^oTENN/q^ ,
80-GT-177
THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS
345 E 47 St., New York, N.Y. 10017
The Society shall not be responsible for statements or opinions advanced in papers or
in discussion at meetings of the Society or of its Divisions or Sections. or printed in
its publications.
Discussion is printed only if the paper is published in an A SME
Journal or
Proceedings. Released for general publication upon presentahon.Full
credit should be given to ASME. the Technical Division, and the authorls).
R. P. Johnston
P. Ortiz
General Electric Co.,
Cincinnati, Ohio
Future Trends in Subsonic
Transport Energy Efficient
Turbofan Engines
Details of the NASA sponsored General Electric Energy Efficient Engine (E
3
)
technology program are presented along with a description of the engine, cycle and
aircraft system benefits. Opportunities for further performance improvement
beyond E
3
are examined. Studies leading to the selection of the E
3
cycle and con-
figuration are summarized. The advanced technology features, cycle and com-
ponent performance levels are also presented. An evaluation of the benefits of the
fully developed Flight Propulsion System (FPS) is made relative to the NASA
program goals by comparing the FPS with the CF6-50C where both are installed in
advanced subsonic transport aircraft. Results indicate that a mission fuel saving
from 15 to 23 percent is possible depending on mission length.
INTRODUCTION
After the 1973 oil emhargo, it became clear
that efforts to develop a more fuel efficient
air transport system needed to be accelerated.
An overall plan to implement this, called the
Aircraft Energy Efficient (AC[) program, was
developed by NASA. The Energy Efficient Engine
(E
5
) project is an important part of ACEE.
The current E
3
had its beginnings in
earlier NASA study contracts such as Studies of
Turbofan Engines Designed for Low Energy Con-
sumption (STEDLEC) (Reference 1), Studies of
Unconventional Engines Designed for Low Energy Con-
sumption (USTEDLEC) (Reference 2) and Energy
Efficient Engine - Preliminary Design and Integra-
tion (E-
3
- PDI) (Reference 3).
In addition, a
compressor study called Advanced Multi-stage
Axial Flow Core Compressor (AMAC) (Reference 4)
influenced the core engine configuration of the
current E
5
. The timing and major contribution of
these studies toward the F
5
is shown in Figure 1.
The StIDLEC study investigated the use of
advanced engine technology in the choice of
cycle and fan pressure ratios, materials and
turbine inlet temperatures. Direct and geared
(for high bypass flows) drive engines of up to
45:1 overall pressure ratio and turbine inlet
temperatures up to 2800°F (1538°C) were evalua-
ted in terms of increased fuel efficiency.
Contributed by the Gas Turbine Division of The American Society of
Mechanical Engineers for presentation at the Gas Turbine Conference &
Products Show, New Orleans, La., March 10-13, 1980. Manuscript received at
ASME Headquarters January 9, 1980.
Copies will be available until December 1, 1980.
E3 Cycle And Configuration Selection
NASA Progarns
1974
1975
1976
1977
1978
STEDLEC
L
1
I
Cycle and Technology
USTEDLEC
Two Spool
AMAC
10 Stage 23:1 Compressor
E'PD&I
E Proposal Cycle
E'CD&I
Minor Chan
FPS Cy
Figure 1
E
3
Cycle And Configuration Selection
US'IEDLEC was essentially an engine configu-
ration study in that gearing, triple spools, prop-
fans, turboprops and exhaust gas regenerators were
studied to see if there were fuel efficiency advan-
tages in any unconventional arrangements of turbo-
machinery. Direct and geared drive two-spool
engines and turboprops were identified as having
high potential for advanced fuel efficient engines.
The Ai
,
WAC study resulted in the selection of a
10 stage 23:1 pressure ratio core compressor coupled
with a 2 stage turbine for the core of the General
lilectric E
a
. This configuration was shown to offer
both fuel efficiency and economic benefits over
earlier core engine configurations studied.
