&4
W
u
33
.3
.2
.1
.09
.08
.07
.06
.05
.04
.03
.02
.01
HC EPAP AS A FUNCTION OF RATED THRUST
NEWLY CERTIFIED ENGINES
NCE STANDARD
1
10,000 20,000 30,000 40,000
Rated Thrust Ibs
Figure 46
-86-
50,000 60,000
-------
5!
ix
w
o
u
10
9
8
7
6
1
,9
,8
.7
6
,5
.2
CO EPAP AS A FUNCTION OF RATED THRUST
NEWLY CERTIFIED ENGINES
NCE STANDARD
NCE TECHNOLOGY
1
10,000 20,000 30,000 40,000
Rated Thrust - Ibs
Figure 47
-87-
50,000
60,000
-------
04
w
X
o
2
PROJECTED NOx EPAP AS A FUNCTION
OF RATED THRUST
PRODUCTION TECHNOLOGY
NCE TECHNOLOGY-
NCE STANDARD
10,000 20,000 30,000 40,000 50,000
Rated Thrust - Ibs
Figure 48
-88-
-------
PSFC AS A FUNCTION OF HORSEPOWER
NEWLY CERTIFIED CLASS P2 ENGINES
.5
.4
t
u
en
.3
.2
.1
10,000 20,000 30,000 40,000
Horsepower
Figure 49
50,000 60,000
-89-
-------
Si
o<
w
u
SB
10
9
8
7
6
5
.3
.2
.1
HC EPAP AS A FUNCTION OF HORSEPOWER
NEWLY CERTIFIED CLASS P2 ENGINES
0 10,000 20,000 30,000 40,000 50,000 60,000
Horsepower
Figure 50
-90-
-------
CO EPAP AS A FUNCTION OF HORSEPOWER
NEWLY CERTIFIED CLASS P2 ENGINES
a
&
w
o
u
10
9
8
7
6
5
10,000 20,000 30,000 40,000
Horsepower
50,000 60,000
Figure 51
-91-
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NOx EPAP AS A FUNCTION OF HORSEPOWER
NEWLY CERTIFIED CLASS P2 ENGINES
20
10
9
8
7
6
5
O4
w ,
X
i
10,000 20,000 30,000 40,000 50,000 60,000
Horsepower
Figure 52
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Section IV. Leadtime Requirements
A. Introductory Remarks
Since the regulations were first promulgated three years ago
(reference 66) the engine manufacturers and NASA have been largely
directing their activities and resources towards compliance with the
1979 (newly manufactured engine) standards for all four of the regulated
pollutants, HC, CO, NOx, and smoke. Total success has been limited in
view of the now appreciated need for some type of staged combustion
system in order to substantially reduce NOx emissions without unduly
compromising the other combustor design criteria. Staged combustion
systems (eg., the NASA Experimental Clean Combustor Program double
annular combustor applied to the General Electric CF6-50) are complex
systems involving unproven design concepts (technology complexity cate-
gory 3). Application in commerical use also requires economically sound
design as well as safe design, thereby burdening the designer with
stringent durability requirements as well as performance requirements,
all of which must be met by the new concept. The manufacturers have
also explored modifications to conventional designs which offer far less
development risk (technology complexity categories 1 and 2), but also
less capacity to control all four pollutants. Specifically, the control
of NOx is largely lost although there are exceptions wherein some im-
provement was achieved nonetheless (but not enough for compliance).
This section attempts to relate this complicated development situation
to a series of time schedules, each associated with one of the established
technology complexity categories of the control concepts which may be
used for compliance.
In Section III, all known emissions control concepts were cate-
gorized according to their complexity of development. The categories
were:
1. Little or no difficulty
2. Only minor difficulty
3. A number of difficulties to be overcome, usually
with a compromise with other design criteria
4. Basic problems not resolved and are expected to be
solved only with great difficulty
The fourth category is quite esoteric (e.g., catalysis) and conse-
quently so little work has been done on methods within this group that
it is not a viable alternative for compliance with any of the standards
that are hoped to be implemented in the near future. As a result of
this lack of sufficient time and also of the sheer lack of information,
there will be no further consideration of this category here.
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Each of the categories, because of their different levels of dif-
ficulty, have associated with-them a typical or average compliance
schedule which would apply roughly to the development of any of the
control concepts in that rating. Naturally, the more complex the tech-
nology level and the control schemes within it, the longer is this
typical compliance schedule. In the work that follows, these compliance
schedules will be developed. However, in ascertaining a compliance date
which would be reasonable for the utilization of a given rating, it is
necessary to examine more than the typical schedule that is prescribed
for that technology level. First, it is necessary to identify where on
that schedule the different engines presently lie. Second, it is neces-
sary to decide if the concept being developed for each engine is adequate
(being complex does not guarantee it is effective) and if not, it is
necessary to select an adequate concept and to re-evaluate the position
of the engine on the schedule accordingly. Third, it may be necessary
to revise the schedule to reflect peculiarities of a given concept or
limitations of a given manufacturer which may radically alter it (recall
that the schedule is only a typical or average schedule associated with
that rating).
Two preliminary points should be made. First, if a set of standards
are proposed based upon the utilization of a given level of complexity
(e.g., .category 1), it must be recognized that there may be a few engines
which, for a variety of reasons, cannot comply with these standards
without utilizing more complex types of control (e.g., category 2).
These engines necessarily would be forced to follow the schedule of the
more complex rating. In a similar vein are those engines which simply
cannot comply with any set of standards employing any demonstrated
techniques within each rating of technology (except perhaps for category
4 which is not available). Such engines would require an indefinite
research time to continue to explore new and heretofore undemonstrated
techniques. In this case, a compliance schedule cannot be readily
identified.
The second point is that any schedule can be identified as either a
normal or an accelerated schedule. Both types will be developed here
but in view of the delays beyond the January 1, 1979 deadline that the
manufacturers have claimed, the emphasis in Section V will be on the
accelerated schedules.
B. Stages of a Technology Development Schedule
There are nine basic steps that occur in the incorporation of a
technical concept into production. They are:
1. Concept verification
2. Combustor demonstration
3. Engine demonstration
4. Production design and procurement
5. Engine testing
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6. Flight testing
7. Qualification
8. Service evaluation
9. Production preparation
C. Technology Category 2 Schedule (Normal)
The presentation below will explain the work tasks involved in each
of these steps. Representative times will be given for a normal schedule
-for a modest technological advance (category 2). Modifications to this
basic schedule for an accelerated pace, a minor technological advance
(category 1), and a major technological advance (category 3) will then
be discussed.
Details of the nine basic steps are as follows:
1. Concept verification. This task is considered a separate step
only if the involved manufacturer is taking demonstrated technology from
another source to incorporate into his engine without prior research
work of his'own on that approach. The purpose of this activity is to
verify that the concept is functional in the combustor and has the
potential for controlling emissions to the necessary level. This is not
a large task; only sufficient testing is necessary here to evaluate the
worth of the concept, whereas in the next step the concept must be
optimized. Six months are sufficient for the analysis of existing data,
planning, hardware production, and limited testing that is necessary.
A manufacturer who has undergone the exploratory development of a
given concept prior to the go-ahead decision need not undergo this step
at all as the necessary verification was performed in the course of the
research effort.
2. Combustor development and demonstration in rig test. At the
start of this task, a number of low emissions technology concepts have
been applied to a combustor rig and their potential evaluated. Generally
such testing does not yield configurations that are optimized with
respect to emissions and engine performance. In fact, such configura-
tions are often totally unflightworthy as well as perhaps marginal in
emissions performance. Once one or two prospective candidates have been
selected for further development, this task begins.
The chosen concept or concepts must be refined in the rig sufficient-
ly to meet the pertinent standards when applied to the actual engine.
Rig testing also permits extensive optimization of the system for safety
and reliability. It further permits simulated altitude performance
testing (including relight) which otherwise would be very costly in
engine tests in altitude chambers, if possible at all.
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Phase II of the NASA programs (ECCP and PRTP) involves the same
sort of effort and in those programs the small engines (PRTP) are allowed
9 months for this work and the large engines (ECCP), 15 months. However,
in the latter case especially, the technology involved is much more
exotic (category 3) and additional time is obviously necessary. It is,
therefore, reasonable to expect the requisite rig testing for category
2 technology to be accomplished in 9 months.
3. Engine Demonstration. Engine demonstration is necessary,
first, to verify that the combustor, which was optimized in the rig
tests of task 2, operates satisfactorily in the full engine, and second,
to redesign any subsystem as necessary (e.g., the fuel control system)
to achieve that satisfactory performance. This is essentially the
"proof of concept" test. Satisfactory performance here means (1) accepta-
ble steady state performance as evidenced by specific fuel consumption
within production specification and acceptable turbine inlet temperature
profile and (2) acceptable transient performance as evidenced by a
acceleration/deceleration times within FAA requirements with sufficient
surge margin. The time allotted to this task is 9 months. For compari-
son, the NASA programs, ECCP and PRTP, allow 15 and 12 months respectively.
(Phase III). However, they involve, in their entirety or in part, more
complex control schemes than are necessary for category 2 sophistication.
4. Production Design and Procurement. This task involves the
design and manufacture (in limited numbers) of all the engine parts
modified by the low emissions system and the associated tooling. Such
engine parts may include, in addition to the combustor liner itself, the
inner and outer casing, diffuser, struts, fuel control, and nozzles.
Manufacturing includes the necessary tooling as well as the limited
production to supply engines for static and flight testing, qualification,
and service evaluation. Proving of the tooling can be accomplished in
parallel with the subsequent steps prior to commitment to production.
As reported in references 44, 48 and 49 this exercise should take a
little over one year, perhaps a quarter of which (design phase) can
easily overlap the engine demonstration period as the production configuration
is finally identified.
5. Engine Testing. This activity involves endurance and cyclic
testing necessary first to generate an adequate safety record for later
flight testing and engine qualification and second to develop a main-
tenance and reliability record for service. Typically about one year is
necessary before sufficient time is accumulated to begin flight tests as
reported by General Electric and Rolls Royce in references 8, 48,and 45.
This type of testing, however, usually continues throughout the production
life of the engine.
6. Flight Testing. This is done primarily to investigate engine
performance at altitude. It is usually done on a corporate owned experi-
mental aircraft or an available military aircraft neither of which are
subject to the FAA requirements of certification. Included in the
engine performance criteria are thrust, specific fuel consumption,
relight, and transient behavior. Environmental factors such as icing,
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etc., may be investigated as necessary. The T2 class engine manufac-
turers generally allocate about three to six months to flight testing
prior to the start of qualification (references 8 and 48) while the Tl
and P2 class engine manufacturers apparently often get by with less,
perhaps three months (references 4 and 33).
7. Qualification. This step involves obtaining the necessary type
certificate, supplemental type certificate, or engineering approval for
the low emissions engine. The amount of effort involved here is largely
dependent upon the type of certification necessary. In any case, FAA
certification involves mostly paperwork as a large part of the necessary
testing has already been done. Following the recommendation of several
of the manufacturers (references 8, 14, 33, 45, and 48), three to six
months are delegated to this task.
8. Service Evaluation. Once an engine has the necessary FAA
certification or approval, it is, strictly speaking, available for
public use. Nonetheless, another stage of testing, that of service
evaluation, has developed as a matter of industry policy prior to full
production. The procedure in service evaluation is to have a limited
number of engines installed on fleet aircraft (one per plane) for a long
enough period of time to judge their performance, reliability, and
maintainability in actual use. While reliability may be considered a
safety issue (the FAA considers service evaluation a vital supplement to
its own required testing for certification), it is intended primarily to
prove the economics of the engine in service. The length of time of the
service evaluation depends upon the rate at which flight time and landing-
takeoff cycles are accumulated in service and the extent to which the
new system differs from the old (an exotic system will be scrutinized
more thoroughly for reliability). A system which has had a history of
difficulty in development will also command a longer service evaluation.
At least one manufacturer had earlier indicated that a one year
service evaluation would be adequate for its engines (reference 4).
This is a fairly short time, giving a high time of perhaps 3500 hours
and 1400 cycles for commercial engines and somewhat less for general
aviation engines. However, this is supported by reference 26. It
should be pointed out, on the other hand, that some service evaluations
take longer; for instance, the JT8D and JT3D smoke retrofit programs
service evaluations took 1 1/2 to 2 years to complete. However, both
represented attempts to install advanced technology in quite mature
engines (can and can-annular combustors) and both experienced a history
of development problems. A similar situation can be expected to occur
with the Allison 501D (P2 class), another very mature engine.
There are, however, two more points that must be weighed. First,
the large commercial engines classes, while of fairly recent origin, are
accumulating flight time very rapidly and by the time the new combustors
are ready for production, these engines will themselves be "mature" (10-
12 years old) so that the airlines will be expecting a high level of
durability. Second, as these large engines will dominate the fleet and
all may eventually be affected by the standards (because of the proposed
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retrofit requirement), it follows that any failure to perform satisfac-
torily may have a considerable economic impact on the airlines. For
these reasons, the airlines contend that an adequate service evaluation
should not be compromised. Any lost time would eventually be compensated
for by the proposed retrofit program. Following recent industry recommenda-
tions (references 28, 40, 44, 45, 48, and 49), two years are allotted
for commercial engines for the service evaluation. The airlines involved
in the service evaluation should be encouraged to see that the two year
period is used to the maximum advantage. Failure to put the greatest
possible time on the engines in service evaluation may lead to the
acceptance of a system with hidden flaws which would eventually prove
unreasonably costly to the airlines.
For general aviation a problem arises in that the annual usage rate
is generally far below that of the commercial engines. This means that
in a given period, considerably less time can be put on the engines.
This is compensated for by the less stringent durability requirements of
private operators. A one year service evaluation is generally considered
adequate for the Tl and P2 classes by the industry (references 4 and 32)
with the exception of General Motors which produces a very mature P2
class engine (the Allison 501D) achieving a very high durability perfor-
mance. This single engine should be treated as a special case, if
necessary.
9. Production Preparation. The manufacturers (eg. reference 45)
have called for a one year lag between the end of the service evaluation
and the actual production on the basis that they do not wish to commit
to final tooling, etc., until the final validation is in (the service
evaluation). An earlier commitment in effect foreshortens the service
evaluation as the "go-ahead" decision must be made before the service
evaluation is concluded.
This normal schedule for category 2 control schemes is summarized
below in figures 53 and 54.
D. Technology Category 1 Schedule (Normal)
It is next necessary to determine the changes from this schedule
that would occur with the utilization of technology from another tech-
nology complexity category. First, a normal schedule for category 1
technology is reviewed.
1. Concept verification. This task can be performed quite speedily
in view of the generally small amount of analysis and hardware develop-
ment needed to begin testing. Three months is adequate.
2. Combustor rig development. By definition, there are no category
1 concepts which materially affect the combustor. Therefore this step
is drastically reduced and constitutes little but verification, where
necessary, that any new components do not adversely affect the combustor
performance (which is unlikely in view of the character of the control
schemes). In some cases, this step can be deleted entirely (eg., in-
creased idle speed). Again, three months is sufficient.
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3. Engine demonstration. The importance of this phase of the
schedule for category 1 technology depends significantly upon the parti-
cular control concept involved. Engine testing for a fuel sectoring
concept is much more important (to check turbine wear) than it is for an
increased idle concept in which case there is no change in the engine or
its performance. In any event, engine testing is of generally less
importance for category 1 control as there are generally no significant
effects on the engine at power levels above idle. Three to six months
is all that is warranted here.
4. Production design and procurement. Again, this task is often
considerably reduced because of the limited nature of the changes to the
engine involved in most category 1 technology. The major exceptions to
this are the external air-assist concept and possibly airbleed in some
applications. This task should take from three to nine months, depending
on the concept involved with still three months overlap with the third
task, as before.
5. Engine testing. While one year's worth of testing may be
desirable for the air-assist concept, it is hardly necessary for the
other concepts of category 1, one-half that time is ample and even
superfluous in certain situations such as the increased idle speed
concept. Generally three to six months should be acceptable with special
consideration being recognized as necessary for development of the air-
assist concept. For comparison, AiResearch (reference 14) suggests
one year for all engine testing prior to certification (Task 3 and 5).
6. Flight testing. Once again the amount of time devoted to this
effort depends largely upon the concept involved, but it is probably
sufficient to devote to this the lower limit of the time that was devoted
to flight testing of category 2 technology, namely three months.
7. Qualification. There is little reason for change here. In
the case of increased idle speed, both steps 6 and 7, flight testing and
qualification, become trivial exercises requiring a minimal amount of
time.
8. Service evaluation. As the purpose of the service evaluation
is to determine through actual usuage the maintainability and performance
of the modification in question, then for category 1 technology wherein
the control concepts usually have little or no effect on the engine per-
formance only a minimal service evaluation is required, ranging from
nothing to one year. The one year figure would apply to the air-assist
concept for which the questions of durability, etc., must be addressed.
9. Production preparation. This stage would not generally be
necessary for most category 1 concepts, but would be required on the
normal schedule for air-assist and airbleed if changes have been made to
the original hardware in the latter case. Depending upon the control
concept in question, zero to six months should be allotted to this.
This normal schedule for technology category 1 control schemes is
summarized below in figure 55.
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E. Technology Category 3 Schedule (Normal)
Next, a normal schedule for category 3 technology is reviewed.
i. Concept verification. The nature of this task is not appre-
ciably different from that for category 2 technology so that the six
month time is adequate here also.
2. Combustor rig development. The time frames of the Phase II
parts of the NASA programs (ECCP and PRTP) are relevant for category 3
as indicated in the discussion earlier. Thus for large engines 15
months is necessary and for small engines, 9 months..
3. Engine demonstration. Again the NASA programs (category 3)
set the pace, 15 months for T2 class engines and 12 months for the Tl or
P2 classes.
4. Production design and procurement. No change from the category
2 situation appears warranted.
5. Engine testing. Probably no change from the category 2
situation is necessary, but in light of the complex fuel control which
is required (doubly so if staging must be performed in flight as with
the GE double annular combustor), additional engine testing may be
warranted. This would then require 18 months for this task.
6. Flight testing. In view of the major engine changes the full
six months should be used prior to qualification.
7. Qualification. No change. FAA qualification is fundamentally
the same: The hardware at this point has proven itself.
8. Service evaluation. The two year service evaluation recom-
mended by the industry is still acceptable in this case although it is
quite likely that problems will arise with such exotic systems in use
which will then require modification to the system, testing, and further
service evaluation. Therefore, it is quite plausible that a three year
period is warranted for this service evaluation.
9. Production preparation. No change from the category 2 time is
needed.
The normal schedule for technology category. 3 technology development's
summarized in Figure 56 and 57.
F. Accelerated Schedules
It is now necessary to investigate accelerated schedules, ones in
which some risk is assumed that if unexpected problems develop, the
impact or the schedule and on the cost is magnified considerably. The
following remarks apply to all of the schedules discussed above. Accel-
erations in a schedule will be given as a percentage of the normal
schedule times.
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1. Concept verification. For those manufacturers who must go
through this step there is not any easy way to avoid part of the work.
What is not done here (analysis, rig hardware manufacturing) must then
be done in step 2. Consequently, all the manufactuere can do is to push
for overtime, shop priority, etc. Total savings may be 20%.
2. Combustor rig development. The NASA (category 3) times are
quite long for what is the required minimum effort for they are based
upon the refinement of at least two different approaches. For those
manufacturers who have not yet conducted this task for one of their
engines, the NASA experience will be highly beneficial and much other-
wise wasted effort can be avoided by careful attention to the lessons of
the NASA ECCP and PRTP. For category 3, the savings may be 33%, while
for the other categories, perhaps 20% savings can be expected through
overtime, shop priority, etc.
3. Engine demonstration. Although NASA has allocated 15 months
for Phase III, it is worthwhile to note that the actual JT9D engine
demonstration is to be completed approximately 10 months after start of
the contract although the report will come out somewhat later. Consider-
ing this as a guideline as to the pace that can be achieved if there are
no hold-ups, it can be expected that an accelerated schedule here for
category 3 might save 33% again. For the other ratings, the same 20%
savings perhaps can be expected for reasons of overtime, etc.
4. Production design and procurement. While the usual accelera-
tion can be achieved (20% savings) through high priority treatment of
the project, further gains can be made by overlapping the engine demon-
stration phase by as much as 50% rather than the nominal 33% (at the
risk of wasted effort if there is failure).
5. Engine testing. Through priority treatment, this task can
easily be cut in half (savings of 50%).
6. Flight testing. Probably the best that can be expected here
is the lower limit of what the manufacturers have indicated they allow,
namely three months. Some acceleration may be possible, but safety
considerations in flight as well as the preparation for qualification to
follow weigh against this.
7. Qualification. This is essentially in the hands of the FAA:
Little time can be gained here from the manufacturers' effort. It is
reasonable to expect, however, that timely preparation of the material
by the manufacturer and expedited handling by the FAA should keep this
time to the minimum normal (3 months).
8. Service evaluation. Service evaluation can always be fore-
shortened, but at the expense of completeness of testing. There is no
way to have the same duration of testing in a shorter period as the
engines are used in airline service and there is no way to accelerate
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their usage. A shorter evaluation may lead to later problems as hap-
pened in the JT8D smoke retrofit program, especially if category 2 and 3
technology is being employed. The service evaluation for category 1
technology is already minimal. At best, for category 2 and 3 technology
a 25% savings can be hoped for if no problems arise during the evaluation.
Even then later problems may plague the airlines until final corrections
are made.
Of course, in the extreme, no specific service evaluation is re-
quired as is the case with a new engine. The initial experience in
using the engine in essence becomes a service evaluation. There is,
naturally, a resultant economic penalty. However, this extreme is not
assumed for the accelerated schedule.
9. Production preparation. For category 1 technology, this step
is not long enough (3 months) to warrant attempts to shorten it. For
technology category 2 and 3, however, some savings could be realized if
the task were shortened by 25% and moved up so that it would begin six
months prior to the end of the service evaluation. However, as noted in
the first discussion on scheduling, an earlier commitment in effect
foreshortens the service evaluation as the "go-ahead" decision must be
made before the service evaluation is concluded.
To compensate, if the airlines ensure the greatest possible in-
service time on the test engines, the probability of later rejection is
vastly diminished. Furthermore, with sufficient motivation, the manu-
facturer can tool for production in less than one year, especially as
some tooling from the manufacture of the service evaluation test engines
is available along with the necessary designs. If a total of nine
months is taken for the production preparation and if this phase is
begun three months prior to the end of the service evaluation, then only
six months beyond the end of the service evaluation is necessary before
production begins.
An alternative point of view which can be taken by the industry is
that a fixed period be allotted for service evaluation and preparation
for production. It is the industry's decision (manufacturers and
airlines) to agree among themselves upon the best breakdown between
service evaluation time and production preparation time. Such a flexi-
ble procedure, of course, allows the industry to minimize its cost
within the given constraint.
Figures 58-62 describe graphically the accelerated schedules for
the three categories of technology under consideration. These schedules,
while optimistic (allowing for no major setbacks), are realistic. In
this regard, the following remarks are relevant from material submitted
to the EPA in response to the NPRM of December 12, 1972, proposing air
pollution standards for aircraft and aircraft engines (FR, Vol. 37, No.
239).
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1. For the time required to go from concept identification
through engine demonstration (1 year, 8 months by Figure 60 or 2 years,
1 month by Figure 62):
Aerospace Industries Association - "It can be conservatively estimated
that at least eighteen (18) months are required to design, fabricate
hardware and conduct sufficient full scale engine testing to determine,
by further development, the useable potential of promising rig-developed
combustor designs or concepts."
Pratt and Whitney Aircraft - "It can be conservatively estimated that at
least two (2) years are required to design, fabricate hardware and
conduct sufficient full-scale demonstration engine testing to determine,
by further development, the useable potential of promising rig-developed
combustor designs or concepts."
2. For the time required to deliver the first production engine
once production design has been initiated (3 years, 1 month by Figures
60 or 4 years, 1 month by Figure 62):
Aerospace Industries Association - "Typically there is a four year lead
time required to develop an aircraft gas turbine engine from the ini-
tiation of engineering drawings until the first production engine is
delivered."
Pratt and Whitney Aircraft - "Past history has shown that it takes at
least four (4) years from design initiation through completion of the
certification testing to the delivery of the first production engine."
General Electric - "With regard to implementation dates, it should be
noted that several years of design, development, evaluation and pre-
production effort are normally required to provide an acceptable com-
bustor for use in a given engine application. Typically, this time span
ranges from three to four years."
G. APU Schedules (Normal)
Generally APUs, like the propulsion engines, must be flightworthy
as they are to be used as auxiliary power backup in flight. This means
that the development times will largely parallel that of the P2 class
with the notable exception that a lengthy service evaluation is not
required: Durability and maintenance of optimum performance of APUs are
not pursued by the users in view of their relatively small impact on the
total fuel consumption and maintenance burden. Furthermore, APUs are
not stressed and cycled in the same fashion as is a propulsion engine.
In view of these considerations, the compliance schedule for APUs would
differ from that for the propulsion engines as follows:
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APU Technology Category 1 and 2 Technology Schedules
2. Combustor rig demonstration. In view of the less stringent
requirements on the APU noted above, this task need not spend as much
time on performance development as is necessary for propulsion engines.
A 33% savings can be realized.
3. Engine demonstration. The remarks made relative to step 2 are
applicable here also.
5. Engine testing. The less stringent requirements on APUs would
lead to less need for this testing; a 50% time savings is possible.
6. Flight testing. This is necessary only for those engines
which have an in-flight function, but again, as reliability and maintain-
ability do not have the importance they have in propulsion engines, a
50% savings in time may be effected.
8. Service evaluation. As indicated above, a lengthy service
evaluation is not required. In fact, for the modest changes involved,
the service evaluation is largely unnecessary.
The other steps that have not been addressed here remain unchanged:
They require the same time for both the propulsion engines and the APUs.
Figures 63 and 64 present the normal APU schedules for category 1 and 2
technology.
APU Technology Category 3 Technology Schedule
Although the performance and service requirements on APUs are not
usually so strict as on propulsion engines, not too much opportunity is
offered by this fact to shorten the development schedule for the appli-
cation of category 3 technology into APUs. This is so because incorpora-
ting this technology into APUs has a number of difficulties unique to
the APU environment that would require special developmental effort.
Specifically, the combustor operating conditions and the combustor size
together create a situation in which much of the category 3 technology,
especially that which effects NOx control, is ineffective or physically
difficult to implement. Therefore, the schedule is likely to follow
that for propulsion engines except for step 8, service evalution.
8. Service evaluation. Because of the radical changes, an APU
employing category 3 technology should be expected to undergo some sort
of service evaluations although not necessarily to the extent of propulsion
engines. Somewhat arbitrarily, a six month period is allotted for this
task.
Figure 65 presents normal schedule for the implementation of cate-
gory 3 technology into APUs.
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H. Accelerated APU Schedules. There is no reason why accelerated
schedules here should differ in character from those for propulsion
engines. Hence, those percent savings are assumed to apply and figures
66, 67, and 68 summarize the accelerated APU schedules for the three
technology categories.
I. Schedules for Newly Certified Engines
It is assumed that any standard for newly certified engines will be
based upon the use of at least category 3 technology in order to achieve
effective NOx control without the use of water injection. The schedule
then should be roughly the same as that which applied to newly manufac-
tured engines with the exception that a service evaluation is not required
here. A service evaluation is a practice which has been established to
provide a measure of the economic impact to the airlines, resulting from
equipment modifications. This knowledge, is quite important to the
airlines as older engines have established high levels of durability and
maintainability which make their operation very economically attractive.
Any change in the engine is expected, or at least, hoped to preserve
that economic performance. In the case of a newly desiged engine there
is no prior performance record; the introduction into service and the
initial experience constitute the baseline economic performance record
which is a service evaluation of sorts.
In addition, it is reasonable to expect that if there is an engine
now in development which is expected to be certified after the January
1, 1981 date specified in the existing regulations, that program has
incorporated category 3 technology into its combustor development effort.
As that technology development has nearly reached the production design
stage (i.e., NASA ECCP), it follows from Figure 57 that newly certified
engines utilizing category 3 technology (NOx control) could be available
by January 1, 1981. For this estimate a normal schedule has been assumed
as it is unlikely that development of a full engine could proceed at any
faster pace.
