Performance of basic XJ79-GE-1 turbojet engine and its components

MISSING IMAGE

Material Information

Title:
Performance of basic XJ79-GE-1 turbojet engine and its components
Series Title:
NACA Research memorandum RM E58C 12
Physical Description:
56 p. : illus., tables. ; 27 cm.
Language:
English
Creator:
Campbell, Carl E
United States -- National Advisory Committee for Aeronautics
United States -- National Advisory Committee for Aeronautics
Publisher:
National Advisory Committee for Aeronautics
Place of Publication:
Washington
Publication Date:

Subjects

Subjects / Keywords:
Airplanes -- Turbojet engines   ( lcsh )
Genre:
bibliography   ( marcgt )
non-fiction   ( marcgt )

Notes

Bibliography:
Bibliography: p. 21.
General Note:
Cover title.
General Note:
"Second printing, for non-military distribution."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003852967
oclc - 11715579
sobekcm - AA00006132_00001
System ID:
AA00006132:00001

Full Text
'tc 4r







/1-7 / -.*1



NACA RM E58C12


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


RESEARCH MEMORANDUM


PERFORMANCE OF BASIC XJ79-GE-1 TURBOJET ENGINE AND ITS COMPONENTS

By Carl E. Campbell


SUMMARY

An investigation to determine the performance of the XJ79-GE-1 tur-
bojet engine and its components, while operating as integral parts of the
engine, was conducted in an altitude test chamber at the NACA Lewis lab-
oratory. Data were obtained over a range of Reynolds number indices
from 0.60 to 0.08 and for various settings of the variable compressor
stators and variable-area exhaust nozzle from fully open to fully closed
positions.

Compressor performance and turbine performance are presented in the
form of performance maps at selected values of Reynolds number index;
the effects of Reynolds number on performance are summarized. The ef-
fects of variable stator angle and high inlet-air temperatures on com-
pressor performance are also shown. Combustor performance is given in
generalized form as a function of the usual combustor parameters. Ex-
haust system data are presented to permit the calculation of over-all
engine performance from pumping characteristics. Maps of engine pumping
characteristics are presented at selected values of Reynolds number
index, and the general effect of Reynolds number on the pumping charac-
teristics is summarized. Over-all engine performance (net thrust and
specific fuel consumption) is presented for a flight Mach number of 0.9
at rated engine conditions over a range of altitudes to illustrate per-
formance losses resulting from decreased Reynolds number index. All
component and engine performance data are presented in tabular as well
as graphical form.


INTRODUCTION

An investigation to determine the performance of the XJ79-GE-1 tur-
bojet engine and its components while operating as integral parts of the
engine was conducted over a range of Reynolds number indices in an alti-
tude test chamber at the NACA Lewis laboratory. This engine incorporates
variable inlet-guide vanes and variable stator vanes in the first six
compressor stages as a means of avoiding part-speed surge. The engine
also has an afterburner and iris-type variable primary and secondary








NACA RM E58C12


exhaust nozzles; however, during the investigation reported herein, the
afterburner was inoperative and the secondary nozzle was removed. The
variable-stator system and the variable exhaust nozzle are normally
scheduled automatically by a combination electronic and hydraulic con-
trol, but they were manually positioned during this investigation.

The performance data were obtained over a range of Reynolds number
indices from 0.60 to 0.08 with the variable stators in the open position
and over a range of stator positions from 00 to 350 at a Reynolds number
index of 0.20. At each Reynolds number index and stator position, data
were obtained over a range of engine speeds at five exhaust-nozzle areas.
Data were obtained at engine inlet temperatures up to 7000 R, but the
bulk of the data was obtained at an inlet temperature of approximately
4160 R.

Component performance data are presented over a range of Reynolds
number indices and variable stator positions. Generalized engine data
are presented in a form that permits computation of engine performance
at operating conditions other than those specifically investigated, and
the method of such computation is illustrated. All component and engine
performance data obtained during this investigation are presented in
tabular as well as graphical form.


APPARATUS

Engine

The XJ79-GE-1 turbojet engine has a length of 207 inches and a
maximum diameter of 32.6 inches at the turbine section. The frontal
area based on the compressor tip diameter is 4.89 square feet. The dry
weight of the engine and its accessories is about 3150 pounds. The
manufacturer's static sea-level military performance rating (nonafter-
burning) is 9600 pounds of thrust with a specific fuel consumption of
0.87 pound per hour per pound of thrust at an engine speed of 7460 rpm
and a turbine-outlet temperature of 10700 F.