Ie the most recent of the preceding E'
studies, F
3
- PDI (Reference 3
5
, the general confi-
3uration and cycle of the Genera, Electric F
3
was
selected.
she eye e utilized projectec alvances is
ges -
cle
Copyright © 1980 by ASME
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aerodynamic technology and the long duct mixed flew
engine configuration to produce an estimated 14.4`0
reduction in installed specific fuel consumption
(sfc) relative to a scaled CF6-50C at a maximum cruise
rating condition, Figure 2 is a cross section of
the engine resulting from this preliminary design
study. Table 1 is a comparison between this advan-
ced engine and the reference scaled CF6-50C.
Advanced E
3
PDI Study Engine
4
^r
Figure 2
Advanced E
3
PDI Study Engine
Engine Cycle Comparison
Ref. CF6-50C
Initial
FPS
(Scaled)
Engine
Installed Fn @ M=.8
lbs (kN)
8610 (38.3)
8610 (38.3)
10668 m (35K), MxCI Hot Day
Core Corr Flow - MxCI lb/sec (kg/sec)
118.7 (53.8)
120.0 (54.4)
Bypass Ratio - MxCI
4.2
6.8
Fan Pressure Ratio - MxCI
1.76
1.65
Overall Pressure Ratio - MxCI
32
38
ASFC - Uninstalled MxCr (Std) So
Base
-13.7
ASFC - Installed MxCr (Std) °/o
Base
-14.4
Table I
Engine Cycle Comparison
Other NASA sponsored material studies
(References 5 and 6), were also conducted during
this period to determine possible system benefits
of certain new materials installed in an advanced
technology transport engine. Several of these
concepts such as Near Net Shape - Rene' 95 powdered
metallurgy disks, directionally solidified turbine
blade alloys and ceramic HP turbine shrouds were
eventually incorporated into the General Electric
E.
The current General Electric E
3
work is being
conducted under Energy Efficient Engine - Component
Development and Integration contract NAS3-20643 to
NASA-Lewis Research Center with Mr. Neal T. Saunders
as the Lewis Project Manager. The purpose of the
contract is to develop and evaluate technology
advances identified in earlier studies. These
advanced technology features are then to be demon-
strated in a series of component tests culminating
in the running of a core engine and an Integrated
Core/Low Spool (ICLS) engine in 1982 to demonstrate
the complete engine system.
ENERGY EFFICIENT ENGINE PROGRAM
There are well-defined NASA goals for the current
E
3
program. In terms of a completely installed
Flight Propulsion System (FPS) on an advanced techno-
logy subsonic transport aircraft, the FPS is to show,
as a minimum, the following benefits relative to a
CF6-50C reference engine:
• 12% reduction in installed specific fuel
consumption (sfc)
i 5% reduction in direct operating cost (DOC)
S 50% reduction in sfc deterioration in service
Other goals are:
B Meet FAR 36 (March 1978) acoustic standards
with provisions for growth
• Meet Proposed EPA (1981) emissions standards
for new engines
The program is structured into four major technical
tasks as follows:
Task
I
-
Preliminary
Design of a fully
developed FPS
based on Task II,
III
and IV results
Task
II
-
Preliminary
and Detailed Design
and testing
of the individual
components
Task
III
-
Testing and
Evaluation of the core
engine
Task
IV
-
Testing and
Evaluation of the ICLS
The timing of the major elements of the various tasks
is shown in Figure 3.
E
3
Program Milestones
Schedule
1978
1979
1960
1981
1982
• Contract Received
•!FPS_ Preliminary_Desg
Rev
lem
♦
ID
A
• Fan Test
Ir`,^'
• Core Compressor Test, Sts. 1-6
• Core Compressor Test, Stga. 1-10
• Combustor Development Tests
• HP Turbine Air Test
• LP Scaled Turbine Air Test
J
• FADEC System Test
j
d
• Thermal Barrier Coating, FPS Decision
• Miser Test
AA
AA
• P
owe
red Nacelle Tests (Langley)
J
•
First Core Test'
• Second Core Test
• ICLS Teat
Figure 3
E
3
Program Milestones
FPS TECHNOLOGY GOALS
12% SFC Reduction
Table II is a comparison of the major cycle and
system parameters for the General Electric FPS and the
reference CF6-50C, A visual comparison of the two
engines is given in Figure 4. Besides the obvious
difference of a separate versus mixed flow exhaust,
the FPS utilizes a higher overall pressure ratio,
lower fan pressure ratio and a higher bypass ratio
in conjunction with higher component efficiencies.