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TECHNOLOGY CATEGORY 2
T2 CLASS ENGINES
NORMAL SCHEDULE
Years
4
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
\
\
\
\
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I
M
o
I
Figure 53
-------
TECHNOLOGY CATEGORY 2
Tl & P2 CLASS ENGINES
NORMAL SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
\
\
\
\
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\
\
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Figure 54
-------
o
CO
I
TECHNOLOGY CATEGORY 1
CLASS Tl, T2 & P2 ENGINES
NORMAL SCHEDULE
Years
3 4 5 6
! 10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation *
Production
Preparation
I
£
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Figure 55
* One year service evaluation is required for the air-assist concept
-------
I
M
o
I
0
TECHNOLOGY CATEGORY 3
T2 CLASS ENGINES
NORMAL SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
\
\
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\
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Figure 56
-------
TECHNOLOGY CATEGORY 3
Tl & P2 CLASS ENGINES
NORMAL SCHEDULE
Years
2345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
\
\
\
\
\
\
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Figure 57
-------
TECHNOLOGY CATEGORY 1
CLASS Tl, T2 & P2 ENGINES
ACCELERATED SCHEDULE
Years
! 3 4 5
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation *
Production
Preparation
S
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Figure 58
* One year service evaluation is required for the air-assist concept
-------
TECHNOLOGY CATEGORY 2
T2 CLASS ENGINES
ACCELERATED SCHEDULE
Years
4
10
t-o
I
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
\
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Figure 59
-------
TECHNOLOGY CATEGORY 2
Tl & P2 CLASS ENGINES
ACCELERATED SCHEDULE
Years
4
10
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I
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
:
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Figure 60
-------
TECHNOLOGY CATEGORY 3
T2 CLASS ENGINES
ACCELERATED SCHEDULE
Years
(
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
T?\/a 1 na-f- 1 OH
Production
Preparation
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Figure 61
-------
i
i-1
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Ul
TECHNOLOGY CATEGORY 3
Tl & P2 CLASS ENGINES
ACCELERATED SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
.
Product ion
Preparation
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-------
TECHNOLOGY CATEGORY 1
APU
NORMAL SCHEDULE
Years
345
10
Concept
Identification
Combust or
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
I
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Figure 63
-------
TECHNOLOGY CATEGORY 2
APU
NORMAL SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
\
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Figure 64
-------
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00
TECHNOLOGY CATEGORY 3
APU
NORMAL SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
\
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Figure 65
-------
TECHNOLOGY CATEGORY 1
APU
ACCELERATED SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
3
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Figure 66
-------
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o
TECHNOLOGY CATEGORY 2
APU
ACCELERATED SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
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Figure 67
-------
I
M
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TECHNOLOGY CATEGORY 3
APU
ACCELERATED SCHEDULE
Years
345
10
Concept
Identification
Combustor
Demonstration
Engine
Demonstration
Production
Design and
Procurement
Engine
Testing
Flight Testing
Qualification
Service
Evaluation
Production
Preparation
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-------
SECTION V: Industry Status
This section presents a summary of the specific control techniques
(see Appendix C for descriptions) employed by each manufacturer for each
engine in attempting to meet the gaseous emission standards. The infor-
mation presented indicates the amount of effort put forth by each manu-
facturer in attempting to comply with the regulations for each engine*,
what has been accomplished, the status of each engine, and what can be
expected. Also, a brief profile of each engine is included. Since
approximately three and a half years have passed since final promulga-
tion of the aircraft emission regulations, an engine's status with
respect to the various control concepts under development is presented
relative to the lead time schedules developed in Section IV. The dates
projected for the implementation of a given control technology concept
into newly manufactured engines are based on an accelerated schedule.
It should be noted that for all of the propulsion engines (Class
Tl, T2, T3, T4, and P2 engines) the control techniques listed for each
engine are dry concepts (i.e. do not include water injection for NOx
control). Further, the prospects for meeting the 1979 NOx standards,
which are given in the profiles, are estimated on the basis of dry
technology. Water injection in sufficient quantities could be used to
control NOx to the required level on any propulsion engine, if necessary,
although care must be taken that not too much water is injected or CO
will climb to excessive levels due to the quenching effect of the water.
Most manufacturers have demonstrated (development status task no. 3)
water injection (category 2 technology) on their respective engines.
Hence the estimated implementation dates for incorporating water injec-
tion into newly manufactured engines are late 1979 for Tl and P2 class
engines (see Fig. 60) and mid 1980 for T2, T3, and T4 class engines (see
Fig. 59). This time estimate is considered valid also for engines that
currently use water injection for thrust augmentation because in the
augmentation application water is used at takeoff only and is delivered
at a constant rate. For NOx control water would be injected at takeoff
and also at climbout and would be injected at a rate proportional to the
fuel flow rate. Hence a water injection control would have to be developed
and tested.
A. Tl Class Engines
Class Tl engines are described (FR Vol. 38, N. 136, July 17, 1973,
p. 19091) as "...all aircraft turbofan or turbojet engines except
engines of Class T5 of rated power less than 8,000 pounds thrust."
Generally engines in this class are used on general aviation jet aircraft.
The gaseous emissions standards listed below for this class of engines
are presently scheduled to go into effect on January 1, 1979.
*Although it is possible that the manufacturers have expended a greater
effort than reported herein, such information has not been made avail-
able to the EPA.
-123-
-------
Standard
EPAP
(Ibs pollutant per 1000 Ibf-hours over the cycle)
HC CO NOx
1.6 9.4 3.7
1. AiResearch
a. TFE-731
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 1 technology by early 1979
NOx - poor with category 2 technology;
- good with category 3 technology by late 1981
Thrust = 3500 Ibs (-2 model)
BPR = 2.66
PR = 13
Combustor type: reverse flow annular
Application: Learjet 35/36 (business jet)
Certification date: August '72
Number delivered: more than 100 for all models combined
as of January '75
Production rate: 25 per month for all TFE-731 models combined
Production category after January 1, 1979: III*
Effort Expended
This engine has been selected for testing in the NASA Pollution
Reduc-tion Technology Program. Phase I of the anticipated three phases
of this program began in November 1974 and was completed in June 1976.
According to AiResearch (Ref. 14) Phase I was to cost 693,000 dollars on
a cost-share basis, NASA 80% and AiResearch 20% (see Appendix A, Table
A-l). During Phase I six versions each of three basic NASA combustor
concepts were selected for screening tests. These concepts vary in
degree of complexity and thus have varying potential for achieving
satisfactory emissions control. The basic concepts selected are listed
below along with a brief description of the features of each version.
NASA Concept 1 retrofittable modifications to production combustor
(roughly category 1 technology)
Version Features
Baseline 1. up to 22% bleed at idle
2. air-assist through secondary passage
of duplex nozzle at idle
* This refers to the production categories (expected production situation)
defined in part B of Section III.
-124-
-------
NASA Concept 1 (continued)
Version
Mod I
Mod II
Mod III
Mod IV
Mod V
Features
3. water-methanol injection at takeoff
provisions made for sector burning at idle
increased swirler airflow
piloted airblast fuel nozzles
improved recirculation pattern by making
hole pattern changes in the primary zone
variable primary-zone equivalence ratio
NASA Concept 2
air assisted/airblast fuel system (roughly category
2 technology)
Version
Baseline
Mod I
Mod II
Mod III
Mod IV
Features
1. 20 (compared to 12 for a production
combustor) fuel nozzles
2. nozzles are combination air-assist/
airblast type instead of the duplex
used in production
3. fuel nozzle swirlers are replaced by
grommets at low power to simulate a
proposed variable geometry device
moved row of primary zone holes upstream on
both inner and outer liners
added row of primary zone holes to outer
combustor liner; increased swirler airflow
doubled number of holes added in Mod II and
reduced size of each; moved holes upstream for
earlier quench; reinstalled baseline swirlers;
larger primary zone holes moved downstream to
original position
reduced number of outer liner primary zone
holes and increased in size; moved them
downstream also; primary zone cooling was
reduced 33%
-125-
-------
NASA Concept 2 (continued)
Version Features
Mod V reduced number of outer-liner primary zone
holes, increased in size, and moved down-
stream; reinstalled larger swirlers; restored
primary zone cooling air to original amount
Refinement I unknown
Refinement II unknown
NASA Concept 3 piloted premixing/prevaporizing fuel system
(category 3 technology)
Version . Features
Baseline 1. 20 simplex fuel nozzles in pilot zone
2. forty individual premix/prevaporizing
tubes in main zone; these contain
simplex nozzles for fuel injection
Mod I increased pilot zone holes on outer liner from
40 to 120; closed pilot zone cooling holes
on outer and inner liner
Mod II added transpiration cooling holes to inner
liner
Mod III unknown
Mod IV unknown
Mod V unknown
Refinement I unknown
Refinement II unknown
Phase II (the combustor-engine compatability testing phase) began
in June, 1976 and is currently at the design and fabrication task.
Candidates for Phase II rig testing are the airblast/air assist system
(NASA concept 2), the premix/prevaporization system (NASA concept 3),
and a combination of these. This phase of the anticipated three phase
project is expected to last 14 months. No dollar estimate of cost is
available for Phase II.
-126-
-------
Several other category 1 technology concepts were tested "in-house"
prior to the start of the NASA work. These were 1) taxi-idle on primary
fuel nozzles only, 2) taxi-idle with air assisted primary fuel nozzles
(parametric study from 0 to 2.54 Ibs/min) and 3) taxi-idle with compressor
bleed (parametric study from 0 to 20% compressor airflow). According to
the data submitted to the EPA (Ref 14) this "in-house" testing began in
1972. Information quantifying the effort expended for this work has not
been submitted to the EPA. However, it should be noted that all of
these "in-house" concepts are category 1 concepts which would require
the least amount of effort to develop and test.
Results and Status
The most significant results of the TFE-731 tests are summarized in
the status table below.
Status (-2 model)
Concept
Standard
Production
(avg. of 6)
2.54 Ibs/min 1
air assist at idle
Piloted airblast* 2
Mod 5
Technology
Category HC
Development
EPAP Status Implementation
C£ NOx (Task No.) Date
1.6 9.4 3.7
6.6 17.5 5.0
Jan. 1, 1979
0.5 6.7 5.6 2 (Fig. 58) early 1979
0.2 9.2 4.2 2 (Fig. 60) mid 1980
Premix/prevap.* 3-4 0.6 9.2 2.7 2 (Fig. 62) late 1981
Mod 3
*NASA PRTP results
The data tabulated above indicates that levels of CO and HC below
the standards can be achieved with either air-assist or airblast. The
air-assist pressure differential (260 psid) required for 2.54 pounds per
minute airflow is considered by AiResearch to be higher than can be
achieved with a reasonable compressor discharge boost system. An
alternate external air-assist system may be required. With both the
air-assist and the airblast concepts, NOx remained above the standards.
The airblast did reduce NOx by 15 percent; however a 35 percent reduc-
tion is required. The premix/prevaporization concept is capable of
providing the needed reduction in NOx while controlling HC and CO to a
level below the standards.
-127-
-------
Projections and Prospects
As noted in the Effort Expended section, the Phase II candidate
combustor concepts are the airblast/air-assist system, the premix/pre-
vaporization system, and a combination of these. Both concepts have
demonstrated the capability of controlling HC and CO to levels below the
standards. So far only the premix/prevaporization combustor has demon-
strated the capability to control NOx to a level below the standards.
It is expected that development of these concepts will be pursued,
especially the premix/prevaporization concept since these concepts offer
the best known possibilities for compliance with the standards.
b. ATF3
Profile
Prospects for meeting the 1979 standards:
HC, CO - unknown because of insufficient data
to make judgment
NOx - unknown because of insufficient data
to make judgment
Thrust = ^ 5000 Ibs
BPR = 3
PR = 22-24
Combustor type: reverse flow annular
Application: potential use for business jets
Certification date: not certified for commercial use
Number delivered: none for commercial use
Production rate: not yet produced for commercial use
Production category after January 1, 1979: IV
Effort Expended
This engine currently has military applications only. However,
because of the relatively low rated SFC (0.45 Ibm/lbf-hr) this engine
has potential application for use on business jets. Probably no effort
has been made to reduce the level of gaseous emissions from this engine
because of its military application. The only emissions testing that
has been done on this engine was conducted to determine the baseline
emissions level.
Results and Status
The level of gaseous emissions produced by this engine is summarized
in the status table below.
-128-
-------
Concept
Standard
Baseline
Technology
Category HC
1.6
5.8
Status
Development
EPAP Status
C£ NOx (Task No.)
9.4 3.4
16.9 4.1
Implementation
Date
Jan. 1, 1979
Projections and Prospects
The potential for reducing HC and CO should be good because of the
relatively high pressure ratio of this engine. However, this potential
may be limited because of a high surface-to-volume ratio of the reverse
flow annular combustor. No estimates of low emissions levels can be
made because of insufficient data reported about this engine.
-129-
-------
2. Pratt & Whitney Aircraft of Canada
a. JT15D
Profile
Prospects for meeting 1979 standards:
HC, CO - fair with category 2 technology by late 1979
NOx - poor with category 2 technology;
- good with category 3 technology* by late 1981
Thrust = 2500 Ibs.
BPR =3.3
PR = 12
Combustor type: reverse flow annular
Application: Aerospatiale Corvette (business jet)
Certification date: September '73
Number delivered: 50 as of February '75
Production rate: unknown
Production category after January 1, 1979: III
*CO is a problem with category 3 technology
Effort Expended
Low emissions research specifically aimed at reducing JT15D emissions
began in 1971 (Ref. 20). For the years 1971 and 1973 a total of 255,000
dollars were spent by PWAC on low emissions research for this engine.
No effort was expended in 1972 for JT15D low emissions testing. Informa-
tion quantifying the effort expended for the years after 1973 is not
available for this engine.
Low emissions concepts (reported in Ref. 20 and Ref. 32) that have
been tested on this engine are 1) compressor bleed at idle and approach,
2) non-optimized airblast nozzles, 3) rich primary plus delayed quench,
and 4) rich primary plus delayed quench plus increased swirl. The
extent of testing of these concepts for optimization is not known. The
most significant results of this testing are tabulated in the status
summary below.
Results and Status
Status (-4 model)
Development
Technology EPAP Status Implementation
Concept Category HC C£ NOx (Task No.) Date
Standard 1.6 9.4 3.7 Jan. 1, 1979
Production 12.4 34.6 3.8
(avg. of 6)
10% bleed at 1 5.2 30.0 4.5 5 (Fig. 58) mid 1978
idle & approach
Rich primary 2 0.14 10.9 4.8 3 (Fig. 60) late 1979
plus increased
(37°) air inlet
swirl angle
-130-
-------
The JT15D has an unusually high idle SFC characteristic for its
cycle which adversely affects emissions, especially HC and CO, as the
data in the summary above indicates. The modified airflow combustor
(rich primary, etc.) has reduced the HC and CO emissions by 99 percent
and 68 percent respectively. However, NOx increased 26 percent. To
achieve these levels of emissions control the primary zone fuel-air
ratio was increased by 20 percent arid the intermediate zone fuel-air
ratio was decreased 50 percent. This combustor also utilized "delayed
quench" which increased the residence times, and swirling air-blast
which improved the spray quality. Problems associated with this concept
are 4.3% SFC penalty and 6% thrust penalty.
Projections and Prospects
The levels of CO and NOx remain above the standards with either of
the reduced emission concepts listed in the status summary. A concept
to consider for CO control is the use of bleed at idle and approach
along with the airflow redistribution. If the air bleed would be as
effective on the modified combustor as it was on the production combus-
tor the first set of values listed in the projected status table below
would result. Another concept to consider for CO control is to use air-
assist either on the production combustor or in conjunction with the
modified combustor. The effectiveness of this concept has already been
demonstrated on the TFE-731 engine. Exact levels of emissions cannot be
estimated for the JT15D engine since it is not known what if any combustor
optimization would be required along with the air-assist.
In order to control NOx it appears that the premix/prevaporization
concept being developed by NASA and AiResearch will be required. With
this technique it is estimated (see Appendix E) that NOx will be reduced
to a EPAP of approximately 3.0. However, the CO EPAP is estimated to be
approximately 15.0. The high CO results from the high idle SFC are
noted above in Results and Status. Some further CO control concepts
would be required. A summary of these projected results is presented
below.
Projected Status
Concept
Standard
Technology
Category
Rich primary 2
plus delayed
quench plus
increased inlet
swirl plus bleed
Premix/prevap. *3-4
*NASA technology
1.6
0.1
Development
EPAP Status
CO NOx (Task No.)
9.4 3.7
9.4 5.7 3 (Fig. 60)
Implementation
Date
Jan. 1, 1979
late 1979
1.0 .15.0 3.0 2 (Fig. 62) late 1981
-131-
-------
The implementation date for the rich primary concept plus bleeds is
estimated to be the same as that for the rich primary only concept since
the addition of bleeds should not require any extra testing or develop-
ment. The premix combustor was placed at task 2 (and hence the 1981
implementation date) since Pratt & Whitney Aircraft of Canada has
probably investigated both premix and staged concepts. EPA holds this
opinion in view of the shared technology between Pratt and Whitney Air-
carft of Canada and Pratt and Whitney Aircraft (East Hartford, Conn.)
where premix and staged concepts have already been tested for use on
the JT9D engines. No data has been reported, however.
-132-
-------
3. Lycoming
a. ALF-502D
Profile
Prospects for meeting 1979 standards:
HC, CO - poor with category 1 technology;
- fair with category 2 technology by early 1979;
- good with category 3 technology by late 1981
NOx - good with category 2 technology by early 1978;*
- good with category 3 technology by late 1981
Thrust = 6500 Ibs.
BPR = 6.0
PR = 10.7
Combustor type: reverse flow annular
Application: Rockwell Sabreliner (proposed)
Certification date: June 76
Number delivered: none
Production rate: unknown
Production category after January 1, 1979: IV
*HC and CO are above standard with this concept
Effort Expended
Information submitted to the EPA (Ref. 35) indicates that a program
to reduce emissions frpm the ALF-502 engine began in 1973. No information
quantifying the amount of manhours and/or money spent in this effort is
available. However, it should be noted that all of the ALF-502 low
emissions effort has involved category 1 and 2 technology concepts which
are the least costly to develop and test.
According to Lycoming (Ref. 35), the initial design of the ALF-502
used dual orifice fuel injectors. Efforts have been made to reduce
emissions by incorporating airblast fuel nozzles, and by making changes
to the primary zone fuel air ratio, liner wall cooling, primary zone
mixing, and primary zone residence time. Lycoming claims (Ref. 35)
these tests were designed to encompass the maximum range of airflow
distribution possible within this engine and produced such small reduc-
tions in emissions that the tests were of little value except to indicate
that the EPA 1979 emissions requirements cannot be accomplished through
these means. The exact number of configurations tested, and their
respective results have not been reported to the EPA (except for the
airblast/ combustor modification concept listed in the status table).
Results and Status
The airblast concept noted above was capable of controlling HC and
NOx to a level below the standards. However, it is unknown whether or
not this configuration represents the optimum airblast-airflow dis-
tribution combination. Lower emissions (expecially HC and CO) may be
possible with the airblast-airflow redistribution concept. Exact levels
cannot be estimated because of insufficient data.
-133-
-------
The available results of the the ALF-502 tests are summarized in
the table below.
Status
Technology
Category
HC
1.6
3.3
2.3
1.5
2.6
Development
EPAP Status
CO NOx (Task No.)
9.4 3.7
23.4 3.5 3 (Fig. 58)
11.9 4.1 5 (Fig. 58)
11.9 3.0 5 (Fig. 60)
17.8 3.2 7 (Fig. 60)
Implementation
Date
Jan. 1, 1979
late 1978
mid 1978
early 1979
early 1978
Concept
Standard
Original
dual orifice
Original air- 1
blast config.
Airblast/com- 2
bustor mod.
Cert, config.* 2
(avg. of 2)
Projections and Prospects
Lycoming anticipates using the technology which has been demonstrated
in the NASA ECCP and PRTP programs to assist in the choice of concepts.
However, the external air assist (technology category 1) should be
considered for reducing emissions, especially HC and CO. As noted in
the JT15D section, the effectiveness of air-assist has been demonstrated
on the TFE-731 engine. Exact values of emissions cannot be estimated
since combustor optimization may be required. It is estimated (similar
to method of Appendix E) that the NASA Tl premix/prevaporization concept
would control all three gaseous emissions to a level below the standards.
The projected levels are presented in the table below.
Technology
Concept Category
Standard
Premix/prevap.** 3-4
**NASA technology
Projected Status
Development
EPAP Status
HC CO NOx (Task No.)
1.6 9.4 3.7
0.5 8.9 2.5 2 (Fig. 62)
Implementation
Date
Jan. 1, 1979
late 1981
EPA understands (reference 91) that Lycoming has looked into
premix and staged combustors for the ALF-502 engine and it estimated
that development has reached the task 2 stage at this point.
*This configuration was labeled certification configuration by Lycoming but,
it is not the configuration that was ultimately certified. Changes in the
liner cooling were made because of a durability problem. Emissions data
from the certified engine are not available at this time.
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4. General Electric
The CF700 and the CJ610 engines have been combined for discussion
purposes because of their basic core commonality and their identical
combustors. Further, General Electric's efforts to reduce emissions
from these engines have also been combined since low emissions concepts
should be equally useful on either engine.
a. CF700
Profile
Prospects for meeting 1979 standards:
HC, CO - poor with either category 1 or 2 technology
NOx - already met
Thrust = 4300 Ibs.
BPR = 1.6
PR = 6.6
Combustor type: straight flow annular
Application: Dassualt Falcon 20 1/1.., \
, ,, , , . -,,-»/ (business lets)
Rockwell Sabreliner 75A ^ J /
Certification date: July '64 (-2B model)
Number delivered: 919 thru 1975; 1107 expected thru 1978
Production rate: combined with the CJ610 - 200 per year
Production category after January 1, 1979: I
b. CJ610
Profile
Prospects for meeting 1979 standards:
HC, CO - poor with either category 1 or 2 technology
NOx - already met
Thrust = 2950 Ibs.
BPR = 0 (turbojet)
PR = 6.8
Combustor type: straight flow annular
Application: Learjet 24/25 (business jet)
Certification date: Dec. '61 (-1 model)
Number delivered: 1648 thru 1975; 1840 expected thru 1978
Production rate: combined with the CF700 - 200 per year
Production category after January 1, 1979: I
Effort Expended
Efforts to reduce emissions from these two engines began in 1973
according to information available (Ref. 5) to the EPA. According to
General Electric (Ref. 5) these efforts have been directed at determining
-135-
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the feasibility and practicality of meeting the emissions standards.
Reduced emissions concepts that have been tested for application on both
engines are 1) three airflow distribution concepts designed to lean the
primary zone 2) three fuel injection concepts to reduce fuel impingement
on the combustor walls, and 3) air assisted fuel nozzles. General
Electric projected in August 1975 (Ref. 4) that the total cost for a
combustor redesign (i.e. airflow redistribution) program through flight
testing (line 5 Fig. 59) would be approximately 6.4 million for both
engines combined, and that the total cost for a fuel nozzle redesign
program (i.e. reduced impingement) through flight testing (line 5 Fig.
58) would be approximately 690,000 dollars for both engines combined.
It is not known how much money was actually spent on low emissions
t es t ing, however.
General Electric has also tested the effects of three engine
operating changes on the CJ610 engine. These changes were increased
customer bleed, increased idle thrust, and operating on primary fuel
injection alone. No dollar or manhour data is available for this test-
ing. However, it should be noted that these changes are relatively easy
to test and therefore the effort required was minimal.
Results and Status
These two engines represent older technology and hence have rela-
tively high SFCs. The HC and CO emissions indices at idle are also very
high due to the very rich idle operation of these engines (primary zone
equivalence ratio of 1.0 to 1.5). This results in especially high CO
emissions. One category 2 concept resulted in reduced CO emissions as
a result of a leaner primary zone at idle. However, this was not
nearly enough to meet the standards. Some success has been achieved in
reducing the HC emission, as a result of the reduced impingement, but
still not to a level below the standards. The available data is summarized
in the status tables below. Apparently General Electric felt it was not
necessary to test NOx control concepts because these engines already
meet the NOx standards.
-136-
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CF700 Status
Technology
Concept Category HC
Standard 1.6
Production 9.1
Reduced spray 1 1.6
cone angle
(fuel injector
mod)
air-assisted 1 11.3
fuel nozzles
addition of 1 17.9
venturi design
to swirl cup
(fuel injector
mod)
addition of 1 5.3
radial air inflow
to swirl cups (fuel
injector mod)
addition of inner 2 9.7
air scoops (air-
flow redist.)
addition of 2
radial air
EPAP
CO
9.4
94
94
99
121
97
78
82
NOx
3.7
2.2
2.2
No
data
No
data
No
data
No
data
No
data
Development
Status Implementati
(Task No.) Date
Jan. 1, 1979
2 (Fig. 58) early 1979
2 (Fig. 58) mid 1979*
2 (Fig. 58) early 1979
2 (Fig. 58) early 1979
2 (Fig. 60) early 1980
2 (Fig. 60) early 1980
inflow to swirl
cups plus airflow
redistribution
*Includes extra six months for service evaluation.
-137-
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CJ610 Status
Technology
Concept Category HC
Standard 1.6
Production 14.7
Reduced spray 1 2.5
cone angle (fuel
injector mod)
air-assisted 1 18.2
fuel nozzles
addition of 1 29.0
venturi design
to swirl cup (fuel
injector mod)
addition of 1 8.5
radial air in-
flow to swirl
cups (fuel
injector mod)
4.2% customer 1 18.8
bleed
increased idle 1 13.5
thrust
primary injectors 1 14.7
only at idle
addition of 2 15.7
inner liner air
scoops (airflow
redisturbution)
addition of 2
radial air inflow
EPAP
CO
9.4
155
155
163
200
160
180
129
202
129
136
NOx
3.7
2.7
2.7
No
data
No
data
No
data
No
data
No
data
No
data
No
data
No
data
Development
Status
(Task No.)
2 (Fig. 58)
2 (Fig. 58)
2 (Fig. 58)
2 (Fig. 58)
3 (Fig. 58)
3 (Fig. 58)
3 (Fig. 58)
2 (Fig. 60)
2 (Fig. 60)
Implementati
Date
Jan. 1, 1979
early 1979
mid 1979*
early 1979
early 1979
late 1978
late 1978
late 1978
early 1980
early 1980
to swirl cups plus
airflow redist.
*Includes extra six months for service evaluation.
-138-
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Projections and Prospects
The CO emissions can be expected to remain high unless the fuel air
ratio can be made lean at idle. It is very difficult to estimate to
what level the emissions can be reduced: Any estimate would require
very large extrapolations of available data no attempt to do so has been
made here.
It should be noted that the implementation dates estimated for the
operational changes are based on Figure 58. However, for the increased
customer bleeds and the increased idle thrust the remainder of the
development tasks (tasks 4 thru 9) could be eliminated. These two
operational changes could be used in current production engines without
further development or testing.
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5. Pratt & Whitney Aircraft
a. JT12A
Profile
Prospects of meeting 1979 standards;
HC, CO - poor with current production combustor
NOx - already met
Thrust = 3300 Ibs. (-8 model)
BPR = 0 (turbojet)
PR = 6.7
Combustor type: multiple can
Application: Lockheed Jetstar , ,, ,.
_,,,_,,. } (business jets)
Rockwell Sabrelmer J
Certification: c. 1960
Number delivered: 2900 as of December 1974
Production rate: unknown
Production category after January 1, 1979: I
Effort Expended*
Pratt & Whitney Aircraft reported in Reference 12 that the only
emissions work conducted on the JT12 engine consisted of a limited
amount of measurements to evaluate the general emissions characteris-
tics. Data reported in Reference 30 indicates that a total of approxi-
mately 15,000 dollars 'was spent for both visible and invisible exhaust
emissions work on this engine during the years 1971, 1972 and 1973.
Results and Status
This engine meets the NOx standard with the production combustor
due to the relatively low pressure ratio (6.7). The HC and CO emissions
are however, very high as1a result of the unfavorable operating conditions
(low pressure and high fuel flow rate).
Since no low emission combustor concepts have been tested on this
engine the only data reported was from the current production,model.
This data is tabulated in the status table below.
Status (-8 model)
Development
Technology EPAP Status Implementation
Concept Category HC C£ NOx (Task No.) Date
Standard 1.6 9.4 3.7 Jan. 1, 1979
Production 5.2 85 3.2
*It should be noted that data reported in Reference 30. indicates that
7,450,000 dollars were spent by Pratt & Whitney Aircraft on basic emissions
research and development (i.e., not engine specific) during the time period
1971-1975. _140-
-------
Projections and Prospects
It is doubtful that any reasonable combustor concept would be
sufficient to bring this engine into compliance because of the unfav-
orable operating environment noted above. It is further doubtful that
any attempt would be made because of the limited remaining production of
this engine.
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6. Rolls Royce
The limited submission of emissions and engine performance data by
Rolls Royce as of the time of this report did not permit the EPA to make
a satisfactory assessment of the status of Rolls Royce. The data
tabulated for each engine in the respective status tables represents the
only data available for each Tl Class Rolls Royce engine. The reduced
emissions technology concepts are unknown but are known to be category 1
or 2 technology. Further, the EPA is not sufficiently familiar with the
Rolls Royce test programs to make accurate judgments with respect to
combustor development times. Therefore the estimates listed are not as
precise as for other engines.
a. M45H
Profile
Prospects for meeting 1979 standards:
HC, CO - good with category 1 or 2 technology by late 1979
NOx - good with category 1 or 2 technology by late 1979
Thrust = 7600 Ibs.
BPR =3.0
PR = 16
Combustor type; straight flow annular
Application: VFW 614
Certification date: August 1974
Number delivered: unknown
Production rate: unknown
Production category after Jan. 1, 1979: III,IV
Effort Expended
Data quantifying the amount of manhours or money spent for reduced
emissions development and testing on this engine have not been reported
by Rolls Royce. Data from only one low emission concept has been report-
ed for this engine. No technical information about the concept is
available. Further it is not known if any other reduced emissions concepts
have been tested for use on this engine.