The XJ79-GE-1 turbojet engine has several minor airflow bleeds that
are used for cabin pressurization, anti-icing, turbine cooling, and
bearing-seal pressurization. These bleed flows are extracted from the
main engine airflow at the seventh and ninth compressor stages and at
the seventeenth compressor-stage seal. The amount of bleed flows dumped
overboard during this investigation did not exceed approximately 1.5
percent of the inlet airflow; the remainder of the bleed flows reentered
the main stream before reaching the afterburner diffuser section.







NACA RM E58C12


Engine Components

Compressor. The seventeen-stage axial-flow compressor has variable
inlet guide vanes and variable stator blades in the first six stages that
are moved simultaneously from the open position to their respective
closed positions. The angle of travel from open to closed for the vari-
our stages is as follows: inlet guide vanes, 470; first stator stage,
44u; second through fifth stages, 350; and sixth stage, 410. All ref-
erences to stator position throughout the report will be in terms of the
second through fifth stages as a matter of convenience; that is, the
closed position will be referred to as 350. The compressor has a con-
stant tip diameter of 29.95 inches through the first fourteen stages and
tapers down to a tip diameter of 29.3 inches at the seventeenth stage.
The hub-tip radius ratios of the first, fourteenth, and seventeenth
stages are 0.36, 0.86, and 0.88, respectively. The compressor was de-
signed to deliver an airflow of 162 pounds per second and a total-
pressure ratio of 12.2 at static sea-level military conditions.

Combustor. The combustor is a cannular type with ten circular
through-flow inner liners. Fuel is supplied to each liner through a
single-inlet duplex fuel nozzle. Ignition is provided by a spark plug
in one of the inner liners and spreads to the other liners through
interconnecting crossfire tubes. The combustor-inlet reference velocity,
based on the full burner section area of 4.33 square feet, is approxi-
mately 89 feet per second at design sea-level conditions.

Turbine. The three-stage impulse-type turbine has a constant
pitch-line diameter on all three stages and tip diameters of 28.4, 29.65,
and 31.05 inches for the first, second, and third stages, respectively.
The hub-tip radius ratios of the first, second, and third stages are
0.81, 0.70, and 0.59, respectively. The increase in annular area through
the turbine occurs entirely in the turbine nozzles. The turbine was de-
signed to operate at a turbine-inlet temperature of 17000 F at 7460 rpm.
Damping rods are installed between adjacent third-stage rotor blades to
reduce blade vibration. These rods are 3/16 inch in diameter and are
situated at approximately 75 percent of the blade height.

Control and exhaust system. The engine control schedules the
variable-stator assembly to vary continuously as a function of corrected
engine speed from the closed position (350) at 64 percent of rated cor-
rected speed to the open position (0) at 90 percent of rated corrected
speed (fig. 1). During the investigation, the original schedule was
altered as shown by the dashed line of figure 1 to achieve a higher
thrust during operation at a high engine-inlet temperature corresponding
to a flight Mach number of 2.0. The primary exhaust nozzle, which is a
convergent, variable-area iris-type nozzle, is scheduled to vary in
gradual steps from an open position at idle conditions to a closed








NACA RM E58C12


position at military position of the power lever. During this investi-
gation, however, the variable-stator assembly and the exhaust nozzle
were manually controlled. The afterburner, which was inoperative during
this investigation, has a maximum internal diameter of about 34 inches
and includes a diffuser, fuel-injection bars, a pilot burner, a three-
ring gutter-type flameholder, and a corrugated and louvered cooling
liner. The variable secondary exhaust nozzle was removed during this
investigation.


Installation

A view of the XJ79-GE-1 turbojet engine installed in the altitude
test chamber is shown in figure 2. The engine was rigidly mounted on a
flexure-plate supported test platform that was connected by a linkage to
a calibrated null-type thrust cell. Dry refrigerated or heated air
entered the engine inlet through a bellmouth Venturi duct, which was
mounted to the engine inlet and test platform. The inlet section is
separated from the exhaust section by the front bulkhead, which incor-
porates a labyrinth seal around the inlet Venturi duct to prevent the
flow of combustion air directly into the exhaust section and to permit
the measurement of thrust forces. The inlet- and exhaust-pressure con-
trols are designed to maintain automatically a constant ram-pressure
ratio and exhaust pressure.


Instrumentation

Instrumentation for measuring pressures and temperatures was in-
stalled at various stations through the engine as shown in figure 3.
The table presented on the figure indicates the number and type of meas-
urements at each station. Total-pressure and temperature probes at each
station were located at the approximate centers of equal annular-area
increments so that measurements could be averaged arithmetically. In-
strumentation was also provided to measure the portion of bleed flows
dumped overboard through the compressor-discharge standpipes.