2
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Comparison of E
3
to Reference CF6-50C
E3
SFC Improvement vs. CF6-50C (MxCr)
CF6-50C
FPS
Rio
A
SFC
Cycle Pressure Ratio, MxCI
32
38
•
Component Adiabatic Efficiencies
-4.1
Bypass Ratio, MxCI
4.2
6.8
•
Mixed Flow Exhaust
-3.1
•
Increased Cycle Pressure Ratio (20%)
-1.0
Fan Pressure Ratio, MxCI
1.76
1.65
•
Propulsive Efficiency (FPR-BPR)
-2.5
Turbine Rotor Inlet Temperature
•
Increased Turbine Inlet Temperature (..170`F) (94°C)
-1.5
SLS/86°F (30°C) Day T/O, °F
(
°C)
2445(1341) 2450(1343)
•
Cooling and Parasitic Flows
-1.0
35K (10668M)/.8Mp/
•
Flowpath Pressure Losses
-0.1
Std. Day MxCr, °C, °F (°C)
2000(1093)
2170(1188)
UNINSTALLED
A
SFC
-13.3
•
Reduced Isolated Nacelle Drag
-0.6
SFC, 35K (10668m)/.8M MxCr,
%
Base
-14.2
•
Integrated Aircraft Generator Cooling
-0.3
Fully Installed,
0
/0
Base
-14.6
INSTALLED
A
SFC IMPROVEMENTS
-14.2
(Nominal Cust. Bid.
&
HP)
•
Customer Bleed and Power Effects
+0.4
Weight, Installed Lb/(kg)
9860(4473)
9300(4218)
•
Regenerative El Fuel Heater
-0.8
(50C Scaled to E' Mxcl Thrust)
FULLY INSTALLED (Cost. Bleed
&
HP)
-14.6
Table
II
Comparison
Of E
3
FPS And CF6
-50C
Table
III
E
3
SFC Improvement vs.
CF6-SOC
Component Efficiencies
C
E
3
/Reference Engine Comparison
E
3
Engine
A
CF6-50C Reference Engine
(Scaled to E
3
MXCL Thrust)
Figure 4
E3/Reference Engine Comparison
The 14.6% projected reduction in sfc for a
fully installed (customer bleed, power extraction,
and ram recovery) FPS comes from many sources as
shown in Table III. Component adiabatic efficiency
improvements are the single largest source of sfc
improvements. Individual component improvements
are given in Table IV. The levels of improvement
were estimated by taking current technology levels
of component performance and comparing them with
the projected performance levels of the FPS with
FPS levels of aerodynamic loading.
The largest improvements were made in the fan and
fan hub regions. Fan tip speeds were set at the most
efficient levels that would provide adequate stall
margin and specific flow. The blade shrouds were
placed in the minimum performance loss position on
the blade. Fan tip clearance reductions from current
levels were possible due to the improved fan casing
deflection control achieved by use of stiffer, lighter
composites and structural integration into the fan
frame. To provide the required core supercharging, a
quarter stage booster was added and loading on the fan
hub reduced. A side benefit of the booster configu-
ration is that about 40% of the booster air is bypassed
into the fan duct resulting in removal of the blade
tip boundary layer air. from the core supply along with
debris that might enter the fan hub region. The
Comparison of E
3
FPS and CF6-50C
Component Efficiencies
35,000 Ft./.8 M Max. Cruise
(1°668 M)
Component
E3 A EFF.
Fan Bypass
+4.8 Pts.
Fan Hub (Booster)
+4.0 Pts.
High Pressure Compressor
-
Adiabatic
-
.3 Pts.
- Polytropic
+
.4 Pts.
High Pressure Turbine
+ .8 Pts.