Results and Status
The data reported for this engine using the unknown combustor
concept indicates that this engine will be able to meet the level of the
standards. It is not known at what task this unknown combustor is in
development. It seems probable that this concept is at the engine
demonstration task or beyond as some emissions development should have
begun prior to selection of the final configuration of this engine.
Allowances in the engine should have been made for this combustor so
that the hardware changes required to implement it (and hence time)
should be a minimum. The status of this engine is tabulated in the
table below.
-142-
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Concept
Standard
Production
Unknown
Technology
Category ![£.
1.6
6.0
1 or 2 0.5
Status
Development
EPAP Status
CO NOx (Task No.)
9.4 3.7
40.7 4.2
5.3 3.3 3 (Fig. 60)
Implementation
Date
Jan. 1, 1979
late 1979*
* This estimate is based on the assumption that the unknown combustor
concept is category 2 technology. Rolls Royce indicated that their
reduced emissions combustors would not be available before 1981. Their
estimate probably is based upon "normal" time schedule.
Projection and Prospects
Since this engine will be able to comply with the standards with the
unknown combustor, probably no further reductions can be expected from
it.
b. RB401
Profile
Prospects for meeting 1979 standards:
HC, CO - good with category 1 or 2 technology by early 1979
NOx - good with category 1 or 2 technology by early 1979
Thrust = 5400 Ibs.
BPR =4.5
PR = 17
Combustor type: straight flow annular
Application: Business jets
Certification date: estimated to be late 1977
Number delivered: none - deliveries are projected to begin in 1978
Production rate: engine still in development
Production category after January 1, 1979: IV
Effort Expended
This engine is still in development (certification is expected in
1977) so it is expected that low emissions combustor concepts are included
in the initial design. Any low emissions development and testing should
-143-
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be part of the normal development and testing of this engine. No data
is available which quantifies the amount of manhours or money spent on
the development of the combustor. However, the amount of time and money
spent should be typical of any new engine combustor development program,
and would not necessarily increase because of designing for low emissions.
Results and Status
Only baseline emissions have been reported for this engine. With
this baseline configuration HC and NOx standards were achieved, and CO
exceeds the standard by about nine percent. The exact EPAP values are
tabulated in the status table below.
Status
Concept
Standard
Baseline
Technology
Category HC
1 or 2
1.6
0.2
Development
EPAP Status
CO NOx (Task No.)
9.4 3.7
10.2 3.6 5 (Fig. 60)
Implementation
Date
Jan. 1, 1979
late 1977*
*Same as the certification date.
Projection and Prospects
According to "Jane's All the World's Aircraft 1975-76" (Ref. 69)
this engine was designed to meet the existing emissions standards.
If the information supplied to Jane's came from the manufacturer
directly, it would suggest that the information supplied to the EPA
is incomplete and that lower levels of CO and possibly NOx might be
expected.
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B. T2 Class Engines
Class T2 engines are defined (FR Vol. 38, N. 136, July 17, 1973, p.
19091) as "...all turbofan or turbojet aircraft engines except engines
of Class T3, T4, and T5 of rated power of 8,000 pounds thrust or greater".
This class consists of large turbojet and turbofan engines intended for
commercial subsonic transports. The gaseous emissions standards,
listed below, for this class of engines are presently scheduled to go
into effect on January 1, 1979. An NPRM proposing that in-service
engines in this class of greater than 29,000 Ibs. thrust be retrofit
with combustors that control emissions to the standards listed below
was issued July 17, 1973. The proposed schedule requires that the
retrofit of these engines be completed by January 1, 1983.
Standard
EPAP
(Ibs pollutants per 1000 Ibf-hours over the cycle)
HC CO NOx
0.8 4.3 3.0
1. Pratt & Whitney Aircraft
a. JT9D-7
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 2 technology by late 1979
NOx - poor with category 2 technology;
- good with category 3 technology by mid 1981
Thrust = 46,150 Ibs.
BPR =5.15
PR = 22.3
Combustor type: straight flow annular
Application : 747-200
Certification date: June '71
Number delivered: 1400 for all JT9D models combined as of early '75
Production rate: unknown
Production category after January 1, 1979: ' III
Effort Expended
Attachment A of Reference 30 indicates that for the period 1971
through 1975 Pratt & Whitney Aircraft spent approximately three million
dollars on development and testing of their own "in-house" low emissions
concepts for the JT9D engines (both the JT9D-7 and JT9D-70). The
concepts tested for JT9D-7 application (reported in Ref. 12) are com-
pressor bleed and horsepower extraction at idle (category 1), air-assist
(category 1), and aerating nozzles (category 2).
-145-
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The JT9D-7 was also selected for testing in the NASA Experimental
Clean Combustor Program (ECCP). This program was conducted in three
sequential phases, each individually funded on a cost share basis, with
a total of 3.25 million dollars spent (80% NASA and 20% PWA). A description
of the efforts, expenditures, and the combustor concepts tested is
presented below for each phase.
Phase I; Combustor Screening - Phase I began in December 1972 and
was completed in June 1974. Available data indicates that NASA provided
a total of 800,000 dollars, while PWA provided the hardware required for
testing.
Phase I consisted of rig tests of 32 configurations of three
distinct combustor concepts to determine which was the best on the basis
of emissions and performance characteristics. The three concepts are
described below.
1. Swirl-can Combustor - Each swirl can is made up of three basic
components: a carburetor, a swirler, and a flame stabilizer. This
combustor is designed such that fuel and air enter the carburetor, mix
when passing through the swirler, and then burn in the wake of the flame
stabilizer. Thirteen swirl-can combustor configurations involving five
major changes were rig tested. These changes were:
1. Changes to the combustor inlet aerodynamics
2. Changes in the carburetor equivalence ratio
3. Changes to flameholding techniques
4. Changes in the fuel injector techniques
5. Addition of liner dilution air
2. Staged Premix Combustor - A description of this type of
combustor is given in Appendix C. During Phase I nine configurations
that included four design modifications were tested. These modifica-
tions involved:
1. diffuser - combustor airflow distribution
2. fuel-air mixture preparation
3. pilot-main burner fuel flow split
4. flame holder design
3. Vorbix - This is also a staged concept designed to provide
relatively long resident times at low power operation to minimize CO and
HC and also to provide rapid burning and quenching of the reactants at
high power to minimize NOx. The pilot burner (used alone at low power)
uses conventional type direct fuel injection (i.e., is not premixed).
Main burner fuel is injected at higher power conditions and is introduced
at the exit of the pilot zone where it is vaporized. Ten configurations
incorporating six design variations were tested during Phase I. These
variations were:
-146-
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1. modifications to FINWALL (cooling) airflow distribution
2. pilot injector type and airflow schedule
3. main burner fuel source density and fuel spray pentration
4. main burner swirler type and orientation and air schedule
5. main burner fuel injector angle
6. pilot-to-main burner fuel flow split
Phase II: Combustor refinement, and optimization: Phase II, which
began in July 1974, was completed in February 1976. During jthis testing
a total of 950,000 dollars was spent on a cost share basis, NASA supply-
ing 63% and PWA, 37%.
This phase of the program emphasized refinement and optimization of
the best Phase I designs. Two combustor designs were selected for
evaluation. These were the vorbix concept and a hybrid concept. A
total of 22 configurations, 15 vorbix and seven hybrid, were rig tested
during this phase. A description of the combustors is provided below.
1. Vorbix - The Phase II vorbix combustor was based on the Phase
I vorbix concept but included the following changes: pilot burner
length increased to lower HC and CO (increased residence times); main
burner length was shortened, liner height reduced, and the throat height
between pilot and main burners was reduced to lower NOx (reduce resid-
ence times).
2. Hybrid - This design attempted to combine the best features of
two Phase I designs. It incorporated a staged premix pilot burner and
a swirl-can mainburner.
Phase III: Combustor-engine testing - Phase III efforts were
initiated in July, 1975 and are nearly completed. A total of 1.5 million
dollars are being spent (excluding hardware costs) on phase III testing,
NASA's share being 1.2 million and PWA's, 300,000; PWA also provided
the necessary hardware. In this phase engine tests of the best Phase
III combustor were conducted, the best having been determined to be the
vorbix concept. Two configurations were tested. Emphasis was put on
validating pollution reductions and on determining the acceleration/
deceleration characteristics of the combustor.
Results and Status
The most successful concept in terms of reducing HC and CO was the
aerating nozzle concept. The HC and CO emissions were reduced by ^ 95
percent and by 90 percent respectively to levels substantially below the
standards. NOx however was reduced only slightly (^5%) with this concept.
The NASA vorbix and hybrid concepts were successful in simultaneously
reducing all three emissions. The vorbix concept was especially effective,
During Phase III testing of the vorbix configuration designated S27E,
compliance with all three gaseous emissions was demonstrated (Ref. 82).
Smoke however was high (SN = 26 compared to standard of 19). This Phase
-147-
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Ill testing is particularly noteworthy since this demonstration was with
a full scale engine. A summary of these results along with results of
the air-assist and aerating nozzle testing is presented in the table
below.
Status
Technology
Concept Category
Standard
Production
Air-assist
Aerating nozzle
combustor
Vorbix combustor*
(S27E)
Hybrid combustor*
1
2
3
3
HC
0.8
5.3
1.4
0.2
0.3
0.7
Development
EPAP Status
CO NOx (Task No.)
Implementation
Date
Jan. 1, 1979
4.3 3.0
14.3 4.9
9.4 4.9 3 (Fig. 58) early 1979**
1.3 4.6 5 (Fig. 59) late 1979
3.3 2.6 3 (Fig. 61) mid 1981
3.3 3.5 2 (Fig. 61) mid 1982
(H-6)
*NASA ECCP Technology
**Includes extra 6 months required for service evaluation
Projections and Prospects
It would seem reasonable to expect Pratt & Whitney Aircraft to
continue developing the vorbix concept since it has shown the capability
of controlling emissions to a level below the standards.
b. JT9D-70
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 1 technology by mid 1979;
- good with category 2 technology by mid 1980
NOx - poor with category 2 technology;
- fair with category 3 technology by mid 1981;
- good with category 3 technology by mid 1983
Thrust = 51,150 Ibs.
BPR =4.9
PR = 24
Combustor type: straight flow annular
Application: 747-400
Certification date: December '74
Number delivered: 1400 for all JT9D models combined as of early '75
Production category after January 1, 1979: III
-148-
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Effort Expended
As stated under part a. (JT9D-7), Pratt & Whitney Aircraft has
spent approximately three million dollars over a five year period
(1971-1975) on their own "in-house" emissions reductions programs for
the JT9D-7 and JT9D-70 engines combined. Data reported (Ref. 12 and 28)
indicates that the air-assist concept, the aerating nozzle concept, the
pretnix concept, and the vorbix concept have been tested for use on the
JT9D-70 engine. This implies that while this engine was not directly
involved in the Clean Combustor program it did benefit substantially
from this program.
Results and Status
Similar to the JT9D-7 engine, the most successful concept in terms
of reducting HC and CO emissions was the aerating nozzle concept. These
emissions were reduced ^ 85 percent and ^ 60 percent respectively. The
resultant levels of these emissions were substantially below the standards.
NOx was reduced approximately 15 percent.
The two staging techniques each reduced NOx by approximately 35
percent relative to the production combustor. However, a 50 percent
reducton is required.
Reported results are tabulated in the status summary below.
Status
Technology
Category
Concept
Standard
Production
Air-assist 1
Aerating nozzle 2
combustor
Premix combustor 3
Vorbix combustor 3
Development
EPAP Status
HC CO NOx (Task No.)
0.8 4.3 3.0
1.2 5.9 5.8
0.3 3.9 5.8 2 (Fig. 58)
0.2 2.3 4.8 3 (Fig. 59)
0.7 3.5 3.8 2 (Fig. 61)
1.0 6.0 3.9 2 (Fig. 61)
Implementation
Date
Jan. 1, 1979
mid 1979**
mid 1980
mid 1982*
mid 1982*
Projections and Prospects
Two other staged combustors should be considered for use in this
engine. These are the ECCP vorbix concept S27E which was demonstrated
in the JT9D-7 engine, and the General Electric double annular concept.
EPA estimates that these would reduce NOx by ^ 40 percent and
*NASA has predicted that premix combustors will not be available until 1985.
The NASA time schedule is considered "normal".
**Includes extra six months for service evaluation.
-149-
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^ 47 percent respectively while controlling HC and CO to a level below
the standards. (Estimates were made in accordance with Appendix E and
Ref. 59). The vorbix would probably be available approximately two
years sooner because of Pratt & Whitney's extensive experience in develop-
ing this concept. Presumably no development of the double annular
concept has been initiated by PWA. A summary of these projections is
presented below.
Projected Status
Concept
Standard
Double annular*
Vorbix*
S27E
Technology
Category HC
Development
EPAP Status
CO NOx (Task No.)
3
3
0.8
0.3
0.3
4.3 3.0
3.2 3.1 1 (Fig. 61)
3.2 3.5 2 (Fig. 61)
Implementation
Date
Jan. 1, 1979
mid 1983
mid 1982
*NASA technology
c.
JT10D
Profile
Prospects for meeting 1979 standards:
HC, CO - good
NOx - good
Thrust = 23,000 Ibs.
BPR = 5.35
PR = 28
Combustor type: straight flow annular
Application: proposed for 707, DC-8, DC-9, 7X7
Certification date: projected for 1979
Number delivered: none - engine still in development
Production rate: none - engine still in development
Production category after January 1, 1979: IV
Effort Expended
Data submitted to the EPA (Ref. 30) indicates that approximately
583,000 dollars was spent on low emissions efforts for the JT10D engine
during the years 1973-1975. The JT10D is in the pre-developtnent demon-
strator phase, and the expenditure noted above is that portion of the
total combustor development cost estimated by Pratt & Whitney Aircraft
(Ref. 87) to be attributed to low emissions efforts.
-150-
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Results and Status
No emissions data has been reported for this engine.
Projections and Prospects
The JT10D is a newly designed engine (first tested in August 1974)
based on JT9D technology. It is being developed for minimum fuel consump-
tion and low emissions and should be expected to comply with the standards
since much of the design work was done after the standards were promulgated.
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2. General Electric
a. CF6-50C
Profile
Prospects for meeting the 1979 standards:
HC, CO - fair with category 1 technology before 1979
NOx - poor with any known control (category 1-3)
Thrust = 49,900 Ibs.
BPR = 4.4
PR = 29.5
Combustor type: straight flow annular
Application: DC-10 Series 30, A300-B
Certification date: c. 1972
Number delivered: more than 400 as of January 75
Production rate: unknown
Production category after January 1, 1979: III
Effort Expended
The bulk of GE's work on this engine since January, 1973 has been
in support of their contractual effort (cost-sharing) with NASA's
Experimental Clean Combustor Program (ECCP). This major program, involv-
ing approximately 3.5 million dollars,(^ 80% NASA and ^ 20% GE plus GE
supplying the necessary hardware required for testing), is divided into
three phases, the first being a preliminary screening phase of a number
of potential concepts, the second being a development phase wherein the
best two concepts were further refined, and the third being an engine
demonstration phase of the best concept. As the program goal was the
development of technology with the capacity to meet the 1979 EPA standards,
it was recognized early that only category 3 concepts would be acceptable.
Hence, the major portion of the effort was directed at such technology.
The screening phase (Phase I) included rig investigation of the
following concepts (reference 1):
(1) Swirl can modular dome
This is essentially a mix of con.tro.1 approaches, all wrapped up
into a single, but extensive nozzle modification: premix (modest degree),
airblast, and lean primary. Seventeen versions were investigated in a
parametric study. (category 2).
(2) Lean dome single annular combustor
This is basically a conventional combustor with the airflow redistri-
buted to provide a very lean primary zone equivalence ratio. The addi-
tional air was admitted through the new carbureting swirl can nozzles,
holes in the dome, and restrictions in the dilution parts of the liner
downstream, as necessary. Four configurations, with and without simulated
bleed at idle, were examined. (category 2-3).
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(3) Double annular combustor -
This concept marks a major departure from conventional combustor
design as fuel staging is involved. As the name suggests, there are, in
effect, two combustors within one, one designed for high power, but low
NOx, operation (lean primary), and the other for near stoichiometric
operation at low power to consume HC and CO better. Six versions were
examined. (category 3).
(4) Radial/Axial combustor -
This is basically just another version of the fuel staging concept.
Seven configurations were investigated. (category 3).
In the development phase (Phase II) of the contract (reference 75),
the two fuel staging concepts (double annular and radial/axial) were
perfected further and subjected to various performance tests (non-
emissions) such as altitude relight, carbon deposition, etc. A total of
66 rig (full and sector) tests were performed through this phase to
investigate the emissions improvements and performance of different
modifications.
In the engine testing phase (Phase III) now underway, one concept
(the best version of the double annular) is being tested in an engine
for emissions and other engine related performance criteria (accelera-
tion, etc.).
Prior to 1973 most of the General Electric effort evidently was
directed towards establishment of baseline emissions and understanding
of the basic physical and chemical phenomena. Additional in-house work
(for which cost information is not available) has been done up to the
present in an effort to establish simpler alternatives to the approaches
pursued in the NASA program (references 8 and 73). NOx control was
explored by techniques of quick quenching, and lean primary, all category
2 concepts. HC and CO control was explored by bleed and sector burning
at idle, both category 1 concepts. The amount of effort in exploring
these category 1 concepts is small in view of their simplicity. Testing
normally involves a parametric investigation of the control. In these
category 1 cases, these were amount of bleed and the size and location
of the sectors. Some additional support testing'is necessary in some
instances, in particular with water injection and sector burning, to
investigate potential durability problems. To EPA's knowledge, GE has
not continued with these studies.
GE is also involved in other government sponsored programs (see
Appendix A) at present directed towards high altitude cruise emissions
(NOx) and very low HC, and CO emissions (i.e., directed at the 1981
standards).
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Results and Status
The most significant results of the above testing are tabulated
below.
Status
Technology
Category HC
0.8
4.3
1 0.3
Concept
Standard
Production
Production 1
combustor plus
sector burn at
idle
Single annular 2-3(est)
lean dome
Double annular* 3
Radial/axial* 3
10.4
0.3
0.9
Development
EPAP Status
CO NOx (Task No.)
4.3 3.0
10.8 7.7
5.3 7.8 3 (Fig..58)
12.7 7.2 2 (Fig. 61)
Implementation
Date
Jan. 1, 1979
late 1978
***
3.0 4.3 3 (Fig. 61) mid 1981**
9.6 4.3 2 (Fig. 61) mid 1982**
*ECCP Technology
**NASA predicts it may take until 1985 to implement staged
combustor techniques. The NASA time schedule is considered
"normal".
***Not likely to comply under any circumstance.
Projection and Prospects
The data above indicate that the sector burn concept is successful
in reducing HC and CO emissions substantially but only HC is reduced to
a level below the standards. This is due in part to the low idle power
(3.4% of rated power) of this engine and in part to the short combustor
(short residence time). If the engine idle speed were increased to say
6% of rated power, the CO emissions should be reduced to an EPAP value
of approximately 3.4 and HC emissions would be reduced to an EPAP value
of approximately 0.2. The effect on NOX should be negligible. Alterna-
tively, increased bleed at idle would provide the same effect without
suffering the adverse consequences of a higher thrust level. Both
solutions result in a fuel consumption penalty, however, at idle.
The single annular lean dome combustor is a considerable failure
and emphasizes the problems to be encountered when the single primary
zoned combustor (i.e., conventional) is pushed beyond its technology
capabilities. Although further development would probably yield some
improvement, it is apparently futile to attempt to reduce all three
pollutants simultaneously with such a configuration.
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The double annular concept produces emissions of HC and CO below
the standards. NOx is reduced by 45 percent but is still above the
standards. A problem with respect to NOx emissions from the double
annular combustor is that it has a higher flame temperature than the
baseline combustor at approach. This is a result of a high fuel air
ratio for pilot-only operation during this mode. In spite of this, the
NOx level shown for the double annular concept is judged by NASA to be
nearly optimum for the level of technology evaluated. The radial/axial
concept is also capable of simultaneously reducing all three emissions,
but none are reduced to the levels required by the standards. According
to NASA both the double annular and the radial/axial concepts represent
major steps for reducing engine emissions without compromising performance.
Either staging concept is faced with certain, as yet unreasolved,
development difficulties, most notably that of the fuel control. Approach
power requires pilot-only operation in order to retain low CO charac-
teristics. However, the fuel flow rate at cruise (high altitude) is
comparable to that at approach; yet, there, the fuel must be distributed
between the two stages and, in fact, more must be supplied to the secon-
dary than to the pilot. A control system then must be able to distin-
guish between these two flight regimes despite the similar total fuel
flows. Either altitude or speed sensing must be linked into the fuel
control. Another problem is that of the transition between pilot-only
(up to approach) and staged (beyond approach) fuel flows. The transi-
tion must occur without thrust perturbations or flame blowout.
b. CF6-6
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with cateogry 1 technology by early 1979
NOx - poor with any known control (category 1-3)
Thrust = 38,900 Ibs.
BPR =5.9
PR = 24.7
Combustor type: straight flow annular
Application: DC-10 Series 10
Certification date: September '70
Number delivered: as of January '75 more than 350
Production rate: unknown
Production category after January 1, 1979: III
Effort Expended
Little actual work on this engine has been expended to EPA's
knowledge (references 6, 7, 8, and 47). Although this engine uses a
different combustor from that in the CF6-50 and the operating conditions
are also different, there is sufficient similarity that the results of
concept selection in the CF6-50 program can be expected to be applicable
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here also. This is an acceptable and economic approach in this instance.
Consequently, the only testing done on this engine to date has been to
explore baseline emissions and the potential of compressor bleed (category
1 technology).
At this point in time, however, it is suprising that the successful
technology concepts identified in the CF6-50 work have not been introduced
into the CF6-6 program. Specifically, engine data for modest technology
(category 1 or 2) application (e.g., sector burning, combined if necessary,
with compressor bleed or increased throttle) might have been available
by now. Also, in view of the initial successes of the double annular
combustor in the CF6-50 (reference 75) and in view of the long development
leadtimes, one might have expected to see the initial steps taken by
this time to incorporate this combustor into the -6.
Results and Status
The limited actual engine data available are presented below:
Status
Concept
Standard
Production
Compressor bleed
Technology
Category HC
0.8
3.4
1.7
EPAP
CO NOx
4.3 3.0
10.0 7.2
8.0 7.4
Development
Status
(Task No.)
4 (Fig. 58)
Implementation
Date
Jan. 1, 1979
mid 1978
Projections and Prospects
The low emissions concepts being evaluated for the CF6-50C engine
are considered to be applicable to the CF6-6D engine and the results are
expected to be similar in terms of emissions reductions. Assuming this,
the following table represents what is potentially available.
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Projected Status
Development
Technology EPAP Status Implementation
Concept Category HC CO NOx (Task No.) Date
Standard 0.8 4.3 3.0 Jan. 1, 1979
Production 1 0.2 4.9 7.3 2 (Fig. 58) early 1979
combustor plus
sector burn at
idle
Production 1 0.1 3.9 6.7 2 (Fig. 58) early 1979
combustor with
sector burning
and compressor
bleed at idle
Double annular 3 0.2 2.84.0 2 (Fig. 61) mid 1982
Instead of combining the two category 1 concepts, sector burning
and compressor bleed, to achieve sufficiently low HC and CO levels, it
is possible also to simply increase the engine power level at idle
instead of bleeding the compressor although this would be less desirable
because of the added thrust. Either way there is a fuel penalty at
idle.
Note that the double annular combustor is not expected to meet the
NOx standard although it does very well. The present development diffi-
culties attendant with this combustor are described in the Projections
and Prospects Section of the CF6-50 discussion.
c. CFM56
Profile
Prospects for meeting the 1979 standards:
HC, CO - fair with category 2 technology by early 1979
- good with category 3 technology by mid 1982
NOx - good with category 3 technology by mid 1982
Thrust = 22,000 Ibs.
BPR =6.0
PR = 25
Combustor type: straight flow annular
Application: proposed for BAC-111, DC-9, 727, A300, 707, DC-8
Certification date: expected late '77 or early "78
Number delivered: none - not yet in production
Production rate: none - not yet in production
Production category after January 1, 1979: IV
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Effort Expended
This engine is in a little different situation than the other GE T2
class engines in that it is not yet certified. This means that the
present combustor is not so inextricably bound to the engine that change
is difficult. Furhermore, as the engine has no in-service history
behind it, the stringent durability requirements of an older engine are
not present. Emissions performance data is scant for this engine (ref-
erences 8, 38, 39, 47, 73, and 76), for just as the effort on the CF6-6
has been small, pending the outcome of the various CF6-50 studies, so
has the effort here. Nevertheless, sector burning has been investigated
in both the baseline combustor (category 1) and a modified combustor
(category 2). The actual extent of this testing (i.e., the degree of
optimization), however, is unknown to the EPA at this time.
Results and Status
The limited actual engine data are presented below:
Status
Concept
Standard
PFRT (Baseline)
Baseline combus-
tor plus sector
burn at idle
Modified PV com-
bustor plus sec-
tor burn at idle
Technology
Category HC
0.8
1.7
0.9
Development
EPAP Status
CO NOx (Task No.)
4.3 3.0
12.8 4.7
10.9 4.8 4 (Fig. 58)
Implementation
Date
Jan. 1, 1979
mid 1978
0.2
9.7 4.8 4 (Fig. 59) early 1979*
*Service evaluation excluded, replaced by entry into service
experience.
Projections and Prospects
This engine produced more CO than is expected from a modern high
bypass ratio engine. This is due to its relatively short combustor and
subsequently short residence time. Efforts to improve combustion efficiency
have resulted in reducing only HC to a level below the standards. The
level of CO listed is based on data submitted to the EPA by General
Electric (Ref. 39). However, at the EPA Hearing on Aircraft held January
27-28, 1976, General Electric recommended that the CO level for newly
manufactured T2 class engines should be set at an EPAP of 5.1 (Ref. 47
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and 48). On the other hand, according to the 1975-1976 "Jane's All The
World's Aircraft" (Ref. 69), this engine was designed to meet the 1979
gaseous emissions standards. Assuming that Jane's obtained its informa-
tion accurately from the manufacturer, it can be surmised that further
CO and NOx improvement can be expected from this engine, and that the CO
emissions indicated for this engine do not represent optimal technology.
Additional improvement in the CO emissions can most likely be
obtained by modest means through the use of compressor bleed in addition
to sector burning although with the combustors now available (baseline
or modified PV) the CO standard would most likely still not be met.
Use of the double annular combustor (technology category 3) in this
engine would probably control all three gaseous emissions to levels
below the standard as is indicated in the table below. The fact that an
improvement is projected for CO emissions through the application of a
category 3 concept (combined with the best compatible category 2 concepts
in the pilot chamber to control HC and CO) suggests that the two combus-
tors on which the sector burning was tried have not been fully optimized
and further CO improvements can be expected with category 2 technology.
This is probably recognized by GE and is the basis for their CO standard
recommendation referred to above. The projected status of the CFM56
engine is summarized in the table presented below.
Projected Status
Development
Technology EPAP Status Implementation
Concept Category HC CO NOx (Task No.) Date
Standard "0.8 4.3 3.0 Jan. 1, 1979
Modified PV com- 2 0.1 7.84.8 3 (Fig. 59) mid 1980
bustor plus sector
burn and compressor
bleed
Double annular* 3 0.4 3.9 2.7 2 (Fig. 61) mid 1982
*ECCP technology
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3. Rolls Royce
As with the Tl engines the limited submission of emissions and
engine performance data by Rolls Royce as of the time of this report did
not permit the EPA to make a satisfactory assessment of the status of
Rolls Royce. The data tabulated for each engine in the respective
status tables represents the only data available for each Rolls Royce
engine. The reduced emissions technology concepts are unknown in specifics
but are either category 1 or 2 technology. Further, the EPA is not
sufficiently familiar with the Rolls Royce test programs to make accurate
judgments with respect to combustor development times. Therefore the
estimates listed are not as precise as for other engines.
a. RB211
The discussion of the emissions programs of the RB211-22B and the
RB211-524 has been combined because of basic design similarities and
hence emissions characteristics. Furthermore, the lack of detailed
information precludes any meaningful discussion at a level which would
distinguish between the two.
Profile RB211-22B
Prospects for meeting the 1979 standards:
HC, CO - fair with category 1-2 technology (by early 1980)
NOx - poor with category 1-2 technology;
- good with cateogry 3 technology (status unknown)
Thrust = 42,000 Ibs.
BRP =5.0
PR = 25
Combustor type: straight flow annular
Application: L-1011-1
Certification date: April '73
Number delivered: 450 including modified RB211-22C engines
as of April '75
Production rate: unknown
Production category after January 1, 1979: III
Profile - RB211-524
Prospects for meeting the 1979 standards:
HC, CO - good with category 1-2 technology (by early 1980)
NOx - poor with category 1-2 technology;
- good with category 3 technology by (status unknown)
Thrust = 50,000 Ibs.