Pressures were measured by null-type diaphragm capsules and recorded
by a digital, automatic multiple-pressure recorder. Temperatures were
measured and recorded by iron-constantan and Chromel-Alumel thermocouples
in conjunction with self-balancing potentiometers. Fuel flow was meas-
ured by a calibrated turbine-type flowmeter. The variable stator posi-
tion and primary exhaust-nozzle area were determined from cold calibra-
tions of output voltages from linear potentiometers in their actuating
mechanisms.








NACA RM E58C12


PROCEDURE

Most of the data were obtained at the minimum inlet-air temperature
consistently available (approximately 4160 R) to extend the range of cor-
rected engine speeds. Engine-inlet pressures were selected in conjunc-
tion with this inlet temperature to give a Reynolds number index range
from 0.60 to 0.08. Some data were obtained at higher inlet-air tempera-
tures up to 7000 R at a constant Reynolds number index of 0.4 in order
to investigate the effect of temperature itself on the reproducibility
of the data.

With the variable stators in the open position, data were obtained
at each Reynolds number index at five fixed settings of the variable-
area exhaust nozzle over an engine speed range from military (7460 rpm)
down to the surge region. At a Reynolds number index of 0.2, similar
data were obtained for other settings of the variable stators down to
the fully closed position. Fuel conforming to the specification MIL-F-
5624A grade JP-4 with a lower heating value of 18,700 Btu per pound and
a hydrogen-carbon ratio of 0.171 was used throughout the investigation.
Definitions of symbols, methods of calculation, and a sample calculation
of engine performance from generalized performance data are presented in
appendixes A, B, and C, respectively.


RESULTS AND DISCUSSION

Component Performance

Compressor performance and turbine performance are presented in the
form of performance maps at selected values of Reynolds number index,
and the effects of Reynolds number on performance are summarized. The
effects of variable stator angle and hot inlet-air temperatures on com-
pressor performance are also shown. Combustor performance is presented
in a generalized form as a function of the usual combustor parameters.
Exhaust-system data are also shown to permit the calculation of over-all
engine performance.

Compressor performance. The compressor performance map at a
Reynolds number index of 0.60 with the variable stators in the open posi-
tion is shown in figure 4(a). At design compressor pressure ratio (12.2)
and rated corrected engine speed (7460 rpm), the corrected airflow was
approximately 159 pounds per second, and the compressor efficiency was
0.784. Compressor efficiency reached a maximum of 0.80 to 0.81 at a
corrected engine speed of approximately 6900 rpm. At a given corrected
engine speed, variation in compressor pressure ratio (as limited by
operation of the compressor and turbine as engine components ) caused
variations in compressor efficiency only of the order of 0.01. The cor-
rected airflow was unaffected by compressor-pressure-ratio variations at









TIACA RM E58C12


high corrected engine speeds; but at a corrected speed of 6600 rpm, the
corrected airflow was lowered about 3 percent by increasing pressure
ratio over the range permitted by the variable-area exhaust nozzle.

The open-stator compressor performance map at a Reynolds number
index of 0.20 is shown in figure 4(b). At the design pressure ratio and
rated corrected engine speed, corrected airflow was about 2 percent lower
and compressor efficiency was approximately 0.02 lower than at a Reynolds
number index of 0.60. Peak compressor efficiency was approximately 0.03
lower and occurred at a slightly higher corrected speed. Variations in
pressure ratio at a given corrected speed had a greater effect on cor-
rected airflow at this lower Reynolds number index (0.20). At a cor-
rected speed of 6600 rpm, increasing the compressor pressure ratio over
the permissible range lowered the corrected airflow about 6.5 percent at
a Reynolds number index of 0.2 compared with a 3 percent reduction at a
Reynolds number index of 0.6.

The compressor performance at a Reynolds number index of 0.20 with
the variable stators fully closed is shown in figure 4(c). The variable
stators are scheduled to be closed only for the low-speed portion of the
map. The compressor performance was mapped over the full range of cor-
rected engine speed and exhaust-nozzle area at this stator position
(fully closed) to permit the determination of compressor performance for
other possible schedules of engine speed with stator position and other
exhaust-nozzle area schedules. With the stators closed, the compressor
operates in a region of considerably reduced airflow and pressure ratio
and the peak compressor efficiency is less than 0.60. Corrected airflow
was sensitive to compressor pressure ratio over the entire corrected
engine speed range. The slope of the corrected speed line at 5200 rpm
illustrates the advisability of scheduling a large exhaust-nozzle area
at this operating condition.