Low Pressure Turbine
+1.1 Pts
Table IV
Comparison Of E
3
FPS And CF6-50C
Component Efficiencies
quarter stage operation also permits proper matching
of the booster to core air requirements by permitting
excess air to bypass the core entrance. This
eliminates any variable geometry bypass provisions
normally required with close coupling of booster and
core compressor.
The choice of the 23:1 pressure ratio 10 stage
compressor (Figure 5) had a significant effect on the
overall FPS configuration and fuel efficiency poten-
tial. Its short length permitted a stiffer, less
deflection prone engine to be designed with just two
major frames. In addition, the work extraction from
the core turbine reduces fan turbine inlet tempera-
tures. When compared to the 14 stage CF6-50C, the
projected polytropic efficiency of the E
3
compressor
is higher although the pressure rise is over 500
greater.
In addition to the attention given to reduction
of aerodynamic losses in all the components, a large
improvement in component efficiencies resulted from a
reduction in blade tip and seal clearance losses. The
reductions came about in three major ways:
• Matching of materials and thermal response
• Low deflection engine mounting system
• Active Clearance Control system (ACC)
_._
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The ACC system is employed over the last 5 stages
of the core compressor, the high pressure turbine and
the low pressure turbine. A schematic of the aft core
compressor ACC system is shown in Figure 6.
E
3
High Pressure Compressor
Figure 5
E
3
High Pressure Compressor
HPC ACC System
FADEC
Modulating Mixing Valve
Controlled
_
y
LPT Purge
Air
Case Cooling Air
Clearance Circuit
Bypass Air
70Mfr^^^^^^f^
Stage 5
Bleed
Figure 6
High Pressure Compressor ACC System
In operation, a modulating valve varies the
amount of cooling air permitted to pass along the
outer surface of the aft inner casing. This modula-
ted cooling varies the radial expansion of the casing
and can then alter the running clearances of the
blades and vane shrouds. The modulation itself is
governed by the engine control. During periods of
higher than normal engine deflection or transient
tip clearance closure, the casing is not cooled
and thereby, becomes hotter and expands. This combi-
nation of heating or cooling allows engine build-up
clearances to be minimized and reduces excess running
clearances during climb and especially cruise.
The ACC system for the turbines is similar in
operation except that controlled fan air is allowed
to impinge directly onto the turbine cases. Opera-
tion of the ACC on a typical turbine stage is shown
in Figure 7. Table V illustrates the expected
performance benefit for each component due only
to the ACC system, The gains are substantial,
especially for the core turbine. A second major
benefit of the ACC system is that deterioration due
to inadvertent tip rubs will be reduced since the
clearances can be opened up during periods of high
maneuver loads or nacelle aerodynamic loads.
Active Clearance Control Operation
Representative Turbine Stage
Uneooted
-
Casing
Diameter
punning Clearance
E
...............................
I
RunnIng
^
CbMM^iFo
Maximum ACC
Acf Cle
ns
v
Capability
Co t
le
t
...
This Diameter
Rotor
Tip
Diameter
L
Ttkeorl
TT
1Climb -
Man. Cr.-.-.. Man. Cr.
100
Time, secs
1000
10.000
Figure 7
Active Clearance Control Operation -
Representative Turbine Stage
Estimated Active Clearance Control
Performance Improvement
Eff-0/a
SFC - %
• HPC
.5
.3
• HPT
1.6
1.0
• LPT
.4
.2
Total
1.5
Table V
Estimated Active Clearance Control
Performance Improvement
The other significant contribution to FPS fuel
efficiency is the mixed flow exhaust system. The
core exhaust is mixed with fan air by a mixer
(Figure 8) to produce additional thrust. Besides
improving overall engine efficiency, the mixer also
provides these benefits:
0 Core thrust spoiling during reverse mode
• Reduction of jet exhaust velocity and noise
A mixing effectiveness goal of 75% at maximum
cruise thrust has been established for the FPS. Scale
model testing is in progress and results, to date,
indicate achievement of approximately two-thirds of
the projected 3.15 cruise sfc improvement.