BPR =5.0
PR = 25
Combustor type: straight flow annular
Application: proposed for B747, A300B
Certification date: c. 1975
Number delivered: unknown
Production rate: unknown
Production category after 1, 1979: III
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Effort Expended
Data submitted by Rolls Royce (Ref. 44 and Ref. 45) indicates that
programs to reduce emissions began in 1973 after the publication of the
original NPRM in December 1972. Details about the RB211-22B program were
not available to EPA at the time of this report.
Based on the limited submittals, it is known that rig testing of
several basic reduced emissions concepts has been conducted in order to
determine their usefulness on RB211 engines. The concepts tested are
(from Ref. 81) 1) fuel preparation and injector modifications (category
1), 2) compressor bleeds (category 1), 3) sector burning (category 1),
4) operating speed (category 1), 5) staging (category 3), and 6) premix
with staging (category 3). It is not known how many versions of each
were tested and whether or not any parametric studies were made. Results
from this low emissions testing have not been reported, and thus the
effectiveness of each in reducing emissions cannot be evaluated.
The only level-of-effort information (from Ref. 45) indicates that
during 1975, 63 percent of the RB211 combustor rig development funding
was allocated to the development of pollution control combustors. No
information quantifying the total dollars or manhours used for emissions
reduction programs in 1975 or any other year is available.
Results and Status
These engines produce much higher HC and CO emissions than is
expected from a modern low SFC T2 class engine. The only low emission
data available, tabulated in the status table below, indicates that all
emissions were reduced including NOx surprisingly; only HC was reduced
to a level below the standards in the -22B, while for the -524, both HC
and CO were reduced to levels below the standards due mostly to the more
favorable operating pressure at idle in that model than in the -22B.
Status (RB211-22B)
Concept
Standard
Production
Unknown
Technology
Category HC
0.8
14.0
1 or 2 0.7
Development
EPAP Status
CO NOx (Tksk No.)
4.3 3.0
20.0 7.3
5.6 6.8 4 (Fig. 59)
Implementation
Date
Jan. 1, 1979
early 1980
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Status (RB211-524)
Concept
Standard
Production
Unknown
Technology
Category HC
0.8
8.3
1 or 2 0.5
Development
EPAP Status
CO NOx (Task No.)
4.3 3.0
11.9 8.9
3.7 7.1 4 (Fig. 59)
Implementation
Date
Jan. 1, 1979
early 1980
Projections and Prospects
Insufficient data prevents any useful projections from being made.
However, because of the design and operational similarities between this
engine and the JT9D or CF6 engines, any of the low emissions combustor
techniques discussed previously for the JT9D or CF6 should be applicable
to this engine, especially the aerating nozzle concept, the double
annular concept and the vorbix concept. Therefore, it is expected that
the emissions levels would be comparable to those of the JT9D engine
with an optimized combustor.
b. Spey
The discussion of the emissions programs and performance of the
Spey 511 and the Spey 555 has been combined because of basic design
similarities and hence emissions characteristics.
Profile - Spey 511
Prospects for meeting the 1979 standards:
HC, CO - poor with any technology
NOx - poor with category 1 or 2 technology;
- unknown with category 3 technology
Thrust = 11,400 Ibs.
BPR = 0.64
PR = 18.9
Combustor type: multiple cans
Application: Grumman Gulfstream II
Certification date: c. 1962
Number delivered: unknown
Production rate: 50 per year for all Spey models
Production category after January 1, 1979: II, III
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Profile - Spey 555
Prospects for meeting the 1979 standards:
HC, CO - poor with any technology
NOx - poor with category 1 or 2 technology
- unknown with category 3 technology
Thrust = 9850 Ibs.
BPR = 1.0
PR = 15
Combustor type: multiple cans
Application: F-28
Certification date: c. 1962
Number delivered: unknown
Production rate: 50 per year for all Spey models
Production category after January 1, 1979: II, III
Effort Expended
Except for the EPAP data presented below, little is known of the
extent of the program for improving the Spey's emissions. Neither the
specifics of the concepts tested nor the extent of testing are known to
the EPA for this family. The manufacturer has claimed to have implemented
an "intensive programme" (reference 44) and certainly sizeable reductions
in HC and CO emissions have been realized, yet because of the basic
combustor design of the production engine and the relatively high SFC
(reflecting its age-certification around 1962), the family is a long way
from compliance with even the HC and CO standards. In summary, little
was known at the time of this report as to the extent of the program for
improving the Spey's emissions.
Nonetheless, the following remarks can be made. First, Rolls Royce
has evidently concluded on the basis of its own analysis that the engine
cannot meet the standards with any technological effort whose development
cost would be recoverable within the estimated remaining production of
the engine. Uneconomical technology includes, apparently, any category
3 concepts and any further category 2 effort. The concepts which have
been tried are (reference 81):
(1) engine speed (category 1)
(2) combustor airflow redesign (category 2)
The engine speed control has been engine tested, the other concept, only
rig tested. The extent of these investigations and the degree of
refinement pursued is unknown. In addition, Rolls Royce evidently
planned to examine in 1975 (reference 81) certain advanced concepts
(variable geometry and staged combustion). However, no results have
been made available to the EPA.
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On the other hand, in response to EPA's inquiry at the January,
1976 hearing, Rolls Royce reported (reference 45) that 78% of the total
combustor rig development funding for the Spey was allocated for pollu-
tion control. The true extent of the commitment represented by this
expenditure, however, cannot be appreciated without knowledge of the
actual monies spent. In view of the maturity of the engine and the
already established excellent durability record of the Spey combustor,
it is understandable that most of the combustor development still
going on is related to emissions improvements.
Results and Status
The only results known to the EPA regarding the Spey family are:
Status - Spey 511
Concept
Standard
Production
Unknown
Technology
Category HC
0.8
35.7
1 or 2 5.0
Development
EPAP Status
CO NOx (Task No.)
4.3 3.0
35.7 7.8
20.7 7.4 unknown
Implementation
Date
Jan. 1, 1979
unknown
Concept
Standard
Production
Unknown
Technology
Category HC
Status - Spey 555
Development
EPAP Status
CO NOx (Task No.)
1 or 2
0.8
46.2
6.0
4.3 3.0
63.2 4.3
23.4 4.2
unknown
Implementation
Date
Jan. 1, 1979
unknown
No estimates of the implementation dates are possible by using the
leadtime analysis in section IV as no knowledge of the development
status of the control concepts is known. The manufacturer has recom-
mended exemption for the Spey family, so no recommended time applicable
for this family is available.
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Projections and Prospects
It is difficult to discuss alternatives in view of the lack of
knowledge of what has been attempted and its level of success. The
following general statements are pertinent, however. The Spey engine
family is very similar in function, design, and performance to the JT8D
family of Pratt and Whitney (compare profiles) except that the Speys are
smaller (50 to 80% the thrust levels of the JTSDs). Both families have
comparable bypass ratios, pressure ratios, turbine inlet temperatures,
and burner design, all of which reflect the fact that both engines are
of the same generation (early 1960s). The air loading of the Spey is
slightly higher (worse) than that of the JT8D (air loading parameter of
0.68 vs. 0.53 for two representative models), but not enough to explain
the relative success of the JT8D over the Spey. Surface-to-volume
effects in the combustors cannot be evaluated as that information is
not available for the Spey.
The prospects for this engine do not appear encouraging on the
basis of what is known, but then again, the causes of the discouraging
data cannot be appreciated. Because of its similarity with the JT8D,
the emissions performance of the Spey should be comparable with the
JT8D, but this has not been realized.
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C. Class T3 engine
Class T3 engines means all aircraft gas turbine engines of the JT3D
model family. This special class was created in order to implement the
special smoke retrofit standard for these engines. The gaseous emissions
standards and compliance date are however, the same as those for the T2
class engines. The standard is listed below for convenience.
Standard
EPAP
(Ibs pollutant per 1000 Ibf-hours over the cycle)
HC CO NOx
0.8 4.3 3.0
1. Pratt & Whitney Aircraft
a. JT3D
Profile
Prospect for meeting the 1979 standards:
HC, CO - poor
NOx - poor
Thrust = 19000 Ibs. (-7 model)
BPR =1.4
PR = 13.5
Combustor type: multiple cans
Application: 707 - 320c, DC-8
Certification date: c. 1960
Number delivered: 8300 for all models combined as of the
end of 1974
Production rate: unknown
Production category after January 1, 1979: I
Effort Expended
Reported data (Ref. 30) indicates that approximately 2.2 million
dollars have been spent on emissions reductions on the JT3D/TF33* engines,
Efforts to develop a low emission combustor have been mainly devoted to
reducing smoke with an aerating nozzle combustor. There has apparently
been no attempt to comply with the gaseous emission standards.
Results and Status
By utlizing the aerating nozzle concept some reduction in HC and CO
has been made at the expense of NOx. The emissions data from this
reduced smoke combustor are presented below along with the emissions
data from the production combustor.
*The TF33 is a military engine. Presumably some of the low smoke effort
has been financed by the Air Force. It is not known how much, however.
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Status
Concept
Standard
Production
(-1 model)
Technology v
Category HC
0.8
34.2
Aerating nozzle
combustor
(-7 model)
18.0
Development
EPAP Status
CO NOx (Task No.)
4.3 3.0
40.8 3.8
Implementation
Date
Jan. 1, 1979
26.2 5.6 8 (Fig. 59) early 1978
Projections and Prospects
It is doubtful that further developments leading to lower emissions
will be made because the aerating nozzle combustor is already being used
in the smoke retrofit program, and also because this engine is nearing
the end of its production. Development of still another low emissions
combustor would probably not be economically justified because of limited
number of engines expected to be produced after January 1, 1979.
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D. Class T4 Engine
Class T4 engines means all aircraft gas turbine engines of the JT8D
model family. As with the T3 Class, this class was created for the
implementation of a special smoke retrofit standard and it must also
comply to the T2 Class emissions standards and implementation date. The
gaseous emissions standard is tabulated below for convenience.
Standard
EPAP
(Ibs pollutant per 1000 Ibf-hours over the cycle)
HC CO NOx
0.8 4.3 3.0
1. Pratt & Whitney Aircraft
a. JT8D
Profile
Prospects for meeting the 1979 standards:
HC, CO - fair with category 2 technology by mid 1979
NOx - poor with category 2 or 3 technology
Thrust = 16,000 Ibs. (-17 model)
BPR =0.99
PR = 16.9 . -
Combustor type: multiple cans
Application: 727, 737, DC-9
Certification date: c. 1962
Number delivered: 7200 for all models combined as of the
end of 1974
Production rate: unknown
Production category after January 1, 1979: II, III
Effort Expended
The JT8D engine was selected as one of the base engines in the NASA
Pollution Reduction Technology Program (PRTP). The JT8D program was
initiated in August 1974, with Phase I of an anticipated three phases
which ended in December 1975. During this program a total of 846,000
dollars (Ref. 87) was spent on a cost share basis, 37% PWA funds and 63%
NASA funds. This program was divided into three basic elements, each
investigating successively more complex combustor concepts. Element I
consisted of investigating modifications to the existing JT8D combustion
system (basically category 1 and 2 concepts). These were: airblast
fuel nozzles, airflow distribution, and fuel-air carburetion. A total
of six configurations were tested. Element II consisted of testing nine
configurations of the vorbix combustor (category 3 technology). This
concept has more potential for meeting the standards than Element I
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concepts, but it is more complex and would be more difficult to adapt to
an operational engine. Element III consisted of testing five configura-
itons of the premix/prevaporization combustor concept (category 3-4
technology). This concept represented the greatest difficulty of those
considered in this program for adapting to operational engines.. However,
it also has the highest potential for meeting the standards.
Available data from Reference 30 indicates that Pratt & Whitney
Aircraft also spent an additional 1.1 million dollars on JT8D reduced
smoke and gaseous emissions efforts during the time period 1971-1975.
Most of this effort seems to have been devoted to testing of the aerat-
ing nozzle concept. Some additional testing has also been done with the
external air-assist concept.
Results and Status
Use of the aerating nozzle combustor concept resulted in an 80
percent reduction in HC, a 70 percent reduction in CO, and a 15 percent
reduction in NOx. However, only HC was reduced enough to meet the
standards. Reductions to the levels indicated are significant because
this engine represents older technology and hence has a relatively high
SFC. However, the JT8D-17 aerating nozzle HC and CO emissions indices
reported are among the lowest from any combustor of technology rating 2
complexity. Further major El reductions would not be expected in the
near future.
The aerating nozzle technique, however, does not adequately control NOx.
Some type of staged combustor such as the vorbix or a premix/prevaporiza-
tion combustor would be required. The vorbix concept that was judged
best by NASA (Ref. 34) reduced NOx by 48 percent. This percentage
reduction in NOx was the best of any of the combustors tested in the
ECCP or PRTP (NASA programs.) NOx reductions to lower levels would
therefore seem unlikely in the near future. Additional development is
still needed to achieve acceptable exit temperature pattern factors.
Further, this concept has an altitude relight deficiency which would
also need to be resolved possibly at the expense of the emissions control.
Available test results are summarized in the status table below.
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Status (-17 model)
Technology
Category
Concept
Standard
Production
Air-assist 1
Aerating nozzles 2
Airflow redistri- 2
bution*
Vorbix* 3
Premix/ 3-4
prevaporized*
Development
EPAP Status
HC C0> NOx (Task No.)
0.8 4.3 3.0
3.1 12.8 6.8
1.2 10.5 7.1 3 (Fig. 58)
0.8 5.0 7.1 5 (Fig. 59)
1.5 6.8 7.4 2 (Fig. 59)
0.2 9.0 4.3 2 (Fig. 61)
0.4 14.3 4.6 2 (Fig. 61)
Implementation
Date
Jan. 1, 1978
early 1979**
mid 1979
early 1981
mid 1982
mid 1982
*NASA technology
**Includes extra 6 months for service evaluation
Projections and Prospects
As noted above the JT8D represents an older technology low bypass
ratio engine and hence has a relatively high SFC. The aerating nozzle
concept in this engine produces among the lowest levels of HC and CO
emissions indices, but only marginally meets the HC standard. Therefore
reductions to levels below those of the aerating nozzle levels listed in
the table above are not likely in the near future. Further, NOx reductions
greater than the 48 percent reduction noted above (achieved as a ..-result
of using the vorbix combustor) are also unlikely in the near future.
These two concepts are not mutually exclusive and could be used together
in a single combustor to control all three emissions. However, it is
doubtful that HC and CO would be as low as those for the aerating nozzle
concept listed, or that NOx would be as low as that for the vorbix
combustor listed. The exact level of emissions for such a combustor
cannot be predicted.
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E. P2 Class Engines
The P2 Class covers all aircraft turboprop engines (see FR Vol. 38,
N. 136, July 17, 1973, p. 19091). These engines are used on medium to
heavy general aviation aircraft with some used on commercial air trans-
ports. The gaseous emissions listed below for these engines are presently
scheduled to go into effect on January 1, 1979.
Standard
EPAP
(Ibs pollutant per 1000 horsepower-hours over the cycle)
HC C£ NOx
4.9 26.8 12.9
1. AiResearch
a. TPE331
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 1 technology by Jan. 1, 1979
NOx - already met
Equivalent shaft horsepower = 900 (-3 model)
PR = 10.4
Combustor type: reverse flow annular
Application: Fairchild Turbo Porter, Short Skyvan, NAR
Turbo Commander 690, etc.
Certification date: February '65 (-1 model)
Number delivered: unknown
Production rate: more than 60 per month for all models
combined as of January '75
Production category after January 1, 1979: III
Effort Expended
AiResearch has conducted low emissions testing on the TPE-331
engine since 1973. The only quantitative financial data available is
that from the EPA ambient effects contract (Ref. 83). Approximately
16,000 dollars were spent on developing the concept of rescheduling the
fuel flow to operate on only primary atomizers during the taxi-idle
mode. Emissions can be controlled to a level below the standards on the
TPE-331 engine with this concept.
Other concepts that have been investigated for use on the TPE-331
engine are airblast fuel nozzles (2 configurations) and taxi-idle power
setting (parametric study). Data quantifying the amount of money and
manhours spent on these studies is not available.
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Results and Status
As noted above, this engine has demonstrated the capability to meet
the standards by the required date with category 1 technology. This was
accomplished by rescheduling the fuel flow to operate on only the primary
atomizers during the taxi-idle mode. The improved fuel atomization
resulted in an increase in combustion efficiency which was sufficient to
achieve the required HC and CO emission reductions. NOx was not a
problem as the current production engine levels are sufficiently below
the standards to offset the slight increase (due to increased combustion
efficiency). The available results are summarized in the table below.
Status (-3 model)
Concept
Standard
Production
Taxi-idle*
power setting
Primary only
at taxi-idle
Airblast
Technology
Category HC
4.9
55
Development
EPAP Status
CO NOx (Task No.)
26.8 12.9
44 8.0
Implementation
Date
Jan. 1, 1979
2.3 15.0 15.7 3 (Fig. 58) late 1978
3.6 12.8 8.9 5 (Fig. 58) Jan. 1, 1979
33
35 9.2 2 (Fig. 60) mid 1980
Projections and Prospects
No further reductions in emissions are likely for this engine since
AiResearch plans to implement the primaries only at idle concept. As
noted above this concept controls emissions to a level below the standards.
This concept is expected to be used in newly manufactured engines after
January 1, 1979.
*Best combination demonstrated was at 7 percent rated power at 100 percent
rated speed.
-------
2. Pratt & Whitney Aircraft of. Canada
a. PT6A-27
Profile
Propsects for meeting the 1979 standards:
HC, CO - good with category 2 technology by late 1979
NOx - already met
Equivalent shaft horsepower = 715
PR = 6.5
Combustor type: reverse flow annular
Application: Beech 99A, DHC Twin Otter, etc.
Certification date: c. November '67
Number delivered: 1000 as of March '75
Production rate: unknown
Production category after January 1, 1979: III
Effort Expended
Available data (Ref. 20) indicates that low emissions testing began
in 1973. Fiscal data is available for that year only, during which approx-
imately 87,000 dollars were spent for low emissions research. Low
emissions data reported indicates that only airflow redistribution
techniques (category 2 technology) have been tested on this engine. At
least five different engine tests of three different combustors were
made to optimize the primary zone fuel-air ratio, delay quenching, and
to improve mixing.
Results and Status
_1 The PT6A27 has a relatively high surface-to-volume ratio (12.35
ft ), and a relatively low compressor discharge pressure (1.8 atm) and
temperature (660 °R) at idle, all of which contribute to high HC and CO
emissions. High surface-to- volume ratio combustors tend to have higher
HC and CO emissions because of wall quenching effects. The low compressor
discharge pressure compounds the problem by causing slow reaction rates.
Emissions levels below the standards have been demonstrated by this
engine using a rich primary, delayed quench combustor. However, this
combustor has a problem of very hot walls which will have to be resolved
before it is flightworthy. Available test results are presented in the
table below.
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Status
Concept
Standard
Production
Lean primary,
delayed quench
(avg. of 2)
Rich primary,
delayed quench
(avg. of 2)
Improved primary 2
zone mixing,
delayed quench
Technology
Category HC
4.9
52
2 4.9
Development
EPAP Status
CO NOx (Task No.)
26.8 12.9
66 7
22.9 No . 3 (Fig. 60)
Data
Implementation
Date
Jan. 1, 1979
late 1979
2.0 19.0 8.1 3 (Fig. 60) late 1979
4.8 31.6 No 3 (Fig. 60)
Data
late 1979
Projections and Prospects
This engine has demonstrated the capability of meeting the standards
with the rich primary zone, delayed quench combustor. It is expected
therefore that Pratt & Whitney of Canada will pursue development of this
combustor. As noted above the problem of very hot combustor wall will
have to be resolved. As a result emissions of HC and CO may increase,
but it is not expected that the increase will be so large that the
standards will be exceeded (see PT6A-41 Status).
b. PT6A-41
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 2 technology by late 1979
NOx - already met
Equivalent shaft horsepower = 903
PR = 8.0
Combustor type: reverse flow annular
Application: Beech Super King Air 200
Certification date: c. 1973
Number delivered: more than 110 as of February '75
Production rate: unknown
Production category after January 1, 1979: III
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Effort Expended
Data reported in Reference 20 indicates that low emissions testing
on the PT6A-41 began in 1973. During 1973 approximately 102,000 dollars
were spent on low emissions testing. No fiscal data is available for
any other time period.
Low emissions data reported for this engine also indicates that
only airflow redistribution techniques have been tested on this engine.
Similar to the PT6A-27 case, four tests were conducted to optimize the
primary zone fuel-air ratio, delayed quenching, and to improve mixing.
Results and Status
As with the PT6A-27 emissions levels below the standards have been
demonstrated with the rich primary zone, delayed quench combustor. The
hot combustor wall problem exists with this engine also, and will have
to be resolved. Available test results are presented in the status
table below.
Status
Concept
Standard
Production
Improved wall
cooling
Rich primary, 2
delayed quench
Rich primary, 2
delayed quench
plus improved
cooling
Hotter interne- 2
diate zone
Projections and Prospects
Technology
Category HC
4.9
>54
2 30.4
2 1.8
2 4.7
EPAP
CO NOx
Development
Status
(Task No.)
26.8 12.9
81 7.2
44.9 7.5 3 (Fig. 60)
16.6 8.3 3 (Fig. 60)
23.2 8.7 3 (Fig. 60)
Implementation
Date
Jan. 1, 1979
late 1979
late 1979
late 1979
3.8 21.4 8.6 3 (Fig. 60) late 1979
Since this engine has demonstrated the capability to meet the
standards with the rich primary, delayed quench combustor it is antici-
pated that development of this combustor will be pursued. The data
above indicates that improved cooling has been added to this combustor
without exceeding the standards. Further development is still needed,
however, as hot spots were still noted.
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3. Avco Lycoming
a. LTP-101
Profile
Prospects for meeting the 1979 standards:
HC, CO - fair with category 1-2 technology by late 1977
- good with category 1-2 technology by early 1980
NOx - good with category 1-2 technology by late 1977
Equivalent shaft horsepower = 620
PR = 8.4
Combustor type: reverse flow annular
Application: Piaggio P168, Britten-Norman Islander
Certification date: end of '75
Number delivered: non - just certified
Production rate: unknown
Production category after January 1, 1979: IV
Effort Expended
Data reported by Lycoming (Ref. 35) indicates that evaluation of
control techniques for reducing emission levels from the LTP-101 combus-
tor have been studied for about four years. No quantitative fiscal data
has been reported, however.
Initially, a basic combustor design was evaluated. Emissions
testing indicated that this combustor has low efficiency at idle.
Experimental developments involved combustion zone aerodynamic develop-
ment (better mixing in the primary zone) and improved fuel injection
techniques (improved pressure atomizers and airblast). Because FAA
certification was scheduled for 1975 Lycoming claims time was not avail-
able to complete emission optimization of the design (Ref. 35), and a
partially improved configuration was chosen as the certification con-
figuration since it had accumulated enough endurance experience.
Results and Status
The certification configuration fails to meet any of the 1979
standards. Overboard air bleed at taxi-idle and higher idle speed will
help reduce HC and CO emissions further, but not sufficiently to reduce
the 1975 certification configuration below the standards according to
Lycoming estimates (Ref. 35).
A summary of the status of this engine is presented below.
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Status
Technology
Concept Category HC
Standard 4.9
Basic design 17.9
'75 cert, config. 1-2* 9.8
Final config. 1-2* 4-5
(estimated by
Lycoming)
Development
EPAP Status Implementation
CO NOx (Task No.) Date
26.8 12.9
44.1 5.3
34.7 4.8 8 (Fig. 58,60) mid 1977, or
late 1977
23-26 6-7 8 (Fig. 58,60) mid 1977,** or
late 1977**
Projections and Prospects
A development configuration (category 1-2 technology) with lower HC
and CO emissions has been demonstrated. Since this combustor produces
emissions below the standards (see table below) it seems that more
effort should be made to produce a flightworthy version of this combustor.
Lycoming has indicated that durability may be a problem with this combus-
tor. The emissions from this combustor are tabulated below.
Status
(development configuration)
Concept
Standard
1975 develop.
configuration
Technology
Category HC
1-2
4.9
3.6
Development
EPAP Status
CO NOx (Task No.)
26.8 12.9
Implementation
Date
17.5 6.2 2 (Fig. 58,60) early 1979
or early 1980
*It is not known if these configurations include airblast fuel nozzles,
**Lycoming claims 5 years are needed for development, certification,
and service evaluation.
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b. T5321A
Profile
Prospects for meeting the 1979 standards:
HC, CO - already met
NOx - already met
Equivalent shaft horsepower = 1800
PR = 8.0
Combustor type: reverse flow annular
Application: military
Certification date: not yet certified
Number delivered: unknown
Production rate: unknown
Production category after January 1, 1979:
III
Status
Technology EPAP
Concept Category HC_ CO
Standard 4.9 26.8
Production 2 17
Development
Status
NOx (Task No.)
12.9
6.0
Implementation
Date
Jan. 1, 1979
c. PLT27
Profile
Prospects for meeting the 1979 standards:
HC, CO - already met
NOx - already met
Equivalent shaft horsepower = 2000
PR = unknown
Combustor type: folded annular
Application: military
Certification date: not yet certified
Number delivered: still in development
Production rate: still in development
Production category after January 1, 1979:
IV
Concept
Standard
Production
Status
Technology EPAP
Category HC CO
Development
Status
NOx (Task No.)
4.9
0.3
26.8
3
-178-
12.9
12.8
Implementation
Date
Jan. 1, 1979
-------
Comment
Current production models of both the T5321A and PLT27 engines meet
the standards. No other low emissions data was reported. Apparently no
effort has been conducted to lower the emissions presumably because
these engines already meet the standards, and also because these two
engines currently have military applications. They are included here to
indicate the low emissions levels that have been achieved in other
Lycoming engines. Possibly this technology may be applicable to the
LTP-101 or ALF-502 engines.
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4. Allison
a. 250
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 2 technology by late 1979
NOx - already met
Horsepower = 400 (-B17 model)
PR = 7-2
Combustor type: single can
Application: Turbostar 502
Certification date: September '65
Number delivered: unknown
Production rate: "small quantities" according to Allison
Production category after January 1, 1979: III
Effort Expended*
Data reported in Reference 57 indicates a total of 1,145,000 dollars
was spent by Allison on the 250 low emissions effort for the time period
1971-1975. Of this total 65% (745,000 dollars) was government funds.
The only low emissions data available for this engine is that which
was developed under contract with the U.S. Army Air Mobility Laboratory.
During this one year program, approximately 97,000 dollars were spent
evaluating seventeen low emissions combustors. Each combustor incorporated
one or more of the following concepts (grouped by combustor area affected):
fuel injector . primary zone . primary dilution zone
air blast/air assist lean increased length
premix/prevaporization rich double combustor volume
staged variable geometry rapid plug flow
early quench
reverse flow
massive recirculation
swirl
heat rejection
water injection
cold air injection
Results and Status
The premix combustor has demonstrated the capability of controlling
exhaust emissions to a level below the standards. It is expected that
*It should be noted that data reported in Reference 57 indicates that
1,633,000 dollars were spent by Detroit Diesel Allison on aircraft
turbine engine related technology programs during the period 1970-1975.
Of this total 1,088,000 dollars (67%) were government funds.
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this is the combustor concept that Allison will use to reduce emissions
from the 250 engine. It should be noted that this combustor increased
NOx by 75 percent relative to the production combustor. However, NOx
remained below the standard. The emissions of the premix combustor are
presented below.
Status (-B17 model)
Concept
Standard
Technology
Category HC
4.9
Production
Premix combustor
EPAP
CO
NOx
Development
Status
(Task No.)
Implementation
Date
26.8 12.9
25 128 5.3
1.2 13.8 9.3 3 (Fig. 60) late 1979*
Projections and Prospects
It is expected that Allison will continue development of the premix
combustor and use it in the 250 engine. No known development problems
exist with this combustor.
b.
501
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 2 technology by mid 1980
NOx - already met
Horsepower = 400 (-D22A model)
PR = 9.5
Combustor type: multiple cans
Application: Lockheed Hercules
Certification date: c. 1968 (-D22A model)
Number delivered: 11,000 (756 and 501 engines combined)
for aircraft and industrial use
Production rate: 3300 during next 10 years
Production category after January 1, 1979: II, III
Effort Expended
Financial data reported to the EPA (Ref. 57) indicates that
2,423,000 dollars were spent during the time period 1970-1975 on 501
low emissions work. Sixty-one percent (1,485,000 dollars) were provided
by the government.
*AlliBon claims that 5-6 more years are needed to produce this combustor.
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The 501 engine was selected by NASA for the Class P2 portion of
their Pollution Reduction Technology Program (PRTP). This effort was
initiated in January 1975 and ended in February 1976. Approximately
242,000 dollars were spent on a cost share basis, Allison supplying 33%
and NASA, 67%. This program is the source of the only low emissions
data reported for this engine.
A total of 18 configurations of three basic low emissions combustor
concepts were tested. These are 1) the reverse flow combustor, 2) the
prechamber combustor, and 3) the staged fuel combustor. A description
of these concepts is given below.
1) The reverse flow combustor - This reverse flow combustor is
not the reverse flow annular type used in such engines as the TFE-731,
the JT15D, and the ALF-502. This reverse flow combustor is a can type
and incorporates a primary zone flow system which increases the amount
of recirculating products, improves fuel-air mixing, and returns the
partially burned products trapped in the cooling film, back into the
reaction. The remainder of the combustor is of conventional design. It
should be noted that each of the five combustor configurations of this
type incorporated either airblast or air-assist fuel nozzles.
2) The prechamber combustors - This type of combustor is called
premix-1 by the EPA and is described in part 13 of Appendix C. In this
program a total of six configurations of two basic designs were tested.
These designs were: a) a short prechamber combustor where fuel and air
are mixed but do not burn, and b) a long prechamber which does act as a
small combustor. Two noteworthy features of both designs are the airblast
fuel nozzles, and a variable geometry band used to open and close dilution
holes (the latter device will not be incorporated into flight hardware
as its main purpose was parametric investigation of airflow distribution).
3) The staged fuel combustor - Fuel staging is described in part
15 of Appendix C. A total of seven configurations of two basic designs
were tested. These were: a) an original staged combustor design, and
b) a modification of the original design. The original design incorporated
features to produce minimum HC and CO emissions. (No attempt to control
NOx was made since an increase in NOx was allowed.) This was accomplished
by providing for a lean pilot zone, low pilot zone airflow loading,
reduced wall quenching, a reverse flow feature in which CO and HC in the
dome cooling are returned to the reaction, and a swirl prechamber which
provides good fuel-air mixing in the combustion zone. The modified
combustor was designed to reduce CO further and reduce NOx. This combus-
tor had an increased pilot zone volume, and a reduced pilot zone equiva-
lence ratio.
No other low emissions testing has been reported on the Allison 501
engine.
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Results and Status
All of the basic designs selected were capable of controlling all
of the emissions to levels below the standards. The reductions of HC
and CO achieved were substantial; however, NOx generally increased
relative to the production combustor level. According to NASA (Ref. 34)
the performance characteristics of all the configurations were acceptable
and the measured exit temperature pattern factors were better than the
baseline combustor. The emissions of the best configuration of each
design are presented below.
Status (-D22A model)
Concept
Standard
Production
Reverse flow
mod IV
Prechamber
mod III
Staged fuel
mod V
Technology
Category HC
4.9
Development
EPAP Status
CO NOx (Task No.)
26.8 12.9
Implementation
Date
14 30 6.2
2 0.3 4.6 7.3 2 (Fig. 60) mid 1980
3-4 0.4 2.1 8.5 2 (Fig. 62) late 1981
3 0.6 5.7 7.2 2 (Fig. 62) late 1981
Projections and Prospects
It is expected that Allison would pursue development of one of the
above designs for the 501 engine, probably the reverse flow concept
since it is the least complex. No development problems have been
reported for any of these designs.
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5. Rolls Royce
a. Dart
Profile
Prospects of meeting the 1979 standards:
HC, CO - poor by late 1978
- fair by mid 1979
NOx - good by late 1978
Equivalent shaft horsepower = 2280
PR = 5.75
Combustor type: multiple cans
Application: Grumman Gulfstream I, Fairchild F-27
Certification date: c. April '53
Number delivered: 6500 as of January '75
Production rate: 50 per year halfway through the next decade
Production category after January 1, 1979: I
Effort Expended
No fiscal data quantifying the effort expended on low emissions
testing on the Dart engine has been reported. Further, the only low
emissions data available for this engine are the EPAP values achieved by
using duplex fuel nozzles in the combustor.
Results and Status
The duplex nozzles reduced HC, CO and NOx, but only NOx was controlled
to a level below the standards. The status of the Dart is presented
below:
Status
Development
Technology EPAP Status Implementation
Concept Category HC_ C0_ NOx (Task No.) Date
Standard 4.9 26.8 12.9
Production 100 125 <12.9
Duplex nozzles 1 66 95 1.4 3 (Fig. 58) late 1978
Projections and Prospects
The levels of HC and CO listed above are very much higher than the
standards. If the fuel flow were scheduled to operate on the primary
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atomizers only during taxi-idle such as in the TPE331, and the percen-
tage reductions in HC and CO, and the NOx increase were the same for the
Dart as the TPE331 the EPAPs listed below would result.
Projected Status
Concept
Standard
Primary
atomizers only
at taxi-idle
Technology
Category HC_
4.9
1 4.3
Development
EPAP Status
CCi NOx (Task No.)
26.8 12.9
27.7 1.6 1 (Fig. 58)
Implementation
Date
mid 1979
These levels may be somewhat optimistic as the rated pressure ratio (and
hence idle pressure ratio) of the Dart is lower than that of the TPE331-3
(5.75 compared to 10.4). Another rating 1 concept worthy of considera-
tion is to increase the taxi idle speed. (The amount of CO reduction is
a function of the amount of increase in taxi idle speed.) This should
be especially effective for this low pressure ratio engine and could be
done in conjunction with operating the primaries only.
It should be noted that at the January 1975 hearings Rolls Royce
requested total exemption for this engine because of the limited produc-
tion. Rolls Royce has estimated the production rate of the Dart at
approximately 50 engines per year halfway through the next decade. t
b. Tyne
Profile
Prospects for meeting the 1979 standards:
HC, CO - unknown
NOx - unknown
Shaft horsepower = 5500
PR = 13.5
Combustor type: straight flow annular
Application: Belfast, CL-44
Certification date: c. 1960
Number delivered: unknown
Production rate: unknown
Production category after January 1, 1979:
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Effort Expended
This engine is not included in the list of engines in Reference 44
for which reduced emissions programs are being conducted. Therefore
probably no effort has been made.
Results and Status
No data, either production or reduced emissions, is available for
this engine.
Projections and Prospects
It is not expected that Rolls Royce will make any effort to reduce
emissions from this engine because of the remaining limited production
of the engine.
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F. Auxiliary Power Units
Auxiliary power units are defined as ". . . any engine installed
in or on an aircraft exclusive of the propulsion engines." (See FR Vol.
38, No. 136, July 17, 1973, p. 19091.) These engines are used to
provide onboard pneumatic energy for jet engine starting, air condition-
ing, heating and anti-icing. Shaft power is provided for generators and
hydraulic pumps. The gaseous emissions standard listed below for these
engines is currently scheduled to go into effect on January 1, 1979.
Standard
EPAP
(Ibs pollutant per 1000 horsepower-hours at maximum power)
HC CO NOx
0.4 5.0 3.0
Certain common design characteristics of APUs, coupled with the
nature of the EPA standard, have created difficulties for APU manufacturers
attempting to comply with the standards. The first characteristic is
that APUs have in general been designed with low operating pressure
ratios, despite the SFC penalty incurred. The reasoning behind this is
threefold: (1) the pneumatic systems cannot take over 3-4 atmospheres
for operation; (2) the total fuel requirements of the APU are much less
than that of the main engines in the aircraft and, therefore, it is more
economical to design for low initial and low maintenance costs (i.e. low
sophistication, in particular, low pressure ratios) than to design for
better fuel economy with more sophistication; and (3) for many APUs
which are fairly small, a higher pressure ratio would mean exceedingly
small components in the high pressure segments leading to excessive
efficiency penalties.
The second characteristic is that being small engines in general,
many APUs have been built with reverse flow annular combustors (because
of the use of radial flow turbomachinery which retains better efficiency
on the small scale than does axial flow machinery). This results in two
problems: The first is that the surface-to-volume ratios of such
combustors are fairly high, leading to a propensity for high emissions
of HC and CO (See Section III); the second is that any attempts to
incorporate a fuel staging scheme into the combustor to control NOx meet
with difficulty because of the limited space available.
The third characteristic is that APU operation is not consistent
with the low emissions technical development which has been directed
towards the cycle type of operation of the propulsion engines. For the
APUs, the test condition is at a specific power point (maximum combined
shaft plus air bleed) and consequently some of the less complex (category
1) concepts used to control HC and CO, such as increased idle speed,
compressor bleed changes, etc., are simply not relevant.
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Another problem is that even water injection for NOx control
is not technically feasible in all cases; it creates a number of com-
plications which compromise its usefulness. In particular, the bulk of
the NOx production is not limited to particular modes as is the case for
propulsion engines, but is uniformly distributed throughout the operation
"cycle" of the APU. This implies that water must be used at all times
if it is to be used at all. As replenishment of the water is not
available immediately upon landing and may often not be available at
all, it then follows that water must be carried on board throughout the
flight. The additional tankage and weight create economic penalties and
the need to prevent freezing of the water requires further care. Mixing
water with methanol compromises CO control even more as the quenching
effect of the water and the generally low operating pressure lead to
incomplete combustion of the anti-freeze. Even straight water, in fact,
in some APUs (notably the low pressure ones) leads to excessive quenching
of the flame and, therefore, high HC and CO.
1. AiResearch
AiResearch has tested a variety of emission control techniques that
could be utilized for APU emissions reduction. These techniques have
been largely tested and evaluated on the GTCP85 APU or on a rig with
conditions simulating that APU. However, as many of these concepts are
applicable to any of the other AiResearch APUs and as the test results
can be extrapolated, AiResearch avoided repetitive testing of the same
concept in each engine unless it showed promise in the first instance.
As the concepts explored by AiResearch can thus be considered to be
relevant to each of its engines, it is reasonable to discuss the issue
of effort expended (in terms of technology exploration) in general terms
prior to the discussion of each engine.
Effort Expended
By and large, the concepts investigated by AiResearch have been
able to meet the HC and CO standards, while the preliminary investiga-
tions of category 3 concepts yielded no improvement in NOx. Water
injection (category 2), tested as an alternative to category 3 complexity,
creates problems in its own right (as discussed earlier).
The techniques explored by AiResearch since 1973 are categorized
into four main groups with the specific concepts listed under each.
These are as follows:
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I. Fuel preparation
1. Airblast (category 2) - A series of tests were made
on a research combustor to investigate the potential
of the pure and piloted airblast nozzle and the air-
assist nozzle. A total of eleven tests of different
injectors were made on a total of seven different
combustor airflow distributions or mixing schemes
for a grand total of 77 configurations. Subsequently
a piloted airblast system was investigated in both
GTCP85 rig and engine testing with three different
combustor airflow redistributions, and an air-
assist/airblast injector was checked in a GTCP36 rig
with several modifications to the airflow pattern.
2. Air-assist (category 1) - This concept was investi-
gated to the extend described in the airblast
discussion above.
3. Premix/prevaporization (category 3-4) - Two different
concepts were explored, both using the GTCP85 rig.
The first was a "scroll" combustor which acted like
a carburetor mechanism. Seven test conditions were
run on a single scroll configuration. The second
was a more conventional prechambered combustor
and it was evaluated at eleven test conditions.
II. Fuel distribution
1. Fuel staging (category 2-3) - A single staging con-
figuration was explored in the GTCP85 rig. The
combustor involved two axially seperated fuel
injection localities within a single combustor
(i.e., the two primary zones were not mechani-
cally separated). A parametric study (8 points)
of the optimal staging ratio was done.
2. Fuel atomizer design (category 1) - Three types
of changes were investigated over the APU operating
range, injection, pressure, distribution of fuel
between the orifices (dual orifice nozzle), and
change in the spray angle.
III. Air distribution
1. Lean primary zone (category 2) - This concept was
explored in the parametric studies of other concepts
(e.g., airblast).
2. Increased turbulent mixing (category 2) - This
concept was explored in the parametric studies of
other concepts.
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3. Improved uniformity in the primary zone (category
2) - Attempts to improve combustion through a more
thorough investigation of the effect of bleed port
location (three tested) and of a lengthened com-
bustor (eight tested).
IV. Water injection (category 2) - While only one
water/fuel ratio was investigated (= 1), two
injection points were checked, compressor inlet
and primary zone, as were two mixtures, pure
water and water-methanol solution for anti-freeze
capability. Water injection was tried in both
the GTCP85 combustor and in the TSCP700 combustor.
In addition, AiResearch ran a large number of tests on various
production type APU engines to determine the effects of production
tolerances, fuel variability, ambient effects, and oil seal leaking, in
order to better define the compliance problem.
Data quantifying the amount of dollars or inanhours spent on any
of this testing has not been reported to the EPA.
The discussion will continue now with an engine-by-engine analysis,
including profiles, results and status, and projections and prospects.
a. GTCP 85
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with either category 1 or 2 technology by 1979
NOx - poor with category 3 technology
Bleed air horsepower = 177-216
Shaft horsepower = 25-100
PR = 3.2-3.5
Combustor type: single can
Application: B727, DC-9, B737, BAG 111 etc.
Certification date: c. 1960
Number delivered: approximately 2500 as of mid '75
Production rate: unknown
Production category after January 1, 1979: III
Results and Status
Below are tabulated most of the concepts explored on the GTCP 85 or
rig equivalent in the AiResearch APU study.
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Status
Concept
Standard
Production
Air-assist
Technology
Category
Atomizer design
-narrow spray 1
-wide spray 1
Airblast 2
Lean primary zone 2
Increased 2
turbulent mixing
Water injection -
standard nozzles at
-H20/fuel =1.0 2
airblast nozzles at
-H20/fuel =1.0 2
Fuel staging 3
Development
EPAP Status
HC OO NOx (Task No.)
0.4 5.0 3.0
0.2 7.5 6.4
Claimed to be
similar to air-
blast. No EPAPs
available.
0.6 9.6 6.2 3 (Fig. 66)
0.1 3.7 6.9 3 (Fig. 66)
0.1
3.9 6.7 4 (Fig. 67)
Implementation
Date
2 (Fig. 66) mid 1978
early 1978
early 1978
mid 1978
mid 1978
Claimed there is 4 (Fig. 67)
no improvement for
APUs. No EPAPs
available.
0.1 5.1 7.9 4 (Fig. 67) mid 1978
0.8 19.4 2.9 4 (Fig. 67) mid 1978
2.0 15.0 2.7 4 (Fig. 67) mid 1978
0.1 4.0 9.2 2 (Fig. 68) mid 1980
Projections and Prospects
There are several possible approaches to compliance with the HC and
CO standards, but none apparently for compliance with the NOx standard
as even water injection is creating an unacceptable HC and CO penalty
from quenching. On the other hand, it is not at all obvious that a
staging technique cannot be better than has been demonstrated to date
despite the obvious technical problems.
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b. TSCP 700
Profile
Prospects for meeting the 1979 standards:
HC, CO - already met
NOx - good with water injection (category 2) by 1979
- good with category 3 by 1981
Equivalent shaft horespower = 910
PR =10.4
Combustor type: reverse flow annular
Application: DC-10, A300
Certification date: c. 1971
Number delivered: approximately 250 as of early '75
Production rate: unknown
Production category after January 1, 1979: III
Results and Status
The TSCP 700 benefits from a relatively high compressor pressure
ratio at its operating condition which is reflected in the low HC and CO
emissions of the production engine; it is also reflected in the adequate
HC and CO emissions performance even with water injection. This higher
pressure ratio engine is probably typical of the new design approach to
be taken in the future (see also the ST6 APU below).
Status
Development
Technology EPAP Status Implementation
Concept Category HC_ CO NOx (Task No.) Date
Standard 0.4 5.0 3.0
Production 0.2 0.8 6.0
Water injection 2 No data 1.6 3.0 4 (Fig. 67) mid 1978
Projections and Prospects
With nothing more than the use of water injection, the TSCP 700 can
comply with the standards. However, water injection is not a particularly
acceptable solution because of the need to carry water aboard and to
prevent it from freezing. Category 3 techniques have not been tried in
the 700 as yet and, for that matter, they are just now being investigated
in reverse flow annular combustors (in the TFE 731) as a part of the
NASA Pollution Reduction Technology Program. On the basis of the prelimi-
nary results of the TFE 731 testing, the following projection for category
3 performance has been made:
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Projected Status
Concept
Standard
Premix/prevap.
Technology
Category HC
0.4
Development
EPAP Status
CO NOx (Task No.)
5.0 3.0
1.7 2 (Fig. 68)
Implementat ion
Date
late 1980
It is important to realize that this projection can be made because the
TFE731 and the TSCP700 have the same type of combustor. Unfortunately,
no such projections can be made between engines of different combustor
types.
c.
GTCP 36
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 2 technology by 1979
NOx - unknown
Shaft horsepower = 192
PR = 3.25
Combustor type: single can
Application: F-28, Gulfstream 11, YS-11, G-222
Certification date: c. 1967
Number delivered: approximately 250 as of mid '75
Production rate: unknown
Production category after January 1, 1979: III
Results and Status
The production combustor of the GTCP 36 engine uses an air-assist
airblast fuel nozzle. Several rig and engine tests were conducted to
determine the emissions levels produced by this combustor during develop-
ment. The best demonstrated emissions along with the production engine
emissions are shown in the table below.
Concept
Standard
Production
Airblast
Technology
Category HC
Development
EPAP Status
CO NOx (Task No.)
0.4 5.0 3.0
0.3 19.8 5.2
0.04 <5.0 5.4 4 (Fig. 67)
Implementation
Date
mid 1978
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Projections and Prospects
Airblast is apparently the way that AiResearch will handle HC and
CO control. The test data suggest that this is unusually effective in
this engine (compare with the GTCP 85 data). NOx control would require
water injection as there is yet no category 3 technology that has been
successfully demonstrated in a can combustor type APU. However, as
water injection has not actually been demonstrated in this engine, it's
effect on the CO level, in particular, is in question.
d. GTCP 30
Profile
Prospects for meeting the 1979 standards:
HC, CO - already met
NOx - unknown
Shaft horsepower = 100
PR = 3.76
Combustor type: single can
Application: DC-7, Jet Star, etc.
Certification date: unknown
Number delivered: unknown
Production rate: unknown
Production category after January 1, 1979:
II
Results and Status
No emissions work has been done on this engine:
Concept
Standard
Production
Technology
Category HC
0.4
0.1
Development
EPAP Status
CO NOx (Task No.)
5.0 3.0
4.8 3.4
Implementation
D.ate.
Projections and Prospects
The manufacturer intends to stop production of this APU (reference
14, p. 2-11) and replace it with the GTCP 36-50. The production configu-
ration already meets the HC and CO standards and the NOx level is close.
If water injection were to be attempted to reduce the NOx level to that
of the standard, it is conceivable that something would have to be done
about the increased CO level. Even with the problem of water injection,
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the CO level is marginal and the need for a greater statistical margin
would require some correction. The airblast concept tested on both the
GTCP 85 and the GTCP 36 was quite successful and is relatively easy to
incorporate. The following EPAPs are projected if that concept were
tried in the GTCP 30:
Development
Technology EPAP Status Implementation
Concept Category HC_ C0_ NOx (Task No.) Date
Standard 0.4 5.0 3.0
Airblast 2 0.01- 1.2-3.5-2 (Fig. 67) mid 1979
0.05 2.5 3.6
e. GTCP 660
Profile
Prospects for meeting the 1979 standards:
HC, CO - good with category 2 technology by mid 1979
NOx - poor with category 2 technology;
- fair with category 3 technology by 1981
Equivalent shaft horsepower = 1100
PR = 3.5
Combustor type: straight flow annular
Application: B747
Certification date: c. 1969
Number delivered: approximately 300 as of mid '75
Production rate: unknown
Production category after January 1, 1979: II, III
Results and Status
No emissions work has been done on this engine to EPA's knowledge
due in part to the conviction by the manufacturer that the post-1979
production will be inadequate to justify the development expense.
Development
Technology EPAP Status Implementation
Concept Category HC CO NOx (Task No.) Date-
Standard 0.4 5.0 3.0
Production 0.2 7.9 4.9
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Projections and Prospects
The manufacturer may hope to obtain relief through a waiver or
a change in the standards rather than undergo the possibly unrecover-
able expense of developing a low emissions combustor for the 660 just
when the production rate of the Boeing 747 would be in its decline. Just
what Boeing would do for an APU in the event of the unavailability of
the 660 is uncertain. On the other hand, based on all of the available
test results from the GTCP 85 and the GTCP 36 APUs, it appears that HC
and CO emissions can be controlled on the engine to a level below the
standards by using the airblast fuel nozzles.
The projected values are:
Development
Technology EPAP Status Implementation
Concept Category HC CID NOx (Task No.) Date
Standard 0.4 5.0 3.0
Airblast 2 0.03- 2.0- 5.1 2 (Fig. 67) mid 1979
0.1 4.1
If this were combined with water injection, however, it is not certain
that the CO level, in particular, would remain below the standard (see
the GTCP 85 and TSCP 700 data). Alternatively, as this is a straight
flow annular combustor, it is possible that some of the category 3 con-
cepts which have been tested on propulsion engines having the same style
combustor may also be applicable here. Due to the smaller size of this
engine, though, it is not possible to make meaningful projections as to
the reductions of any of the pollutants.
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2. Pratt & Whitney Aircraft of Canada
a. ST6
Profile
Prospects for meeting the 1979 standards:
HC, CO - already met
NOx - poor with category 2 technology;
- good with category 3 technology by 1981
Shaft horsepower = 720
PR = 7.8
Combustor type: reverse flow annular
Application: L-1011
Certification date: c. 1979
Number delivered: approximately 200 as of early '75
Production rate: unknown
Production category after January 1, 1979: III
Effort Expended
No emission data either from current production or reduced emissions
concepts has been submitted to the EPA. However, as this engine is
directly derived from the PT6 turboprop/turboshaft series; it may be
surmised that the manufacturer intends to rely on his low emissions work
on the PT6 to solve any problems relating to the ST6.
Results and Status
The status of this engine was estimated from PT6A-41 data, employing
the best category 2 concept which has been investigated in the PT6.
Development
Technology EPAP Status Implementation
Concept Category HC C0_ NOx (Task No.) Date
Standard 0.4 5.0 3.0
Production 0.3 2.6 4.8
Airflow 2 0.02 1.0 5.2 3 (Fig. 67) late 1978
distribution
Projections and Prospects
Both the production and the airflow redistribution concepts control
HC and CO to a level below the standards. However, NOx remains high as
in other APUs. Fuel staging or premix/prevaporization concepts may help
to reduce NOx because of this engine's relatively high compressor pressure
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ratio (7.8). The only category 3 concepts that have been investigated
on a reverse flow annular cbmbustor such as this engine have been those
tested on the TFE 731 as a part of the NASA Pollution Reduction Technology
Program. The following emissions projection is made for the ST6 using
that technology:
Concept
Standard
Premix/prevap.
Technology
Category HC
0.4
EPAP
CO NOx
5.0 3.0
Development
Status
(Task No.)
Implementation
Date
1.3 2 (Fig. 68) late 1980
It is worthwhile to note that any category 3 concept for this engine
must be written off completely against the APU version as the turboprop
version already meets the NOx standard. It is, in fact, interesting
that the same basic engine can comply in the P2 class, yet with the
same high power operating conditions, fail to meet the APU class
standard.
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3. Solar
a. Titan-39
Profile
Prospects for meeting the 1979 standards:
HC, CO - unknown
NOx - unknown
Shaft horsepower = 40
PR = 3.0
Combustor type: single can
Application: business jets
Certification date: unknown
Number delivered: unknown
Production rate: unknown
Production category after January 1, 1979: III
Effort Expended
The only data available for this engine is production engine emissions
based on testing conducted in 1973. No improved combustor technology
data has been reported. Solar has stated (Ref. 60), however, that major
redesigns of the combustion system are anticipated in order to meet the
standards.
Results and Status
Development
Technology EPAP Status Implementation
Concept Category HC CO NOx (Task No.) Date
Standard 0.4 5.0 3.0
Production 0.6 21.0 4.4
Projections and Prospects
Nothing is known.
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78. Letter from H. C. Eatock (PWAC) to R. W. Hunt (EPA) dated May 14,
1976 Subject: Engine Dta.
79. The NASA Pollution - Reduction Technology Program for Small Jet
Aircraft Engines - A Status Report (NASA TM X-73419), July, 1976,
James S. Fear.
80. Letter from G. Opdyke (AVCO Lycoming) to R. Sampson (EPA) dated
Jan. 8, 1974 Subject: Emissions Data.
81. Letter from R. Fletcher (Cranfield Institute of Technology) to
E. K. Bastress (Arthur D. Little, Inc.) dated Jan. 29, 1975.
82. Experimental Clean Combustor Program, Phase III, Sixteenth Monthly
Progress Report, Oct. 28, 1976, Pratt & Whitney Aircraft
83. EPA No. 68-03-2156, Garrett AiResearch contractor, "Effect of'Ambient
Condition in Aircraft Engine Emissions."
84. Pollution Reduction Technology Program, Turboprop Engines - Phase I
Final Report, March 1976, Detroit Diesel Allison Division of
General Motors
85. Monthly Technical Progress Narrative No. 16, Small Jet Aircraft
Engine Program (NASA), April 15, 1976, Garrett Corporation.
86. Monthly Technical Progress Narrative No. 17, Small Jet Aircraft
Engine Program (NASA), May 17, 1976, Garrett Corporation.
87. Dex transmittal, A. Nelson (PWA) to E. Danielson (EPA), Nov. 16, 1976
88. Telecon, A. Nelson (PWA) to E. Danielson (EPA), Nov. 17, 1976
-206-
-------
89. Rolls Royce, CRR 60124, Interim Report, Exhaust Pollutant Measure-
ments on RB 211-226 Production Engines, July 1974.
90. Aircraft Emissions, Impact on Air Quality and Feasibility of
Control, EPA (no date).
91. Telecon, G. Opdyke (Lycoming) to E. Danielson (EPA), Dec. 1, 1976.
92. "Control of Air Pollution from Aircraft and Aircraft Engines,
Proposed Emissions Standards for In-Use Aircraft After January 1,
1983", Federal Register, Vol. 38, No. 136, July 17, 1973, p. 19050.
-207-
-------
Appendix A
PROGRAMS FOR REDUCTION OF ENGINE EMISSIONS
This appendix summarizes all the programs known to the EPA which
bear upon the compliance of aircraft gas turbine engines with the EPA
regulations. These programs are broken into two categories, engine
related programs and combustor research programs. The former are
directed exclusively towards the 1979 EPA standards and the latter are
applicable to all the standards in general as such programs serve more
to advance understanding than to concoct a specific solution.
Engine Related Programs
Government Sponsored Engine Programs
The government has sponsored and continues to sponsor several major
engine related low emissions combustor development programs. A summary
of these programs is provided in Table A-l. Research and development in
these programs were conducted by private companies, and in some cases
the company partially funded these programs on a cost sharing basis.
Industry Sponsored Engine Programs
Most of the manufacturers have initiated programs of their own to
identify and refine the technology necessary to satisfy the 1979 stan-
dards. In general two approaches were taken. These were modification
to current production hardware, and new combustor design concepts.
Again, as indicated above, some of these programs were conducted on a
cost sharing basis with the government. A summary of these programs is
provided in Table A-2.
A-l
-------
Table A-l
Engine Pvelated Programs - Government Sponsored
Agency/Company
NASA Lewis/Garrett AlResearch
(Small Jet Engine Program)
NASA Lewis/Pratt & Whitney
Aircraft (Can-Annular Program)
NASA Lewis/Detroit Diesel Allison
(Turboprop Program)
>
i
NASA Lewis/General Electric
(ECCP)
NASA Lewis/Pratt & Whitney
Aircraft (ECCP)
Army/Detroit Diesel Allison
Objective
Demonstrate combustor technology,
for the TFE 731 engine, which is
capable of meeting Tl class stand-
ards.
Demonstrate combustor technology,
for the JT8D engine, which is
capable of meeting the T4 class
standards.
Demonstrate combustor technology,
for the 501D-22 engine, which is
capable of meeting the P2 class
standards.
Demonstrate combustor technology,
for the CF6-50 engine, which
is capable of meeting the 1979
T2 class standards.
Demonstrate combustor technology,
for the JT9D engine, which is
capable of meeting the 1979 T2
class standards.
Develop and demonstrate tech-
nology sufficient to obtain a 50
percent overall reduction in
T63 (class P2) engine emissions
with no increase in any individual
pollutant.
Status
Phase I (combustor screening of
three phases is essentially com-
pleted.
The combustor concepts being
screened are:
1) modified existing combustor
2) alrblast system (meets HC
and CO standard)
3) prenix/prevaporization (meets
HC, CO and NOx standard)
Phase I (combustor screening)
is completed.and the report pub-
lished. The second phase has been
cancelled. The combustor concepts
screened were:
1) modified existing combustor
(meets HC standard)
2) vorbix combustor
3) premix/prevaporization
Phase I (oombustor screening) of
three phases is completed, and the
report published. The combustor
concepts screened were:
1) reverse flow
2) staged fuel
3) prechamber (meets HC, CO, NOx
standard)
Phase III (engine demonstration)
of three phases is approximately
70 percent complete. Pursuing
double annular combustor concept.
(Meets: HC, CO standards)
(Fails: NOx standard)
Phase III (engine demonstration)
of three phases is nearly com-
pleted. Pursuing vorbix com-
bustor concept.