The propulsive efficiency improvements over the
CF6-50C are the result of the higher bypass flow
ratio and the lower fan pressure ratio, At maximum
climb, for instance, the fan bypass ratio of the FPS
is 60% higher than that of the CF6-50C, The increase
in propulsive efficiency coupled with the increased
pressure ratio and cruise turbine inlet temperature
produces a 5% reduction in sfc as compared to the
CF6-50C. Currently, the uninstalled FPS sfc, as
shown in Table III, is estimated to be 13.3% lower
than the uninstalled CF6-50C,
4
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Mixer and Rear Frame
-.
-1
m 4
Mixer
-
-^'"--
i
Centerbody
Rear Frame
Figure 8
Mixer and Rear Frame
Nacelle drag improvements over the CF6-50C were
accomplished by reducing the maximum nacelle diameter,
relative to the fan diameter, increasing the nacelle
slenderness ratio and reducing the frontal area. The
nacelle diameter was reduced by integrating the fan
casing and frame directly to the outer nacelle walls
and through extensive use of lighter and stiffer
composite materials. Frontal area was also reduced
by installing the accessory gearbox within the core
cowl volume instead of the fan case.
When the reduced nacelle drag and benefits due
to elimination of fan air cooling of current
technology constant speed drives for the FPS
Variable Speed Constant Frequency (VSCF) aircraft
generator are combined, the FPS installed sfc bene-
fit relative to the CF6-50C is 14.2 percent, as
shown in Table III.
If a fully installed FPS (customer bleed, power
extraction and ram recovery) is considered, an
advanced fuel heater/regenerator system increases
the net sfc benefit to the 14.6% shown on Table III.
A schematic of the fuel heater/regenerator as
installed on the FPS is given in Figure 9.
Fuel Heater/Regenerator
To Aircraft ECS
ECS
Precooler
Fan Air Overboard
Fan Air Inlet
Fuel
Regenerative
Heater
Heated Fuel
To Engine
Customer
ECS Bleed
Intermediate
Fluid
Figure 9
Fuel Heater/Regenerator
The regenerator takes advantage of the heat in
the Environmental Control System (ECS) air that is
normally lost to the engine cycle. By transferring
the excess heat to the fuel, low grade heat is added
to the engine in the thermodynamically most desirable
location, the combustor. Also, the current require-
ment for fan air to cool the ECS air is reduced, and
at most mission power settings, eliminated. Table III
showe +0.4% sfc penalty for E'
S
relative to the
CF6-50C for customer bleed at constant thrust. This
penalty is exceeded by the benefits of the regene-
rator.
The individual control functions that must be
maintained for the FPS to achieve fuel efficient
operation through wide variation of altitude and
thrust have been increased significantly as shown
in Figure 10,
Control System Outputs
Compressor Clearance
HP Turbine
Valve P osition J
Clearance
Start Bleed
Reverse
Valve Position r LP Turbine
Valve Position
Position
Clearance
\
Valve
r
Pos tion
'^ t
^.
/}
.++'R/
Thrust
Fuel Flow
/
8 Flow S lit
Core Stator
Vane Position
JJ Functions Not on CF6
Figure 10 Control System Outputs
Because of the number of controlled functions
required, increased power management complexity,
and more convenient aircraft interfaces, a Full
Authority Digital Electronics Control (FADEC) has
been selected for the General Electric E
3
, A
schematic of the FADEC control function, Figure 11,
illustrates the initial concept of reliability
through the use of an active standby FADEC. As
experience with and reliability of the FADEC grow,
more economical methods of ensuring essential
reliability would be utilized. Other control
functions, not now envisioned, could also be
added due to the inherent ability of a digital
control to be programmed to accept new duties.
Full Authority Digital Control
Stator
Actuation
Inputs
Power
Supply
Active
Clearance
En ine
Primary
Valves
Sensors
FA DEC
Start
Active
Bleed
Standby
H
valves
Aircraft
FADEC
Commands
Thrust
Reverser
Condition
Monitoring
Figure 11 Full Authority Digital Control
5
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