(Meets; HC, CO, NOx standards)
Program completed. Best concept
is the premix combustor. (Meets:
HC, CO and NOx Standards)
Funding/Duration
693K Phase I /FY 75-76 (17 months)
513K Phase I /FY 75-76 (16 months)
242K Phase I /FY 75-76 (13 months)
1000K Phase I /FY 73-74 (18 months)
950K Phase II /FY 75 (12 months)
1200K Phase III /FY 76 (12 months)
BOOK Phase I /FY 73-74 (19.5 months)
950K Phase II /FY 75 (12 months)
1200K Phase III /FY 76 (14 months)
97K/FY 72-73 (12 months)
-------
Table A-l continued
Agency/ Company
Army/Williams Research
/Garrett AiResearch
/Pratt 4 Whitney Aircraft
/AVCO Lycoming
Army/Detroit Diesel Allison
EPA/Garrett AiResearch
Objective
Develop new high performance
turboshaft engines with low
emissions. . Engine has applica-
tion as turboprop (class P2)
or as an APU.
Refine best designs and concepts
from T63 program above and test in
a Model 250-C-20B (class P2)
engine.
Develop full scale engine data
to determine the effects of
ambient conditions on emissions
for the TPE331 (Class P2) and
GTCP85 (APU) engines.
Status
Program nearing completion
Williams concept - meets standards
AiResearch concept - fails standards
Pratt & Whitney concept - no data
AVCO Lycoming concept - meets standards
Funding/Duration
Unknown
Testing has been completed and
data is in final stages in pre-
paration for report. No data
published as yet.
Testing has been completed and a
final report is being prepared.
The emissions standards were
achieved in the TPE331 engine by
rescheduling the fuel flow to
operate only on primary atomizers at
taxi-idle.
293K/FY 74-76 (23 months)
Joint project 170K
TPE331 extension 16K
FY75-76UO months)
-------
Table A-2
Engine Related Programs - Industry Sponsored
Objective
Status
Funding/Duration
Garrett AiResearch
Garrett AiResearch
Develop combustor technology for
the TPE 331 engine which allows
it to meet P2 class standards.
Develop combustor technology
for the TFE 731 engine which
allows it to meet Tl class
standard.
Best concept is rescheduling fuel Unknown/1973-1974 (2 Years)
flow to operate on the primary
atomizers only at taxi-idle. (Meets:
HC, CO, NOx standards)
Best concept demonstrated was
the air assisted fuel nozzles.-'
Garrett AiResearch
Garrett AiResearch
(Small Jet Engine Program)
Pratt & Whitney Aircraft
Pratt & Whitney Aircraft
(ECCP)
Pratt & Whitney Aircraft
(Can-Annular program)
Pratt & Whitney Aircraft
Develop combustor technology for
APUs which allow them to meet the
APU standard.
Demonstrate combustor technology
for the TFE-731engine, which is
capable of meeting the Tl class
standards
Develop combustor technology for
the JT9D engine which allows it
to meet 1979 T2 class standards.
Demonstrate combustor technology
for the JT9D engine which is cap--
able of meeting the 1979 T2 class
standards
Demonstrate combustor technology
for -the JX8D engine which is
capable of meeting the T4 class
standards.
Develop low smoke combustor for
JT3D (class T3) engine.
Best concepts demonstrated were:
GTCP 36 - airblast - meets KC and
CO standard; fails
NOx
GTCP 85 - increased fuel spray
cone angle - meets HC
and CO standard; fails
NOx
GTCP 30 -/production combustor -
TSCP 700 «meets HC and CO stan-
\dard; fails NOx
" GTCP 660 - production combustor -
meets HC standar'd; fails
CO and NOx
See Table I
Best concept demonstrated was
aerating fuel nozzles combined
with-modified airflow distribution.
(Meets: HC, CO standards)
(Fails: NOx standard, unless water
is used.)
See Table I
See Table I
Best concept demonstrated combined
delayed dilution with aerating
fuel nozzles. This concept signi-
ficantly reduced HC and CO while
increasing NOx.
(Fails: HC, CO, NOx standards')
Cost shared with NASA
Company share unknown
See Table I
Unknown/1972-1975 (3 Years)
Cost shared with NASA
Company share unknown
Cost shared with NASA
Company share unknown
See Table I
See Table I
Unknown/1963-ly74 (6 Years)
-------
Table A-2 continued
Company
Pratt & Whitney Aircraft
United Aircraft of Canada
Limited
United Aircraft of Canada
Limited
Objective
Develop low emissions combustor
technology for the PT6A-27 and -41
(class P2) engines
Develop low emissions combustor
technology for the JT10D (class
Tl) engines
Status
JT12A class Tl engine is expected
to be out of production and there-
fore no effort was made to reduce
emissions from it.
The best emissions results were
obtained with the combustor air-
flow redistributed. Results
were:
PT6A - 41 - meets HC, CO, and NOx
standard
PT6A - 27 - meets CO and NOx stan-
dard; fails HC
The most effective concept
demonstrated was 10% compressor
bleed at idle and approach.
(Fa~lTsT~~HC~,~COy NOx standards)
Fund ing/Dura t ions
190K/1973 (1 Year)
255K/1971, 1973 (2 Years)
General Electric
General Electric
(ECCP)
Detroit Diesel Allison
(Turboprop Engine Program)
Determine the most effective
method of reducing emissions in
the CJ610/CF700 (class Tl) engines
Demonstrate combustor technology
for the CF6-50 engine, which is
capable of meeting the 1979 class
T2 standard.
Demonstrate combustor technology
for the 501D-22 engine, which is
capable of meeting the P2 class
standards.
Reducing the fuel spray cone angle
was the most effective concept
demonstrated. With this method
HC levels were reduced_60 percent
relative to current engine (still
fails to meet standard) while NOx
remained acceptable (meets stan-
dard) and CO was unchanged (fails
to meet standard).
See Table I
See Table I
Unknown/1973-1974 (2 Years)
Cost shared with
NASA
Company share
unknown
Cost shared with
NASA FY 75-76
Company share 81K
Phase I (13 mos.)
See Table I
See Table 1
-------
Combustor Research Programs
Government Sponsored Combustor Research Programs
Several non-engine related combustor development programs have also
been sponsored by the government. These programs include grants to
private industry, research laboratories and universities, and NASA and
Air Force "in-house" programs. A summary of these programs is provided
in Tables A-3, 4, and 5.
Also, the government has sponsored a number of analytical combustor
modeling studies to support the development of necessary low emissions
technology. These studies have also been conducted by private industry,
research laboratories and universities, and NASA and the Air Force. A
summary of these programs is provided in Tables A-6, 7, and 8.
Industry Sponsored Combustor Research Programs
Although such efforts certainly exist within the industry, the EPA
has no explicit knowledge of them, perhaps because they may be appendages
to larger engine related programs.
A-6
-------
Table A-3
Government Sponsored Combustor Development Program
Private Industry Non Engine Related
Program/Sponsoring Agency
Investigator
Objective
Status
Augmentor Emission Measurement
Program/USAF
Augmentor Emissions Reduction
Technology/HASA Lewis
Advanced Concepts to Reduce NOx
in Aircraft/NASA Lewis
General Electric
Company
General Electric
Company
Solar Division of
International Harvester
Develop methods for deter-
minatibh'*of the emissions
levels of afterburning
engines
Reduce emissions from duct
burning engines
Reduce ilOx at supersonic
cruise conditions
Program completed.
Approximately 80 percent
complete
Program completed.
690K FY 73-75
500K FY 75-76
250iC FY 74-75
Low HOx Emission Combustor/NASA
Lewis
Full Annular Low HOx Emission
Combustor/HASA Lewis
Fundamental Low Power Emission
Study/NASA Lewis
Combustor Exhaust Odor Intensity
and Character Study/NASA Lewis
Supersonic Cruise Combustor
Pollution Technology/iJASA Lewis
General Applied Science
Laboratory
Solar Division of
International Harvester
General Electric
Company . _
A. D. Little
Pratt & Whitney Aircraft
General Electric
Company
Demonstrate that premix concept
will result in low MOx at '
supersonic cruise conditions.
Investigate application of jet-
induced-circulation coiabustor
concept to full annular Combustor
Study advanced concepts for
reducing low power emissions that
will permit the design of combus-
tors meeting the 1931 standards.
To determine the intensity of
odors of exhaust gases from
various turbojet engines or com-
bustors at several engine opera-
ting conditions, and to determine
ambient odor intensities at stra-
tegic airport locations.
,To investigate combustion concepts
specifically designed for low HOx
emission at supersonic cruise
operating conditions. This is an
addendum to ECCP.
Program completed.
Single source RFP has been
issued.
Contract just signed.
Program completed.
Program completed.
78K.FY 74
75K FY 75
40K FY 76
300K FY77+
114K FY 72-73
25K FY 74-75
200K to
each investiga-
tor FY 73-74
Advanced, High-Temperature-Rise
Combustor/AMRDL
Low Emission Jet Combustor/AMRDL
Low HOx Emission Combustor for
Automobile Gas Turbine Engines/
EPA-AAPS
Hydrogen Enrichment for Low
Emission Jet Combustion/
.s'ASA Washington, D.C.
Garrett AiResearch
Jet Propulsion Labora-
tory
United Aircraft of
Canada Linited
Jet Propulsion Labora-
tory
Minimize HOx, CO, and HC and
smoke levels for 2-5 Ib/sec
coHbustors at 12-16 atm inlet
pressure
To generate and evaluate uncon-
ventional combustor design con-
cepts.
Develop reduced emission and low
fuel economy automotive gas tur-
bine engine.
Project completed.
Program is 60 percent completed.
Project completed.
Experimentally evaluate the H~ - Currently, experimental evalua-
enrichoent concept with a research tion of the ^-enrichment concept
combustor. is underway with a research com-
bustor utilizing premixed H^/
fuel/air mixtures.
415K FY 70-73
Unknown
248iC FY 71-74
150K/year
FY 7J, 74,
75
Independent Research and Develop-
ment /DOD
Engine contractors
with DOD
Private industry is granted a
limited amount of government funds
to conduct independent research
programs some of w'nich will sup-
port low .-;nissions studies
-------
Table A-4
Government Sponsored Combustor Development Programs
University Grants
Program/Sponsoring Agency
"Investigator
Objective
Status
Mixing in High-Intensity Combustors/
NASA Lewis
Flow Processes in Combustors/
NASA Lewis
Study of Techniques for Lean
Combustion Systems/NASA Lewis
Prof. R.A. Strehlow
at University of
Illinois
F. Gouldin,
S. Leibovich,
F. Moore at
Cornell university
M. Branch,
A. Oppenheira,
R. Sawyer at
University of Califor-
nia, Berkeley
Influence of Intensity and Scale
of Turbulence on Emission Con-
centrations in the Primary Zone
of a Gas Turbine Combustor/NASA
Lewis
Prof. A.A. Kovitz
Northwestern
University
at
To develop means to probe
molecular-level niixedness in
high-intensity coiabustors
.Measure and predict velocity and
turbulence levels in swirling re-
circulating flows with and without
chemical reactions.
Study combustion processes of fuel
lean, turbulent flames at realis-
tic gas turbine conditions to de-
velop stabilization techniques
and improve idle performance.
To determine the effect of turbu-
lence on emission concentrations
within a-premixed combustor using
gaseous propane as a fuel.
The effect of turbulence on 51K/year
molecular mixing is being in- FY 73, 74,
vestigated. - 75, 76-
Several experiments in progress 64K FY 75
75tC FY 76
Assembly of test equipment is in 75K FY 75
progress. 73K FY 76
Project has been completed and a 48K/year
technical report has been pre- . FY 72, 73,
pared. . JJi
-------
Table A-5
Government Sponsored Corabustor Development Programs
NASA/Air Force
Evaluation of Techniques for
Pollution Reduction
Investigator
NASA Lewis Research
Center
In-house Exhaust Emissions
Investigation
Swirl Can Combustors
Twin-Ram Induction Combustors
Basic Pollution Research
AFAPL
NASA Lewis Research
Center
NASA Lewis Research
Center
NASA Langley Research
Center
Single Combustor Rig Investigation AFAPL
of Turbine Engine Exhaust Emis-
sions
Gas Turbine Catalytic Combustor
Premixed Variable Geometry Com-
bustors
Multiple Point Fuel Distribution
AFAPL & Engelhard
Industries
NASA Lewis Research
Center
NASA Lewis Research
Center
Objective
Reduce emissions from turbojet
coiabustors and reheat burners.
Provide limited investigated to
further understanding of aircraft
engine exhaust problems.
Status
Continuing "in-house" effort.
The effect of exhaust gas re-
circulation was examined and a
final report is in preparation.
Study of catalytic Combustors
was studied and the report issued.
Altitude emissions of refanned
JT8D was studied and the report
issued.
Program began April 1975 and
will continue for 3 years.
250K FY 73
193K FY 74
50K FY
100K FY
75
76
Develop short-length high per- Continuing "in-house" effort.
formance turbojet engine Combustors
having low levels of exhaust emis-
sions suitable for Mach 3 flight.
Develop a short double-annular
Combustors for advanced turbojet
engines.
a) Study reaction kinetics of NO
formation in fuel rich systems
behind shock waves.
b) Study effort of very high
pressure on NOx formation.
c) Study the kinetics of soot
formation and decay using laser
light-scattering techniques.
d) NOx formation studies in a
jet-stirred reactor.
To investigate novel control and
measurement techniques. '*".'? L-'
Investigate the catalytic
combustor concept.
Investigate the pollutant
reductions possible with premixed,
variable geometry Combustors.
Study ways of dividing and
simultaneously regulating one
flow stream into several equal
flow multiple streams.
Task completed FY 74. Reports
on final stage is complete.
Continuing "in-house" effort.
Unknown
185K FY 73
125K FY 74
40K FY 75
50K FY 74
50K FY 73
23K FY 74
60K FY 75
Continuing "in-house" effort.
Continuing "in-house" effort.
Project Completed.
Project Completed.
Unknown
Unknown
125K FY 74
150K FY 73
-------
Table A-6
Government Sponsored Combustor Modeling Programs
Private Industry
Investigator Objective
Combustors Design Criteria
Validation/AMRDL
Development of a 3-D Combustor
Flow Analysis/AFAPL/FAA
Garrett AiResearch
United Technologies
Research Center
Low Power Turbopropulsion Combustor Pratt & Whitney Air-
Emissions /AFAPL craft
Combustor Flow Analysis Procedure
AFAPL/FAA
Analysis and Investigation of
Exhaust Emissions from Non-
Afterburning Aircraft Ga&:
Turbine Engines During High
Power/AFAPL
Turbojet Combustor Pollution
Formation/NASA Lewis
.United Aircraft
Research Laboratory
Pratt & Whitney Air-
craft
General Applied
Science Laboratory
Develop and validate existing
analytical combustor design
procedures which can be used
to significantly shorten the
design time of 2-5 Ib/sec. com-
bustor s.
To modify and extend an existing
three-dimensional Navler-stokes
calculation procedure which in-
cludes the effects of chemical
reaction, turbulence transport,
radiation energy flux, and
droplet burning, and vaporization.
To identify and develop improved
component design techniques which
will increase part-power per-
formance.
To refine an existing analytical
technique and develop a computa-
tional procedure for predicting
combustion system performance.
Optimize design features through
insight gained from analytical
modeling of the combustor at
low power and from experimental
results.
Develop a modular, computer- program
to describe the coupled flow and
combustion kinetics''in .a turbojet
combustor..
Status
Preliminary version of computer
program is operational.
Program completed.
25K FY 74
91K FY 75
Project completed.
Project completed.
Project completed.
Project completed.
100K FY 71
600K FY 72
315K. FY 73
178K FY 74
85.8K FY 72
75K FY 73
25K FY 74;
72
20K FY 73
Analysis of Combustion Systems
with Swirl/NASA Lewis
Advanced Technical
Laboratory
Perform numerical.modeling of
swirling flows using finite
difference techniques for pre-
dicting the effect of swirl on
combustor performance and emis-
sions.
Project completed.
84K FY 71-73
-------
Table A-7
Government Sponsored Combustor Modeling Programs
University Grants
Investigator
Objective
Status
Study of Air Pollution from
Aircraft/MSA Lewis
Study of Parameters Affecting
Emissions of Gas-Turbine
Engines/NASA Lewis
Basic Chemistry of Aircraft
Pollutants/NASA Lewis
Semi-Empirical Correlations
for Aircraft Emissions/EPA
John B. Heywood at
MIT
Prof. C.W. Kauffraan
University of
Cincinnati
Professor A. Berlad
SUNY-Stonybrook
Prof. A.M. Mellor
Purdue University
To predict NOx and CO levels in
combustors and to investigate
fuel atomization, fuel vaporiza-
tion, and the mixing process in-
side the conbustor.
Analyze combustor emission data
and formulate a mathematical
expression that correlates
measured emissions with primary
operating conditions.
Survey current reaction rate
literature on cheaical reaction
affecting atmospheric ozone
levels, establish analytical model
of ozone balance in the atmosphere,
and indicate required experimental
rate data needed to solve model.
To determine a scaling theory for
the prediction and optimization of
pollutants for aircraft gas turbine
combustors
Continuing research effort.
Data in process of being
analyzed.
Project completed.
66K FY 74
79K FY 75
115K. FY 76
33K FY 75-76
43K FY 73-75
Continuing research effort.
33K
40K
46K
60K
55K
FY 72
FY 73
FY 74
FY 75
FY 76
-------
Turbine Engine Exhaust Emissions
Technology
Aircraft Engine Emissions Cor-
relation Technique
Three Dimensional Combustion
Flow Analysis
Combustor Analytical Modeling
of Pollution Formation
Table A-8
Government Sponsored Combustor Modeling Programs
HASA/Air Force
Objective
Investigator
AFAPL
Air Force Weapons
Laboratory
Wright-Patterson AFB
NASA Lewis Research
Center
Develop ambient temperature and
humidity correction factors for
CO and HC.
To predict CO, HC, and HOx and
particulate emissions indicies
for turbine engines used by the
Air Force.
Develop a mathematical model
of the combustion process for
annular and can-annular combustors.
To study the effects of mixing
and fuel preparation on the
formation of emissions.
Status
Project completed.
Continuing "in-house" effort.
Project completed.
Project completed.
Funding
40K FY 72-73
2K FY 74-75
5K FY 76
BOK FY 72-74
50K FY 74
>
i
-------
Appendix B
Data Bank
The appendix summarizes the pertinent technical data for the
combustor operating conditions and combustor emissions performance at
various power levels for each production or baseline engine and various
advanced emissions control concepts applied to that engine. For the
advanced concepts, the best demonstrated concept of each complexity
category (see Section II, Table III-l) is listed and its data tabulated.
Data may come from actual engine test, engine rig test (raw or extrapo-
lated data), or projection from related rig test. In addition, engine
data was sometimes given only in terms of EPAP, thereby forcing an
estimate of the associated Els to be made. Consequently it must be
recognized that not all of the data recorded herein are equally accurate
or valid although considerable effort was made to cull the wheat from
the chaff.
B-l
-------
ENGINEl
S/V - 9.92 ft
PRODUCTION
TFE 731-2
-1
T rated: 3500 Ibs.
V 13
IDLE POWER:
5-7 % rated
. P3 T3 Ha Mf
MODE (qtra) (°R) (///S) ft f/a (#/hr) >;, 1-V
IDLE
APPROACH
CLJ.MHOUT
TAKEOFF
ADVANCED:
IDLE :
APPROACH
CLIMHOUT
TAKEOFF
ADVANCED:
IDLE
'
APPROACH
CLlMttOHT
TAKEOFF
1.87
4.85
11.9
13.0
Air-a
(270
Pilo
(w:
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1180
sist 1
sig)
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prodi
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19.2
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9.97
6.1
18.0
20.2
'
-------
TFE
L'N'GINE:
S/V = 9.92 ft.
PRODUCTION
731-2
-1
T rnted:
* : 13
3500 Ibs.
IDLE POWER:
5.7 % rated
MODE
IDLE
APPROACH
CLTMiJOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
f CLIMI5GUT
TAKK01-T
ADVANCED:
IDLE
APPROACH
CI.lMnOUT
TAKEOFF
P3
Cltro)
1.87
4.85
11.9
n.o
Prem
pr
T3
(°R)
640
867
1147
1180
.x/pre-v
same
as
>ductic
M
(///S)
4.61
12.0
26.2
28.2
aporis
n
ft
.646
.199
.051
.044
ed com
f/a (
0098
0120
.0148
.0155
mstor
Mf
#/hr)
(Catej
ory 3-
I
4)
l-nc
.035
.006
.0003
.0003
.009
HC
22.4
2.9
.09
.09
2.07
El
CO
61.4
18.3
1.3
1.2
30.74
N0x
3.0
6.9
17.1
19.2
3.77
-------
ENGINE: ATF3
S/V =
PRODUCTION
T rated: 4,050
23.0
IDLE POWER:
% rated
V
MODE
IDLE
APPROACH
CLlMiiOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMBOUT
T A KM) IT
ADVANCED:
IDLE
AIT' ROACH
CLIMBOUT
TAKEOFF
P.
(Atio)
23.0
T3
H
(///S)
40.0
Jl
f/a
Mf
(#/hr)
1820
l-nc
HC
El
CO
>
NO
-------
ENGINE: JT15D
S/V = 11.4 ft
PRODUCTION
-1 T rated: 2500 Ibs
it : 10.1'i'
c
IDLE POWER:
8 % rated
MODE
IDLE
APPROACH
CLIMiiOUT
TAKEOFF
ADVANCED**
IDLE
APPROACH
CLIMliOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMnOHT
TAKEOFF
P3
(q'tra)
1.82
4.51
9.31
10.12
10% c(
and a]
? 1
4.5
Airf!
' T3
(°R)
659.1]
905.9
1118.1
114}. 3
mp. bl
proach
7O
906
sa
ow dis
M
(0/s)
5.49
11 ,.10
20.58
22.23
eed -at
5.73
12.21
ne as
:ribut
sa
.-
d
produ
ft
1.0412
.2543
.0849
.0740
Idle
.7490
.2897
>roduc1
.on/ In;
ie' "
:tion.
f/a
.0099
.0125
.0174
.0182
Ccate
,0123
.0132
ion.
ector
.
Mf
l#/hr)
215
^81_
1247
1405
gory =
253
581
nodifi
1)
nation
\
I . '
,
(d
\
tegory
.
= 2)
' '
l-nc
.0515
.0084.
.0003
.0002
.02f^
.0063
.pm1;
..0006
.0005
HC
.34.8
3.3
0
0
12.7
1.1
0.4
b
0
0
El
CO
. 90.5
23.8
1
i.o:
fiS.9
17.2
25.8
6.4
2.6
, 2.2
NO
X
2.5
5.1.
9.1.
10.1
i
V
6.2
3.1
5.8
7
.9.8
**HC and CO Els estimated.
-------
ENGINE! CJ610-2C
'1
8.51 ft.
PRODUCTION
T rated: 2950.1bs.
6.8
IDLE POWER:
6.64 ' % rated
P, ' T- M 1
3 3 a
MODE (Atm) (°R) (#/S) ft f/a (j
IDLE
APPROACH
CLIMBOUT
TAKEOFF
ADVANCED:
IDLE
t
APPROACH
CLIMUOUT
TAK1-OFF
ADVANCED:
IDLE
APPROACH
CL1MK01IT
TAKEOFF
E
t
Vi
Reduc
Air i
ngine
le sam
dt_&
ed spr
low di
lata e
i as t
ty con
;tribu
sentii
e CF7C
; angl<
:ion ((
lly1
0
(Gate
ategoi
%
A/hr) 1- c
510
1025
2430
2780
gory =
7 = 2)
1)
.052
.023
.006
.006
.039
.020
.nnfi
.006
.004
.020
.006
.006
El
HC CO
18.0
2.7
0.2
0.1
3.0
.6
.T
.1
14.7
2.9
0.2
0.1
1550
88.0
27.0
27.0
155
:85.6
77 n
27.0
134.9
74.3
27.0
77. n
NO
X
0.9
1.5
3.7
4.2
0.9
1.5
* 7
4.2
0.9
1.5
3.7
.L. 7
ta
-------
ENGINE: CF700-2D
S/V = 8.51 ft"--
PRODUCTION
T rated: 4250 lbs '
IDLE POWER:
V
6-24
M
El
MODE (Atra) (°R) (///S) ft f/a ^hour)
l-nc
HC
CO
N°
IDLE
APPROACH
CLIMiJO'JT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIM150UT
TAICI-OFF
ADVANCED:
IDLE
APPROACH
CL1MHOUT
TAKEOFF
1.50
2.72
5.78
6.94
Reduce
Airf
605
716
890
980
t? spra;
.
Prc
.ow dii
Kiwv
10.6
18.6
38.4
41
r cone
Same
as
duct j.c
tribut
as
2.745
1.344
.518
.337
angle
*
n
ion (C
.012
.014
.017
.018
(Categ
itegor
.
460
919
2322
26.07
3ry = :
' = 2)
)
' \
I
,
\
i
."
,
' '
'
.052
.016
.0059
.0052
.039
.0145
.0059
.0052
.044
.013
.006
.005
18.0
1.4
0.1
0.1
3.0
.57
0.1
0.1
ILJ,
1.5
0.1
0.1
155.0
62.0
25.0
22.0
L55.0
J0.3
!5.0
22.0
134.9
51.0
25.0
22.0
.90
1.8
4.3
5.6
.90
1.8
4.3
5.6
.9
1.8
4.3
.5.6
.
-------
ENGIN'E: M45H
S/V =
PRODUCTION *
T rated: 7500 Ibs
IT : 16 V;
IDLE POWER:
5% rated
MODE
IDLE
APPROACH
CLlMiJOUT
TAKEOFF
ADVANCED:**
IDLE '
M APPROACH
oo CLIMKOUT
TAKEOFF
ADVANCED:
IDLE .
'
APPROACH
CL1MHOUT
TAKEOFF
P3
Wtra)
Jnknov
V
<°R)
n (&
.
M
<0/s)
."'
.tegorj
..
ft
3v4
1.0
0.22
0.20
'" 1.-2
f/a
*
) '
1
,
Mf
(#/hr)
3450
. \
i
\
.
. .'
t
' '
l-nc
.090
.019
.0029
.0026
.0088
.*
HC
81.6
11.0
1.0
0.9
2.56
El-
CO
80
40.5
8.5
8.0
28.2
NO
X
i
'
' - - -,
* Not yet in production
** HC and CO Els. estimated
-------
ENGINE: RB401
S/V =
PRODUCTION *
T rated: 5000 Ibs.
IT : '17.0;'
c
IDLE POWER:
' % rated
MODE
IDLE
APPROACH
CLlNiJOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMIiOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CUM BOUT
TAKEOFF
*3
17.0
3
<°R)
.
M
(///S)
.
.
..
ft
'
f/a
.
Mf
(#/hr)
\
\
' .
.'
(
l-nc
HC
El
CO
.
N0x
'
w
* Not yet.in production
-------
LNGIN'E: ALF502D
s/v = 11-76 ft.
<
PRODUCTION *
-1 T rated: 6500 Ibs.
* : 10.7
IDLE POWER:
6.2 % rated
El
MODE (a!trn) (°R) (tf/S) ft f/a (#/hr) l-nc HC CO
IDLE
APPROACH
CLI.MI50UT
TAKEOFF
ADVANCED:**
IDLE
w APPROACH
M CLIMBOUT
/>
TAKHOFF
ADVANCED:
IDLE
APPROACH
c LI M no: IT
TAKEOFF
2.3
5.2
10.3
0,7
Air b
Airbli
695
900
1110
1125
.ast nc
st (Ca
9.6
19.8
33.3
34.5
zzle (
tecory
no lin
2)
_nm
.0117
.0217
.0240
iar ch
377
844
2456
272O
nge)
Catego
ry 1)
.0360
.0206
\J L U
.0142
14.?
-
-
9.9
same
qcj
JUCU-LU
JJ.
0.4
0.1
0
inn «
-
-
-
51.3
i
41.1
9.2
0.4
. 0
NO
X
9 S
5.0
13.0
13.2
4.4
2.8
5.6
8.0
.8.4
* Not yet in production
** HC and CO Els estimated
-------
ENGINE: CF6-6D
S/V = 7.56 ft'1
PRODUCTION
T rated: 38900 Ibs
» : 24.7.';'
c
IDLE POWER:
3-34% rated
MODE
IDLE
APPROACH
CLIMHOUT
TAKEOFF
ADVANCED:
IDLE '
APPROACH
CLIHliOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CL1MROUT
TAKEOFF
P3
(Atra)
2.73
24.52
Comprf
Pr
' V
<°R>
783
1403
ssor B
.
same
as
)ducti<
M
(///S)
22.3-
156.6
.eed
m
ft
.326
.0140
[Categc
f/a
.0132
.0244
ry - 1
.
Mf
(#/hr)
1063
3864
L1329
13750
1
1290
. \
\
.
.'
,
' '
l-nc
.032
.0029
.00066
.00066
.016
Proc
HC
20.5
1.80
.60
.60
7.7
same
as
uction
t
El '
CO
61.2
6.00
.60
.60
39.7
NO
X
4.59
10.75
27.29
34.00
i
4.6
..
'
-
-------
ENGINE: CF6-50C
S/V = 8.28 ft'1
PRODUCTION
. - -I.-J.
T rated: 49900 Ibs
v 29-8/;-
IDLE POWER:
3-39 % rated
P3 ' T3 Ma - Mf
MODE (iitra) (°R) (///S) ft f/a (#/hr) ' l-nc
IDLE
APPROACH
CLIMHOUT
TAKKOFF
ADVANCED:
IDLE
APPROACH
CLIMI50UT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CL1MBOUT
TAKEOFF
2.92
11.7
25.9
29.8
?roduc
sector
Doubl
(fu<
72
1134
1415
1476
tion c
burn!
.
Pro
>. anna]
1 stag
Pro
36.1-
125-. 0
261. 0
269.0
nnbusti
ig at
same
as
luctio
ir com
i-ng)
same
»
as
luctio
.576
.079
.022
.017
r plus
die C
i
mstor
i
.0098
.0118
.0181
.0195
]atego]
(Cate
,
12Q6
5280
15662
18900
T = i:
?ory =
3)
\
.
'«
.
.'
,
' '
'
.043
.0010
.0001
.0001
.0105
Prc
.0064
.0007
.0001
.0002
El
HC CO
30
0.01
0.01
0.01
2.2
same
as
iductic
2.2
0
0
0.06
73.0
4.3
).3
).02
36.0
n
L9.3
3.1
0.4
0.52
N0x
2.5
10.0
29.5
35.5
>
3.6
3.0
8.9
13.3
16.9
*
**
to
** HC and CO Els estimated
-------
ENGINE! CFM 56
S/V
10.68 ft
-1 T rated: 22,000 Ibs. IDLE POWER:
TT : 24 .
PRODUCTION*
4 % rated
P ' T
M
El
MODE (qtra) (°R) ry = 1
:or bui
)
ning (
\
-
,
]atego
\
.
:y_= 2
.'
f
*
'
.0249
.0016
.OOQ5
.0005
.0175
Pro
.0138
9.14
' ' -
.29
^ 0
^ 0
- 5.0
same
as
luctioi
1.14
uO
%0
^0
72,85
5.71
2.29
2.29
56.5
53.3
M)
^0
, ^0
N0x
2.14
7.3
16.1
18.2
i
4.4
*v 4.4
* 7.3
\J.6.1
0,18.2
'
_
.
* Engine not yet in production
**HC and CO Els estimated
-------
ENGINE: JT9D-7
S/V = 5.88 ft"
PRODUCTION
T rated: 46150 Ibs,
tr : 21.1
IDLE POWER:
8.2' % rated
. P3 ' T3 Ma - ' ' Mf
MODE (ACID) (°R) (///S) ft f/a (#/hr) l~nc
IDLE
APPROACH
CLIMiJOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMIJOUT
TAI'J-OFF
ADVANCED:
IDLE
APPROACH
CLIMIJOUT
TAKEOFF
3.95
8.5
18.5
21.1
Air a
Pro
Aerati
Pro
835
1055
1320
1380
S3 1st
same
as
ductio
ig noz
Rjimp
as
luctio
51.4
99-. 4
189.0
208.6
[Categ
-
i
;le co
.. -
i
.226
.074
.021
.017
>ry = !
*
ibustoi
.0100
.0130
.0194
.0215
) .
(Gate
.
1849
4648
13193
16142
gory =
2)
\
.
^\
.
.'
,
' '
'
.044
.0031
.0002
.0001
El
HC CO
29.8
1.0
0.1
0.05
7.0
Pr
1.1
77.0
9.6
0.5
0.2
45.7
same
as
oducti
6.9
.
NO
X
3.1
7.8
21.4
29.4
i
3.1
>n
2.9
7.3
20.1 ;
27.6
*
'
-------
ENGINE: JT9D-7 Continued
S/V - 5 88 ft."I T rated: 4615° lbs"
V 21-1
PRODUCTION
IDLE POWER:
8.2 ' % rated
' T,
MODE
(Atm)
M,
(///r.) ft
-Mf
f/a
EI
l-nc
HC
CO
. IDLE
APPROACH
CLIMJJOUT
TAKEOFF
ADVANCED:
IDLE
Cd
G APPROACH
CLIMCOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMHOUT
TAKEOFF
3.95
8.5
18.5
21.1
Vorb:
8.35
1055
1320
1380
x comb
a
a
51.4
99.4
189.0
208.6
istor
UK
i
.226
.074
.021
.017
[fuel «
L *
.0100
-M30
.0194
.0215
taginj
1849
4648
13193
16142
) (Cat
jgory
: 3)
.044
nmi
.0002
0001
.0058
29.8
i n
0.1
o.os
1.70
77.0
Q f>
0.5
0.7
18.7
3.1
7 ft
21.4
7Q.4
'
*
* Estimated
-------
ENGINE: JT9D-70
c/r = 7.14 ft'1
w /
PRODUCTION
T rated: 51,150
IT : 24.35
IDLE POWER:
6.5% rated
P- T0 M M
3 3 a f
MODE Oitra) (°R) (j»/S) ft f/a f#/hrj
IDLE
APPROACH
CLJMiiO'JT
TAKEOFF
ADVA:;CED:
IDLE
APPROACH
£CLI:;;;GUT
^A'-OFF
ADVANCED:
IDLE
ArPXO/.CH
CLlMiVO'JT
TAKEOFF
1.74
9.66
21.29
24 3^
Air A
Aerat
820
079
L357
A1A
ssist 1
P
.ng No:
prc
45.5
108.3
211.4
?*&. 1
ozzle
same
as
oduct
zle C
camp
as
due tic
(Categ
on
mbusto
n
.011
.015
.021
.rm
ory 1)
r (Cat
1800
5850
L5980
9380
igory
l-nc
0
.0138
.0007
.0002
.0002
.0063
.0035
.0008
.. 0003
.0002
El
HC CO
6.8
0.45
0.15
0.15
1.3
same
as
jroduc
0.75
0.31
0.20
0.18
34.0
1.3
0.3
0.2
22.3
:ion
12.4
2.2
0.45
0.3
NO
X
3.2
8.1
24.2
31.0
3.4
5.1
8.0
17.5
21.3
-------
ENGINE: JT9D-70 continued
S/V = 7.14 ft'1 T rated: 51,150
* : 2'4.35
PRODUCTION
IDLE POWER:
6.5' % rated
P, ' T, M M.
3 3 a f
MODE (qtra) (°R) (///S) ft f/a r///hr^ l-nc
IDLE
APPROACH
CLT.MJ50UT
TAKEOFF
ADVANCED :
IDLE
M APPROACH
£ CLIMEOUT
TAKEOFF
ADVANCED:
IDLE .
APPROACH
CL1MBOUT
TAKEOFF
3.74
9.66
21.29
24.35
Stage
820
1079
1357
1414
d Prem
sa
a
prnrf
45.5
108.3
211.4
234.1
Lx (Ca
le
j
iot- inn
:egory
.011
.015
.021
.023
3) - E
1800
5850
15980
19380
stimat
;d
nn8
.0007
.0002
.0002
.0019
.0101
.0025
El-
HC CO
fi 8
0.45
0.15
0.15
.73
5.88
1.18
0.20
q&.n
1.3
0.3
0.2
5.6
21.5
6.0
0
NO
X
1 7
8.1
24.2
31.0
3.5
7.5
14.3
16.1
n -v 98% at
^60% power
^ 2+ % SFC Penalty
-------
ENGINE: JT10D
S/V =
PRODUCTION
T rated: 22-24,000 Ibs.
POWER.
' % rated
MODE
P, ' T. M
3 3
(Atra) (°R) (tf/S) ft
El
f/a
HC
CO N°x
- IDLE
APPROACH
CLIHBOUT
TAKEOFF
ADVANCED:
IDLE-
₯ APPROACH
00 CLIMCOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CL1MBOUT
TAKEOFF
,
..
.
\
i
,
\
>.
.
.'
, '
'
*
*
.
1
^
.
'
.
-
-------
ENGINE: JT3D-7
"1
=7.14 ft
PRODUCTION
T rated: 19000 Ibs-
IT : 13.5-;
IDLE POWER:
5' % rated
. P3 ' T3 Ma - . -HE .
MODE (Atm) (°R) (0/S) ft f /a (#/hr>_ l-nc
IDLE
APPROACH
CL1MJ50UT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMUOUT
TAKKOIT
ADVANCED:
IDLE
APPROACH
CL1MBOUT
TAKEOFF
2.04
5.58
11.84
13.61
Smoke!
Aerat
650
911
1133
1185.8
ess bu
ne noz
sajae
as
Produc
33. »9
91. -13
168 M
186. 8f
rner c
zle
:ion
..
.5799
.1572
.0497
.0389
m, (C
.0083
.0094
.0135
.0148
ategor
.
1013
3084
8188
9956
= 2>
\
i :
\
.
.'
,
"
.1399
.0064
.nnn«
.0006
.0818
.0034
.0003
.0002
El-
HC CO
123
2.1
0.4
0.5
69.1
0.5
0.1
0.1
139
19,5
1.9
0.9
91.9
12.8
1.1,
0.5
.
NO
X
2.2
5.3
9.6
12.7
3.3
8.0
14.4
19.1
-------
ENGINE: JT8D-17
S/V =7.69 ft"1
PRODUCTION
T rated: 16000 Ibs
» I 17.0 .
IDLE POWER:
6-5 % rated
MODE
P ' T
F3 T3
M
El
(qtra) (UR) (///S) ft f/a (#/hr)
l-nc
HC
CO
NO
IDLE
APPKOACH
CLlMiJOUT
TAKFOFF
ADVANCED:
IDLE
APPROACH
CLIMBOUT
TAKEOFF
ADVANCED:
IDLE
APPKOACH
CLIMBOUT
TAKEOFF
2.45
6.60
14.76
17.01
Air as
Pro
Aerat:
Pro
08
57
1212
1281
sist ((
same
as
ductio
ng noz
same
as .
luctio
26.8-
68.5
133.2
146.0
ategoi
i
:le coi
,
i
.533
.144
.041
.031
y = l)
ibustoi
.0119
.0114
.0165
.0188
(Gate
,
1 1 SO
2810
7910
9980
»ory =
2)
. \
'
.
\
.
.'
,
' '
'
.0156
.0021
.0003
.0002
.0072
.0046
.0009
.0006
.0007
8.78
0.50
0.05
0.05
2.3
0.5
0.05
0.05
1.93
0.50
0.30
0.29
34.0
7.2
L.O
).7
22.5
7.2
1.0
0.7
12.7
,2.1
-1.4
1.7
3.4
6.9
15.6
20.3
-6.9
15.6
20.3
2.8
7.9
17.7
.20.1
'
_
w
-------
ENGINE) JT8D-17 Continued
S/V - 7.69 ft."1 T
PRODUCTION
V
1600Q lbs'
IDLE POWER:
6.5 ' % rated
17.0
MODE
IDLE
APPROACH
CLIMI50UT
TAKEOFF
ADVANCED:
IDLE
eo
£ APPROACH
CLIMBOUT
TAKEOFF
ADVANCED :
IDLE
APPROACH
CLIMBOUT
TAKEOFF
P3
(Atra)
, ^
6.60
14.76
17.01
Vorb:
' T3
(°R)
70fi
957
1212
1281
x (fue
pr
M
(///S)
26 8
68.5
133.2
146.0
L stag
5PT71P
as
)ducti
ft
.sn
.144
.041
.031
ing) ((
>n
f/a
.miq
.0114
.0165
.0188
ategoi
Mf
(#/hr)
11 SO
2810
7910
9980
y = 3)
\
l-"c
0^56
.0021
.0003
.0002
,0049
HC
8.78
0.50
0.05
0.05
.06
El
CO
u.o
7.2
1.0
0.7
18.7
4 Q
7.8
6.8
N0x
3.4
6.9
15.6
20.3
2.6
<; 7
9.1
11.5
.
"
-------
ENGINE: RB211-22B
S/V =
PRODUCTION
T rated: 42000 Ibs
ir : 25.6V
IDLE POWER:
4.75% rated
P3 ' T3 Ma - Mf
MODE fa'tra) (°R) (///S) ft f/a (#/hr) ' l-°c
IDLE
APPROACH
CLlMiJOUT
TAKFOFF
ADVAIICED :
IDLE
H APPHOACH
JS CLIMBOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMnOUT
TAKEOFF
25
Unkn
wn (C
.
i
ategor
-
<
1.4.
0.1
r 1-2)
1
1474
4499
12644
15140
\
i
.
\
.
.'
t
'
.106
..011&
.0010,
,
.0091
.001C
.0005
.0004
El
HC CO
92.9
_.7.8
0.3
0.4
3.2
0.4
0.2
0.2
1Q6__
21.4
2.1
0.4
27.2
2.8
1.3
1.1
.
NO
X
2.2
6.5
23.8
31.3
i
3.1
9.8
28.5
32.8
«
-------
ENGINE: RB211-524
S/V =
PRODUCTION
T rated: 50000 .Its-
tf ! "28 :;
c
IDLE POWER:
6% rated
P ' T
*3 13
M
M
MODE (qtra)
a - f
(///S) ft f/a
El-
l-"c HC
CO
IDLE
APPROACH
CLIMiiOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMiiOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMiiOUT
TAKEOFF
28
Unkni
wn (Ca
.
tegory
i
0.33
0.043
= 1-2
,
\
i .
,
«
' .
.'
(
*
.0077
.0008
.0005
.0004
2.8
0.3
0.3
0.2
*
22.5
2.3
1.1
1.0
1
3.4
10.6
31.2
36.2
,
'
w
u>
-------
ENGINE: Spey 511
S/V =
PRODUCTION
T rated: 11400 Ibs.
ir : 19.0.-J.
IDLE POWER:
6 % rated
P, ' T, M M,
3 3 a - f .
MODE Oitra) (°R) (0/S) ft f/a (#/hr) l-nc
IDLE
APPROACH
CLIMBOUT
TAKF.OFF
ADVANCED:
IDLE '
APPROACH
CLIMBOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMBOUT
TAKEOFF
Unkno
wn Cat
.
-.
.
agory '.
..
0.68
--2)
.
6950
. \
I
,
\
.
.'
t
*
.126
.0064
.0007
.0005
.0191
.0086
.0030
.0005
El
HC CO
116
2.1
0
0
10.2
4.5
2.3
0
*
105.6
19.7
2.7
2
44.0
20.0
4.1
2.3
N0x
1
1.6
9.9
18.0
20.1
.
.
'
-------
ENGINE: Spey 555
S/V =
PRODUCTION
P3 '
T rated: 9850 Ibs
M
IDLE POWER:
M
'6% rated
El
MODE (atra) (°R) (///S) ft f/a (#/hr) l-nc HC CO
IDLE
APPROACH
CLIMUOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
,1, CL1MEOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CL1M1JOUT
TAKEOFF
Urikno
m (Ca
'
.
.
.tegorj
..
1-2)
755
1655
4501
5516
. \
\
.
- .'
f
*
.0223
.0102
nns?
.0036
12.6
5.0
i f>
2.7
48.7
25.2
8.6
5.1
x ?
'
1.1
4.2
11.2
14.9
'
"
-------
ENGINE: Dart RDa7
S/V =
i
PRODUCTION
T rated: 2280 shp
IDLE POWER:
% rated
v5-75
MODE
P ' T
3 :3
(qtra)
M_
(///S) Ji f/a
M
El-
HC
CO N°x
IDLE
APPROACH
CLlMiiOUT
TAKEOFF
ADVANCED :
IDLE
APPKOACU
T CLIMBOUT
<" TAI'.l-OFF
ADVANCED :
IDLE
APPROACH
CL1MKOIJT
TAKEOFF
5.75
1530
,
-------
ENGIN'E: Tyne
S/V =
i
PRODUCTION
T rated: 5500 shp
IDLE POWER:
' % rated
V 13'5
MODE
P ' T
3 13
(a'tra)
M
M
El
ft f/a (#/hr)
l-"c HC
CO N°x
IDLE
APPROACH
CLIMBOUT
TAKFOFF
ADVANCED:
IDLE
APPROACH
₯ CLIMBOUT
"° TAKEOFF
ADVANCED:
IDLE
-
APPROACH
CLIMIiOtJT
TAKEOFF
13.5
2475
.
-------
ENGINE: TPE 331-2
S/V - 15.18 ft"1
PRODUCTION
T rated: 715 shp
tr : 8.5 V;
c
IDLE POWER:
5 % rated
MODE
IDLE
APPROACH
CLIMliOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMUOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CL1MBOUT
TAKEOFF
P.
(Atra)
2.70
7.21
7.98
8.13
Primar
2.79
7.10
7.94
8.01
' V
770
1050
1066
1070
/ only
769-
1043
1054
1060
M
(///S)
2.43-
6.0-5
5*97
5.68
at ta:
2.69
6.04
5.82
5.88
..
a
.369
.093
.074
.068
i idle
.385
.097
.075
.074
f/a
.0120
.0101
.0173
.0198
. (Cat
.0099
.0103
.0180
.0194
.
Mf
(#/hr)
105
220
372
. 405
egory
96
224
377
411
= 1)
. \
.
\
.'
, *
'
L-nc
.0947
.0062
.0007
.0006
.0099
.0066
.0005
.0004
HC
91.24
2.67
.41
.39
5.41
3.25
.14
.13
El
CO
54.13
16.61
1.36
95
52.20
L6.15
L.58
L.01
N0x
2.58
8.29
9.93
10.23
i
4.41
8.80
9.79
10.24
>
'
,
_
oo
-------
ENGINE: TPE 331-3
S/V = 15.18 ft'1
PRODUCTION
T rated:
ir : ' 10
840 shp
IDLE POWER:
5 % rated
P, T- M . ' M,
3 3 a f .
MODE (qtra) (°R) (///S) ft f/a (#/hr) l-nc
IDLE
APPROACH
CLIMiJO'JT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIMI50UT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CUMP.OUT
TAKEOFF
3.07
9.11
9.99
10.09
Primal
3.07
8.84
9.67
9.74
793
1122
1142
1145
y fuel
792.
1124
1142
1145
2.6L
7.31
7.. 15
7.62
nozzlf
2.50
-
7.67
7.15
7.15
..
.301
.065
.052
.054
s only
.418
.071
.055
.054
.0119
.0095
.0159
.0167
at ta
.0109
.0096
.0172
.0179
112
250
409
458
98
265
443
461
' (Cat<
\
.
-gory =
.'.
1 )
.
.
'
.084
.002
.0004
.0003
.0101
.0014
.0003
.0004
El
HC CO
79.11
.62
.14
.10
6.20
.21
.18
.29
61.53
6.94
.97
.77
19.97
5.06
.65
.63
N0x
2.87
9.90
11.85
12.36
4.41
11.09
13.16
13.36
.
.
^D ,
-------
ENGINE:
S/V =
PRODUCTION
T5321A
T rated: 1,800 eshp-
IT : 8.0 .-;
IDLE POWER:
7 % rated
MODE
IDLE
APPROACH
CLlMiJOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CLIM110UT
TAKK01T
ADVANCED:
IDLE
APPROACH
CL1MHOUT
TAKEOFF
P3
(atra)
2.6
4.81
7.75
8
' T3
(°R)
758
895
1072
1100
M
(///S)
6.0-
9.2
13.9
14.1
i
ft
1
f/a
.0130
.0148
.0207
.0217
Mf
(ff/hr)
1026
\
i
.
,>
.
.
(
'
l-nc
HC
t
El
CO
.
N0x
i
-------
ENGINE: T53-L-13A
S/V =
PRODUCTION
T rated: 1400 shp
IDLE POWER:
7 % rated
V
MODE
M.
(qtra) (°R) (///S) ft f/a (#/hr)
El
l-nc
HC
CO
NO
IDLE
APPROACH
CLlMiSO'JT
TAKEOFF
ADVANCED :
IDLE
K APPROACH
£ CLIMBOUT
TAKKOFF
ADVANCED:
IDLE
APPROACH
CLlMliOUT
TAKEOFF
2.38
4.39
7.07
7.3
733
868
1040
1066
4.7
7.2
10.8
11.0
..
.0130
.0148
.0207
.0217
,
220
380
775
800
\
'-
.. .
.
(
'
.0162
.0017
9.52
0.39
-
-
33.94
5.97
0.97
0.84
.
2.71
4.78
7.58
8.23
'
.
-------
ENGINE: LTP101
S/V = 15.00 ft'1
PRODUCTION
T rated: 620 eshp
ff : 8:4 V;.'
IDLE POWER:
4.6 % rated
- P3 ' T3 Ma - Mf .
MODE (4tra) (°R) (///S) ft f/a (#/hr) l-nc
IDLE
2X IDLE
APPROACH
CLlMiJOUT
TAKF.OFF
ADVAIJCLB:
IDLE
^ APPKOACII
u, CLIMI50UT
TAI'.KOIT
ADVANCED:
IDLE
APPROACH
CLlMliOUT
TAKEOFF
3.33
4.01
5.03
8.0
8.3
Inje<
Airf]
830
910
935
1059
1078
tor mo
DW dis
2.32-
2.78
3.26
4.. 64
4.78
liflca
-
ribut
..
ion ((
*
on (d
.0115
.0114
.0133
.0185
.0195
ategor
tegory
67
172
307
340
7 = D
- 2)
\
i
/.
.
.'
(
' '
.030
El
HC CO
19.2
i
45
. .
N0x
_
.
-------
ENGINE:
S/V = 9.76 ft
PRODUCTION
All 250B15B
-1
T rated: 400 shp
»-! 7.2 ''.'
IDLE POWER:
10 % rated
. P3 ' T3 Ma - Mf .
MODE (qtra) (°R) (///S) ft f/a (///hr) l-nc
IDLE
APPROACH
CLlMiSOUT
TAKF.OFF
AD VANG ED:
IDLE
APPKOACH
CUM BOUT
TAKEOFF
ADVANCED;
IDLE
APPKOACH
CL1MROUT
TAKEOFF
2.79
4.33
6.83
7.14
Free!
745
850
1010
1030
amber
1.55
2.27
3.26
3.35
;ombus
.-
.485
.265
.125
.114
:or; S/
.346
.189
.089
.081
.0112
.0142
.0207
.0220
V = 9.
.
63
85
245
. 265
LI ft
.^
(cate
. \
gory=
\
2)
' .
.'
(
' '
.0403
.0158
.0024
.0020
.0050
.0018
.0018
.0014
El
HC CO
20.18
5.18
0.38
0.25
2.00
0.23
0.060
.0436
*
97.35
48.6
9.0
7.83
14.01
7.0
7.4
5.76
.
N0x
1.45
2.18
5.95
6.59
1.58
3.40
5.52
5.81
.
-------
ENGINE:
S/V =
PRODUCTION
All 501D22A
T rated: 4400 shp
IDLE POWER:
ir
9.7
M
M
5 % rated
El
MODE (qtra) (°R) (///S) ft f/a (///hr) l-nc HC CO
IDLE
APPROACH
CLTMi'.O'JT
TAKEOFF
ADVANCED :
IDLE
APPROACH
CLIMIiO'JT
TAI'.KOFF
ADVANCED:
IDLE
APPROACH
CLIMUOUT
TAKEOFF
3.64
8.30
9.45
9.70
levers
Fuel
795
1059
1091
1099
e flow
Pi
stagin
15.0-
33.0
33.0
33.0
combu
same
as
oduct:
', (Cat
same
..
as
roducl
itor ((
on
'gory =
(
ion
. 0113
.0096
.0185
.0200
ategor
3)
,
610
1140
2198
2376
y = 2)
\
.
<
.
.
,
-
.026
.0029
.0013
.0007
.0014
.0009
.0004
.0003
.0017
.0009
.0013
.0005
17.6
1.96
0.89
0.28
0.2
0.3
0.2
0.1
.
0.36
0.58
0.40
0.21
43.6
5.10
2.06
2.04
5.1
2.6
L.I
L.I
5.85
L.67
i.26
L.49
N0x
3.53
7.49
9.22
8.88
3.9
5.8
10.8
11.0
5.15
6.88
6.48
.7.35
'
BE
-------
ENGINE I PT6A-27
S/V = 12.35 ft.
PRODUCTION
-1 T rated: 680 shp
6.5
IDLE POWERS
3.37
% rated
P, ' T., M Mf
3 3 a r
MODE (atm) (°R) (/>/G) ft f/a (#/hr) l-nc
IDLE
APPROACH
CLI.MiiOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CL1MBOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
CUMr.UUT
TAKEOFF
1.90
4.15
6.24
6.48
Impro
Delay
6.93
906
1010
1023
ved Pr
ed que
2,5
5.0
5.9
.mary ;
ich (C;
lone Mi
tegorj
^0107
.0112
.0183
.mRQ
xing.
2)
115
215
400
425
1
.059
onny
.0003
.0002
El
HC* CO*
50.2
9 9
0
P
64 J)
23 0
L-2
1.0
,
NO*
X
2.4
/i Q
7 n
7.8
*EPAPS calculated from this data do not agree with those from GAMA (Ref. 32) which are quoted in
Table III-l. The Els above were taken from Reference 20.
-------
ENGINE: PT6A-41
S/V = 10.75 ft'1
PRODUCTION
T rated:
ir : 8 . 0
900 eshp
IDLE POWER:
3' % rated
P3 ' T3 Ma Mf
MODE (qtra) (°R) (///S) ft f/a (ff/hr) ' l-nc
IDLE
APPROACH
CLIMHOUT
TAKEOFF
ADVANCED:
IDLE
APPROACH
C LI!-! BOUT
TAKL'01-T
ADVANCED:
IDLE
APPROACH
CL1MROUT
TAKEOFF
2.00
4.89
7.76
8.12
Airflc
2.11
3.71
7.25
7.52
628
837
982
1001
w dist
689-
845
1077
1097
3.53.
6.39
8-. 22
8.49
rlbuti
3.50
5.15
7.32
7.58
*
1.22
0.30
0.13
0.12
m (Ca
0.98
0.39
0.11
0.10
.0116
.0118
.0161
.0168
;egory
.0107
.0101
.0161
.0166
.
147
273
473
510
= 2)
134
188
'424
453
. \
i
.
\
.
(
' '
'
.12
.029
.003
.0028
.0059
.0016
.0004
.0004
El-
HC CO
101.6
22.7
2.02
1.75
1.78
.14
.03
0.0
115.3
34.8
6.5
5.1
18.65
6.29
1.72
1.60
.
N0x
1.98
4.65
7.56
7.98
3.73
5.17
8.80
9.31
.
I
LO
ON
-------
ENGINE: T-39
S/V =
PRODUCTION
T rated: 40 shp
ir :
c
POWER
ADVANCLD:
T
u>
-
ADVANCED :
P3
G\tra)
1 T
3
M
(///S)
ft
f/a
i
l-nc
HC
El
CO
.
N0x
-------
ENGINE: ST6
s/v _
T rated: 720 shp
v 7-9
PRODUCTION
*
P, ' T- M M
3 3 a f
POWER (qtro) (°R) (///S) ft f/a (#/hr) 1_1C
720 HP
ADVANCED :
w
U)
ADVANCED:
_
7.9 .
Airfl
1020
w dist
7.9
ributi
on (Ca
:egory
480
= 2)
1
El
HC CO
.
N0x
-------
ENGINE: TSCP700-4
S/V = 8.25 ft
PRODUCTION
-1
T rated: 772 HP
ir : 10.37;
P, ' T, M
3 3 a -
POWER Oitra) (°R) (#/S) ft f/a l-nc
639 HP
704 HP
. - 772 HP
ADVANCED:
-Rig '
Rig
ADVANCED:
.
6.96
7.16
7.77
Water
9.00
7.bii
992
1001
1028
inject
1105
lnu
.
6.7-5
7.00
7.29
ion. (C
9.71
8.00
.058
.056
.048
ategor
.043
.047
.020
.020
.020
' = 2>
.016
.018
,
' \
* "
.
\
i t
.'
,
*
'
.0005
.0006
.0005
. Q002
.0002
El
HC CO
.183 '
.244
.189
.053
.049
1.550
L.446
L.328
.737
.540
N0x
7.108
7.377
7.806
'
12.04
11.93
_ .
_
.
w
-------
ENGINE: GTCP660-4
s/v = 9.50 ft'1
PRODUCTION
T rated:
ir 3.5
1100 HP
POWER
958 HP
1045 HP
. 1141 HP
ADVANCED:
ADVANCED:
P3
(Atra)
3.05
3.24
3.40
' T3
(°R)
854
863
870
'
Mn
(ff/S)
.
13 .-.81
14.98
16.01
.. -
ft
.430
.428
.403
f/a
.019
.018
.018
,
Mf
(#/hr)
923
984
1024
. \
i
\
t ,
- .'
t
*
l-nc
.0025
.0025
.0022
HC
.224
.400
.305
El
CO
9.608
8.972
8.349
.
N0x
4.924
5.119
5.320
t
.
'
.p-
o
-------
ENGINE: GTCP30
S/V »
PRODUCTION
T rated: 78 HP
ff : 3'. 76.;
C
POWER
78 HP
ADVANCED:
ADVANCED:
; >.
P3
(ijtra)
2.43
2.48
2.52
' T3
<°R>
687
790
794
M
0?/s)
.
.947
.963
.954
i
ft
1
f/a
.0198
.0199
.0201
,
Mf
(#/hr)
67.5
6,9.0
69.0
' \
1 : '
.
\
.
.'
,
'
' '
l-nc
.0014
.0014
.0013
HC
.110
.139
.139
El-
CO
5.576
5.549
5.259
.
N0x
3.506
3.620
3.706
*
.
.
ta
-------
ENGINE: GTCP 36-50
S/V =
PRODUCTION
T rated: iqo
ffc: 3".25 .'
P ' T
F3 .3
M
a
El
POWER
79 HP
94 HP
103 HP
ADVANCED:
Rie
Rig
ADVANCED :
(Atra)
3.32
3.47
3.57
Airblj
(°R>
882
893
906
st (C?
'
923
926
(0/s)
.'
1.91
L.96
1.88
tegdry
2.35
1.77
'
t
ft
-
= 2)
f/a
*
.017
.018
.020
.0113
.0207
.
.
. \
\
.
.'
,
' '
l-nc
.0057
.0043
.0028
.0101
.0018
HC
.346
.141
.044
1.042
.041
CO
23.066
17.793
11.645
38.88
7.395
X
4.100
4.332
4.253
i
5.175
5.111
'
.
NJ
-------
ENGINE: GTCP85-98
S/V = 8.70 ft'1 .
PRODUCTION
T rated: 293 HP
ff 3.36;;
P ' T
* X
M
El
Power (qtra) (°R) (///S) ft f/a
l-'c HC
CO N°x
243 HP
277 HP
293 HP
ADVANCED:
160 HP
237 HP
ADVANCED :
3.30
3.60
3.74
Wide s
Spray
3.51
3.29
863
870
876
ngle f
cone
.
1000
990
.
3.9-7
4.26
4.36
el .
Categc
4.21
3.79
i
.409
.370
.350
ry = 1
.301
.310
.019
.019
.019
)
.0147
.0178
.
\
i
,
\
i ,
.'
t
1
.0023
.0022
.0020
.0025
.0019
.'182
.163
.145
.173
.077
9.164
8.87
7.972
9.809
7.876
5.411
5.593
5.856
5.73
5.780
-
-------
Appendix C
Glossary of Emissions Control Concepts
1. Fuel Atomizatlon - Pressure atomizer nozzle design changes can
lead to changes in the character of the fuel droplet size distribution.
This in turn affects the evaporation time and strongly influences the
amount of HC left unburned. To a lesser extent, the change in the
evaporation rates affect the local fuel-air mixture ratio and thus the
local temperature which would likely affect the CO and NOx levels.
Decreasing the flow number (equal to the fuel flow rate divided by
the square root of the injector pressure differential) reduces the
droplet size and in turn the evaporation time. This results in a hotter
local mixture which consumes the HC and CO. NOx, however, increases.
This approach is not universally profitable, however, as experiments
indicate that at low combustor inlet temperatures, no degree of atomization
will improve the droplet evaporation as 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.
2. Increase in Idle Speed - As engine power is reduced, HC and CO
levels generally increase as a result of lower temperatures and pressures
at the combustor inlet. By increasing the engine speed at idle, therefore,
the levels of HC and CO should be reduced. However, the level of NOx
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. This problem does not exist with turboprop (class P2) engines
as the thrust can be held nearly constant by properly varying the
propellor pitch with engine speed. Also, there is an attendant increase
in fuel consumption and noise with increased engine speed.
3. Airbleed - The use of airbleed is merely a means of increasing
the idle speed, particularly of a jet, without suffering the problems of
increased thrust. The additional power is directed towards compressing
and then dumping extra air instead of towards thrust. However, the use
of airbleed results in increased fuel consumption which offsets some of
the expected reductions in emissions (as with any increase in the idle
speed).
C-l
-------
4. 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 towards the fuel injectors
to help break up the fuel droplets and thus eliminate locally rich hot
spots. By eliminating the locally rich hot spots, NOx levels may be
reduced. The high velocity airblast effect is maintained at all power
levels especially at idle wherein the airblast technique (See 7) is
ineffectual. This technique is therefore, very effective in reducing HC
and CO emissions at low power and in low pressure ratio engines.
The results of experimental testing of combustors equipped with
air-assist systems indicate that substantial reductions in low power
levels of HC and CO emissions can be achieved with this technique.
However, NOx appears to increase, apparently due to a higher flame
temperature.
The use of air assist would have a large impact on aircraft hard-
ware systems because of the requirement to redistribute and compress
externally with an auxiliary compressor a part of the compressor air
flow although its effect on the combustor itself is minimal.
5. Fuel Sectoring - Fuel sectoring is a method to improve the
combustion conditions at idle and thereby improve HC and CO emissions.
In essence it is like fuel staging in that fuel is selectively removed
from part of the combustor and injected into another in order to achieve
a desirable fuel/air ratio. Specifically, at idle when the combustor is
burning quite lean and at a low flame temperature, the combustion effi-
ciency is poor, resulting in much HC and CO, because of inadequate heat
to vaporize the fuel and to stimulate the CO -> CO- 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 flow number of the nozzle is necessarily reduced,
improving atomization and (2) the fuel/air ratio is increased (richened)
so that a hotter flame exists, improving vaporization of the fuel and
enhancing the CO -» C0? reaction.
Hardware changes are minimal and involve only replacement of half
the fuel nozzle valves with ones designed to open at a higher pressure.
As such this control technique is particularly attractive for retrofit,
especially in view of its nearly universal,usefulness.
6. Nozzle Design - Pressure atomizer nozzle design changes in-
tended 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 angle spray cone reduces wall wetting
which in turn helps to reduce HC. Increasing the angle of a fuel rich
narrow angle spray cone is also effective in reducing HC and CO levels
since more complete combustion results from the better fuel distribution.
An optimum angle must be found empirically.
C-2
-------
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.
7. Airblast - An alternative to the air-assist technique (See 4)
is the airblast concept. This system is the same as the air-assist
system except for the method used in creating the high velocity air
around the fuel injectors. Here, the pressure differential that exists
between the compressor and the combustor is employed to achieve high
velocity air through a venturi system at the combustor inlet.
The basic concept is relatively simple since it requires only the
addition of the venturi tubes. However, alone it usually improves the
emissions only slightly. It proves to be necessary in most cases also
to optimize the airflow distribution of the liner which is the reason
this concept is placed in the second category of complexity.
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 may be expected that
NOx will increase as a result of the better combustion efficiency. At
low power especially in low pressure ratio engines, the pressure dif-
ferential across the injectors is reduced causing the air velocity
around the fuel injectors to be reduced. Therefore, the airblast effect
on fuel atomization tends not to be as effective at low 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.
8. Increased Combustor Length - Increasing the combustor length
increases the residence time of the reactants. This technique the-
oretically reduces the levels of CO and HC by allowing the reactions to
proceed further toward the goal of C0» and H«0. NOx, however, might
increase for the same reason. Results of testing this combustor technique
indicate that the CO and HC levels decreased as expected.
A consideration in using this technique is that the length of the
combustor casing would be increased, a potentially major change in an
engine which may be reflected as a change in the engine length and/or
center of gravity, both of which would affect engine installation.
9. Water Injection - Water injected into the primary zone of the
combustor results in a lower primary zone temperature. Test results
with water injection indicate that the lower primary zone temperature in
turn results in a 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.
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, however, water injection
results in increased thrust, and hence the payload possibly can be
C-3
-------
increased); (2) Higher fuel consumption is required to maintain turbine
inlet temperatures; (3) Precautions must be taken to prevent ice
formation in the water injection system for operation at ambient tempera-
tures 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 $0.30 per gallon) depending upon the location. Logistics
for demineralized water may be a problem also, especially for those
aircraft using smaller fields.
10. Rich Primary - Theoretically, reducing the primary airflow in-
creases the local fuel/air ratio and hence the primary zone temperature.
At low power, this is beneficial in that the higher temperature enhances
the CO > C0? conversion when 0 became available in the secondary and
aids in fuel droplet evaporation, thereby improving the consumption of
HC in the secondary. NOx may be expected to increase slightly. At high
power, the rich primary is very hot, producing large quantitites of NOx.
If the equivalence ratio is greater than one in the primary zone, smoke
becomes a problem, requiring complicated air flow patterns in the secon-
dary and dilution zones to consume it. Very limited results of testing
this combustor concept verify the expected CO reductions and the ensuing
smoke problem.
11. Lean Primary(l)*-There are two philosophies behind the lean
primary depending upon the pollutant to be controlled. The first idea
is not particularly difficult to implement and therefore it rates as a
category 2 (modest complexity) scheme. The other concept rates as a
category 2-3 scheme. If a primary zone is stoichiometric or rich (See
10), then there is possibly considerable CO emission because there is
insufficient 0 to complete the reaction to CO ( CO + 0 -» CO ) . While
this reaction can be completed further downstream when more air is added
there is the danger that this air addition may either quench the reaction
altogether or may occur too late for the reaction to go to completion
prior to exhausting into the turbine wherein quenching will surely
occur.
The solution here is to lean the primary sufficiently (6, the
equivalence ratio, is = 0.9) so that there is adequate 0- yet enough
heat for the C09 reaction to occur at reasonable speed. This prevents
the formation or large amounts of CO which must somehow be consumed
later. Because of the availability of 0. and the high temperature, any
remaining HC can also be expected to be consumed. This would be most
noticeable if the original primary had been extremely rich (^ £ 1.5).
12. 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 CO. conversion to approach equilibrium and for unburnt hydrocarbons
to be consumed. The temperature in the intermediate zone should be,
however, low enough so that NOx formation rates are slow. Very limited
Number refers to one of two different degrees of control generally
recognized possible with this concept. Lean primary (2) is described
is described in 14.
C-4
-------
test data indicates that CO and HC are reduced, as anticipated, with
only marginal increases in NOx. The difficulty lies in adjusting the
air flow into the intermediate zone properly 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.
13. Premix (1)*- There are two philosophies behind the premix
burner depending upon the pollutant to be controlled. The first ap-
proach is not difficult to develop into a combustor (geometry perm-
itting) and therefore it rates as a category 2 (modest technology)
concept; the other, while more effective against all pollutants, is
quite complicated to handle and therefore rates a category 3-4 (See 16).
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 may be acceptable (slightly lean of stoichiometric in
current practice), 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 or deterioration burning, the
premixing can be compromised 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.
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 a category 3-4 scheme, as noted above. The
excessively rich premix zone used here can lead to carbon deposition,
however.
Implementation of this scheme into an existing design can be
difficult in that the rather major combustor modifications must normally
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 can be
made longer with the premix zone merely being tacked onto the existing
geometry (with some airflow adjustments).
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 (where the fuel is injected). 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. An emissions
penalty is paid, however, through increased CO, a result of the lower
temperature quenching the CO > CO- reaction. This tradeoff has been
witnessed in experiments. Used by itself, therefore, this concept is
generally unacceptable. If CO can be controlled by some other combustor
concept, then leaning the primary can be adapted for NOx reduction.
* Number refers to one of two different degrees of control generally recog-
nized possible with this concept. Premix (2) is described in 16.
**See also 11.
C-5
-------
As with the premix concept discussed below (16), a very lean primary
zone is subject to flame stability problems especially at low powers
when the primary ±a at its leanest condition. Also, the larger amounts
of airflow into the primary leave less air to be used further down the
combustor for cooling the liner or adjusting the temperature profile
into the turbine. This often leads to durability problems.
15. 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 to the pilot
stage, thereby allowing a much higher local fuel/air ratio that would be
possible if the fuel were distributed throughout the combustor. This
mixture is then able to burn hotter, enhancing the CO -> CO- conversion
and droplet evaporation (aiding HC burning).
At high power, the fuel is distributed between the two stages in
such a way so as to minimize the peak temperature. This aids in pre-
venting NOx production. Generally, though, this type of combustor also
incorporates air distribution features (See 10) or fuel preparation
features (See 16) which further aids low NOx levels. Staging requires
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.
16. Premix (2)*- Fuel and air are mixed in a prechamber prior to
entering the primary combustion zone. This premixing allows combustion
to occur at a much leaner condition where NOx formation rates are slower.
Experimental test results indicate that NOx reductions can be achieved
where there is sufficient combustor inlet temperature to vaporize fuel
droplets in the prechamber. This means that this technique is most
applicable to high pressure ratio engines, which produce the required
high combustor inlet temperatures.
With the premix concept careful attention must be given to the
prechamber 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 lean
primary zone combustor stability may be a problem leading to altitude
flameout and difficulties in relighting.
The premix concept requires a significant change to the combustor
liner geometry since the premix chamber must be included in the combustor.
In some cases this has led unavoidably to a longer overall combustor and
thus in the combuster outer casing.
17. Variable Geometry - Variable geometry provides control of the
primary zone such that stable efficient combustion with minimum emissions
occurs 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 stoichio-
metric at idle power for HC and CO reductions, and to be lean (but
*See also 13.
C-6
-------
stable) at high power for NOx reductions.
Very limited testing with this type of combustor has been conducted
and therefore not much data are available. The data do indicate that HC
and CO can be reduced with the variable geometry combustor. However,
NOx improvements have not been significant, owing to inadequate designJ
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-
pressor stators and variable turbine nozzle guide waves do exist.
18. Catalysis - Catalysis is a process by which a special substance,
usually a solid substrate, causes the acceleration of a chemical reaction
while not being permanently affected. Catalysis is often used on current
automobiles to limit the emissions of HC and CO by the placement of the
catalyst in the exhaust gas so that these pollutants, which are products
of incomplete reaction in the cylinders, can be consumed.
In an aircraft engine, however, the catalyst would be 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 HC 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.
C-7
-------
Appendix D
DERIVATION OF THE NOx TECHNOLOGY CURVE (PRODUCTION ENGINES)
The curve through the data points in Figure 5 (production engine
data) represents the level of technology achieved in current engines.
The curve is established from the technology lines in Figures 22, 25,
and 26 (the latter two being used to establish Figure 23) plus the
following assumptions and observations:
1) Where known, the climbout contribution to the NOx EPAP is
typically 56% of the total for Tl engines and 53% for T2 engines;
2) The climbout pressure ratio is roughly 92% of the rated pressure
ratio for Tl engines and 87% for T2 engines; and
3) The climbout TSFC is assumed to be equal to the rated takeoff
TSFC.
As an example, consider a 20,000 pound thrust engine (T2 class).
First, from Figure 26, the rated pressure ratio is 23.0. The climbout
pressure then is 20.01 (23.0 x .87), and the corresponding El is 24.8
lb/1000 Ib. fuel (Figure 25). The rated (=climbout) SFC from Figure 22
is .383 Ibm/lbf-hr and hence the climbout (85% rated thrust) fuel flow
is 6511 pounds per hour. For a 6% idle thrust, the impluse is 1776.7
Ib-hrs. The climbout contribution to the EPAP is then (6511 x 24.8 x
2.2 r 60) * 1776.7 = 3.33. Total EPAP is 3.33 * .53 = 6.28 for a 20,000
pound-thrust engine which falls right in the middle of the data. Actually
this is to be expected as the scatter of engine data points about the
mean curves on each of Figures 25, 25 and 22 is quite small. Figure 25,
in particular, is a modified Lipfert curve which represents an excellent
correlation. The rated TSFC variation shows excellent correlation
except for the engines of older design.
D-l
-------
Appendix E
NOx Calculations for Newly Certified T Class Engines
The data which forms the basis of this NOx projection comes from
the CF6-50C engine employing the GE/NASA double annular combustor con-
cept. Testing was done as a part of the NASA Experimental Clean Combustor
Program. In order to extrapolate this data from the demonstration
engine (CF6-50C) to the hypothetical engine, it is necessary to know
first the operating conditions of the hypothetical engine and second a
means of extrapolating. The extrapolation method is found in reference
59 wherein the double annular data is also found. The conversion formula
offered is:
EI(NOx)NCE = EI(NQx) [^NCE\ /Texit. NCE\ /Tinlet, NCE "Tinlet, CF6
CF6*PCF6,
(n = 0.2 on approach and 0.5 elsewhere)
The operating conditions at partial power are estimated by ratioing down
the known (postulated) conditions at full power according to ratios that
exist in current engines. Specifically, given the rated pressure ratio
(RPR) and the TIT (turbine inlet temperature), the following steps are
taken:
(1) From the rated pressure ratio, T. . at takeoff is calculated
(n = 0.89). inlet
c
(2) With T (TIT) specified and T 1 known at full power, the
rated SFC is calculated (Figure 45) and from the specified size of
the engine, the fuel flow rate (Mf) at full power is calculated.
(3) For the large engine at the other power modes, P, T. , T . ,
and M are estimated by employing the partial power to full power
ratios found on the JT9D-7, while for the small engine the same
procedure is followed using the TFE 731 ratios.
For the two engines postulated, this procedure yields the operating
conditions presented in Table E-l in which the CF6-50C conditions and
El(NOx) are also listed. The above conversion formula then gives the
El(NOx) at each power and for each NCE: This is also listed in the
table. As EPAP is computed by:
E El x Mf
EPAP = modes
£
, Thrust x time
modes
with units of the individual terms and magnitudes of the LTO cycle terms
consistent with the requirements of reference 66, calculation may proceed
directly.
E-l
-------
Table E-l
Mode
Demonstration Engine (CF6)
PR CIT* TIT EI(NOx)
M,
Large Engine
PR CIT TIT EI(NOx)
Small Engine
PR CIT TIT EI(NOx)
Idle 2.92 770 1550 3.0
Approach 11.7 1130 2040 8.9
Climbout 25.9 1420 2710 13.3
Takeoff 29.8 1480 2860 16.9
1050 6 930 1510 5.7
3360 16.1 1225 1940 10.8
9520 35.1 1530 2650 18.7
11650 40 1600 2860 24."/
140 3 745 1575 2.9
440 7.5 955 2020 5.8
1170 18.3 1265 2580 7.9
1310 20 1300 2680 9.3
*A11 temperatures are in degrees Rankine.
-------
Appendix F
Newly Certified P2 Class Standards - Computation
In the computation of the technology lines for the emissions per-
formance of newly certified P2 class engines, two engines are used, one,
small (5000 horsepower), and the other, large (50,000 horsepower). The
specifications of these engines are presented in Table III-7. In order
to compute the emissions it is necessary to know both the emissions
performance of the combustor and the fuel consumption at each of the
four power modes of the landing-takeoff (LTD) cycle. Furthermore, to
know the emissions performance, it is necessary to convert data from a
known application (i.e., operating conditions) to the application in
question. This means, of course, that it is necessary to know the
operating conditions for both the initial application and the one under
investigation (again, for each mode, as necessary).
50,000 horsepower engine
Table F-l summarizes the operating conditions and engine performance
of the large P2 engine under consideration. The takeoff (rated) pressure
ratio (PR) and turbine inlet temperature (TIT) are hypothesized (Table
III-6) and the air mass flow rate selected to supply the necessary total
power output. The fuel flow rate is based then on the TIT requirement.
For the intermediate powers, the pressure ratio is based upon scaling
the intermediate power pressure ratios from known engines, and the
combustor inlet temperature (CIT) is then determined by assuming an
adiabatic efficiency of 0.85. The air flow rate is essentially constant
for the flight modes as the compressor runs at a constant speed; at
idle, the air flow rate is reduced. The fuel flow rate is selected to
provide the necessary net power output specified by the mode (e.g.,
climbout is at 90% maximum rated power). The power specific fuel con-
sumption (PSFC) is then calculated by conventional means.
5,000 horsepower engine
Table F-2 summarizes the operating conditions and performance of
the small P2 engine (the lower limit). The takeoff (rated) pressure
ratio and turbine inlet temperature are hypothesized (Table III-6), and
the rated air flow rate selected to provide the necessary power, as
before. The intermediate power point numbers are found in the same
fashion as those for the large engine.
Emissions - HC, CO
The technology necessary for the control of HC and CO has already
been demonstrated in several engines (JT9D-70, JT8D-17, TFE 731-2) and
it forms the basis of the standards for the competing T2 class newly
certified engines (Section III, part D). The technology lines are given
F-l
-------
Mode
PR
CIT (°F) TIT (°F)
Table F-l
Large Advanced P2 Class Engine
It
air (Ibs/sec) fuel (Ibs/hr)
Power *(horsepower)
rated PSFC
Idle
Approach
Climbout
Takeoff
15
34.2
39
40
1200
1520
1580
1600
1780
2220
2770
2860
105
235
235
235
3,280
9,070
17,550
18,820
3,000
15,000
45,000
50,000
6
30
90
100
1.09
0.60
0.39
0.38
Mode
PR
5R)
Table F-2
Small Advanced P2 Class Engine
air (Ibs/sec) fuel (Ibs/hr)
Power *(horsepower) % rated
Idle
Approach.
Climbout
Takeoff
7.5
17.1
19.5
20
970
1240
1290
1300
1600
1990
2580
2680
11.0
24.2
24.2
24.2
350
980
1890
2030
300
1500
4500
5000
6
30
90
100
1.18
0.65
0.42
0.41
* After reduction gear losses
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on Figures 50 and 51. This is achieved by the use of aerating (air-
blast) or air-assist (external boost) nozzles with optimally adjusted
liner airflow distribution. Emissions index (El) data from those lines
must be converted from the operating conditions of the demonstration
engines to the conditions of the advanced P2 class engine in question
(Tables F-l, F-2). Fortunately, only idle need be considered, for as
reflected in equation (2) virtually all of the products of incomplete
combustion (HC and CO) produced from a low emissions combustor over an
LTO cycle will be produced at idle, the point of the most adverse
conditions.
Table F-3 below summarizes the predicted emissions index performance
of the two F2 class engines under investigation, as extracted from
Figures 50 and 51 using the idle PR data in Tables F-l and 2:
Table F-3
Large Engine:
P2 engine
PR EI(HC) El(CO)
15 0.2 11.8
Small Engine:
P2 engine
PR EI(HC) El(CO)
7.5 0.3 9.1
F-3
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Emissions - NOx
The expected technology for NOx control comes from the application
of such control techniques as staging, lean burning, and premix/prevapor-
ization such as are being demonstrated in rig and engine tests as part
of the present NASA Experimental Clean Combustor Program. Emissions
index data for the most successful development so far, the double-
annular staged burner (CF6-50 application) can be converted to reflect
the operating conditions of the advanced P2 class engines under consideration
here by the formula (reference 59):
El (N0x)2 = El (NOx), { ^
(n = 0.2 on approach and 0.5 elsewhere).
Table F-4 shows the results of this conversion. Note that all modes
must1 be considered unlike the case for the HC and CO emissions, for here
the pollutant is produced in significant quantities in all modes.
Predicted Emissions Performance
It is now possible to calculate the emissions performance of the
two hypothetical advanced P2 class engines. The emissions performance
is measured here by the EPA parameter (EPAP), the parameter with which
the EPA aircraft engine emissions standards are based (reference 66).
(For the P2 class the units of the EPAP are pounds of pollutant per LTO
cycle per 1000 horsepower x hours over the LTO cycle.) The formula is
simply
EPAP
El x Mf
E Power x time
modes
with units of the individual terms consistent with the specified units
of EPAP. The summation is over the four modes of the LTO cycle. The
Els are specified in Tables F-3 and F-4, the fuel flow rates and powers
in Tables F-l and F-2, and the times in mode in reference 66. For HC
and CO, the Els in the flight modes are assumed to be negligible. The
results are, for the small engine,
EPAP (HC) =0.1
EPAP (CO) =2.9 (5000 horsepower)
EPAP (N0x)= 5.9 ,
and for the large engine,
EPAP (HC) .1
EPAP (CO) = 3.6 (50,000 horsepower)
EPAP (N0x)= 13.4
From these points, the technology curves are generated. For the NOx,
F-4
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Mode
Demonstration Engine (CF6)
PR CIT* TIT EI(NOx)
Table F-4
Large Engine
PR CIT TIT El (NOx)
PR
Small Engine
CIT TIT El(NOx)
Idle 2.92
Approach 11.7
Climbout 25.9
Takeoff 29.8
770
1130
1420
1480
1550
2040
2710
2860
3.0
8.9
13.3
16.9
15
34.2
39
40
1200
1520
1580
1600
1780
2220
2770
.2860
17.6
25.3
23.1
24.7
7.5
17.1
19.5
20
970
1240
1290
1300
1600
1990
2580
2680
7.3
11.5
8.7
9.3
*A11 temperatures are in degrees Rankine
01
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curve in particular (Figure 59), the EPAP is taken to be constant above
20,000 HP given the same PSFC characteristics. For such engines, NOx
control techniques should be equally effective as surface to volume
(S/V) effects do not substantially affect the NOx emissions index.
F-6
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Appendix G
Calculation of NOx Trend Line for
Category 3 Technology
In order to reasonably predict the performance of fuel staging in
general terms for existing engine models (describing a trend line on
Figure 33, EPAP (NOx) vs. thrust), it is necessary to be able to extra-
polate the test data to expected values in the representative (or "trend")
engine in question. To accomplish this extrapolation, the best presently
known formula is given in reference 59:
/ x
(-|~f)
o.5 , v r CIT -CIT
EI(NOx) - EI x -~ x -~ x exp
ref AV Pref j A\__< / ~ ~"l 518 | (1)
where the subscript "ref" indicated the reference engine on which the
testing was actually performed and
P = combustor pressure
TIT = turbine inlet temperature
CIT = combustor inlet temperature (compressor discharge pressure).
From Figure 40 it is determined that the NASA GE double annular combustor
offers the greatest potential for NOx control of the three concept/engine
combinations shown (conclusion based upon the fact that it has the
lowest average slope, 3 El(NOx), over the pressure range tested). Other
3 (PR)
concepts were tested in each of these three engines in the course of the
NASA study, but this graph presents only the best (selected for further
development). The reference conditions then become for the takeoff
condition of the CF6-50C (reference 59):
P .. =29.5 atm
ref
TIT ,. = 2855°R
ref
CIT . = 1475°R
ref
El = 16.9
ref
The calculation proceeds as follows:
1. From Figures 28 and G-l, the rated pressure ratio (PR) and turbine
inlet temperature (TIT) are found for the "trend line" engine at
the rated thrust level in question.
The combustor inlet temperature (CIT) is calculated from:
III n !
I(PR) Y _ ! -+l (2)
₯-Lo
where T °° ~ 519°R and r\ is taken to be 0.89, a good average, especially
for T2 class engines. c
G-l
-------
3. This data is then used in equation (1) to give the rated or take-
off EI(NOx) for that thrust engine using the double annular combustor.
4. For the other modes (idle, approach, climb-out), which must also
enter into the calculation, the Els are found by ratioing the rated
EI(NOx) computed in step 3 according to the ratios experienced by
the CF6-50C reference engine in its modes.
5. Determine the fuel flow rate (Mf = TSFC x thrust).
6. Estimate Mf at the other modes by ratioing the TSFC according to
the average TSFC ratio of a cross section of production engines,
specifically,
(TSFC) mode . . (TSFC) mode ( ,
Mode (TSFC) take off*1 ' (TSFC) take off ^ '
Takeoff 1.000 1.000
Climb-out 0.967 0.970
Idle . 1.732 2.022
(Again, M,. , = TSFC . x thrust , ).
6 ' f mode mode mode
7. Calculate the EPAP by the formula,
E , EI(NOx) x M.; x TIM
EPAP(NOx) m°de f
m . (thrust level) x TIM
(TIM = time in mode)
For example, a 30,000 Ib-thrust engine would have, by Figures 28 and G-
1,
PR =23.4 atm
TIT = 2750 °R
Equation (2) then gives CIT = 1380°R and Equation (1) gives
EI(NOx) = 12.07 for takeoff. From reference 59 or Figure 40 the ratios
of the partial power Els to the takeoff El for the CF6-50C these engine
are,
El (mode)
Mode
Takeoff
Climbout
Approach
Idle
El (takeoff)
1.000
.787
.757
.178
Applying these ratios to the new 30,000 Ib-thrust engine given,
Mode EI(NOx)
Takeoff 12.07
Climbout 9.50
Approach 9.14
Idle 2.15
G-2
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Equation (3) is then employed.to compute the EPAP. The particulars
are presented in Table G-l. For Tl engines, the times in modes
will differ, of course, according to the specifications in FR vol.
38, No. 136, pp. 19088.
Mode
Idle
Approach
Climb-out
Takeoff
TOTAL
26
4-
2.2
0.7
1155
3360
9120
11100
Table G-l
El(NOx)
2.15
9.14
9.50
12.07
Ibs NOx
1.08
2.05
3.18
1.56
7.87
% power
6
30
85
100
impulse
780
600
935
350
2665
I Ibs NOx
EPAP = ,E impulse, =2.95
1 1000 J
G-3
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TURBINE INLET TEMPERATURE
vs.
RATED PRESSURE RATIO
- PRODUCTION ENGINES
3000
0)
M
d
P
(0
2500
-P
0)
H
XI
M
2000
10 20
Rated Pressure Ratio
Figure G-l
G-4
30
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