Compressor efficiency and corrected airflow for open stator opera-
tion are shown in figure 5 as functions of Reynolds number index for
constant values of corrected engine speed and compressor pressure ratio.
Reynolds number variations had no appreciable effect on compressor effi-
ciency or corrected airflow for values of Reynolds number index greater
than approximately 0.4. At a corrected engine speed of 8000 rpm and a
pressure ratio of 12.50, reducing the Reynolds number index from 0.4 to
0.08 lowered the compressor efficiency by 0.05 and lowered the corrected
airflow approximately 7 percent. The decrease in efficiency and cor-
rected airflow with decreasing Reynolds number index was greater at the
lower corrected engine speeds. Compressor performance losses due to
Reynolds number effects were also greater at higher compressor pressure
ratios, especially in the low corrected-engine-speed region.







NACA RM E58C12


Corrected airflow data obtained over a range of inlet-air tempera-
tures corresponding to flight Mach numbers up to 2.0 did not appear to
generalize within the expected accuracy of the airflow measurements
(fig. 6(a)). The apparent trend of increasing corrected airflow with
increased inlet temperature was still doubtful because the compressor
pressure ratio was not quite constant in spite of the constant exhaust-
nozzle area, and the hot-temperature data were in the low corrected-
engine-speed region where pressure ratio was shown to have an effect (see
fig. 5). However, after the effects of compressor pressure ratio were
eliminated (fig. 6(b)), the trend of increasing corrected airflow still
existed and amounted to approximately a 3 percent increase in corrected
airflow for an increase in inlet temperature from 4160 to 7000 R.
Although the reasons for this effect are not understood, it is believed
that the accuracy of the data involved establishes the existence of this
trend.

The effect of stator position on compressor performance is shown in
figure 7 for a range of exhaust-nozzle areas at a Reynolds number index
of 0.20. Corrected airflow with the variable stators closed was approxi-
mately 38 percent of the open stator value at a corrected engine speed
of 8000 rpm and about 44 percent of the open stator value at a corrected
engine speed of 6700 rpm (fig. 7(a)). The corrected engine speeds at
each stator position that correspond to the control schedule values (fig.
1) are shown on the figure by vertical dashes. The corrected airflow
for off-schedule operation or for altered schedules may be approximated
by linear interpolation.

Compressor pressure ratio (fig. 7(b)) with the variable stators
closed was about 36 percent of the open stator value at a corrected
engine speed of 8000 rpm and about 42 percent of the open stator value
at 6700 rpm. At the corrected speed of 6700 rpm, the closed nozzle pres-
sure ratio of 9.1 could be reduced to 8.3 by opening the exhaust nozzle
or to 3.85 by closing the stator vanes. The low-speed stall-line inter-
cept with open stators and an exhaust-nozzle area of 2.81 square feet
and the approximate scheduled operating line of the variable stators are
indicated on figure 7(b) to illustrate the necessity of an antistall
device on this high-compression-ratio compressor.

The effect of stator position on compressor efficiency is shown in
figure 7(c). Peak compressor efficiency was lowered from 0.77 to 0.75
by closing the variable stators from 00 to 180, but dropped off rapidly
to less than 0.60 when the stators were fully closed to 350. The cor-
rected engine speed at which the engine is scheduled to operate for a
given stator position was in the region of peak compressor efficiency
for the two closed-stator positions investigated.







IACA RM E58C12


Ccmbustor performance. The variation of combustion efficiency with
the combustion parameter wa,2T9 is shown in figure 8. This parameter
is approximately proportional to the combustor-inlet parameter PT/V
and is more convenient in conjunction with over-all engine performance
calculations. Combustion efficiency varied from 0.97 at the highest
values of Wa,2T9 to 0.90 at a combustion parameter of approximately
15,000, which essentially covered the range of engine and flight condi-
tions investigated with the variable stators in the open position. With
the stators closed at a Reynolds number index of 0.20, the combustion
parameter could be reduced further, and the combustion efficiency dropped
off rapidly with decreasing engine speed and increasing exhaust-nozzle
area to about 0.33 at a combustion parameter wa,2T9 of 4500. However,
this condition would not normally be encountered in actual flight and is
presented as isolated combustor performance beyond the usual range of a
combustor operating as an integral part of the engine. The combustor-
inlet conditions of pressure, temperature, and reference velocity at this
point were approximately 750 pounds per square foot absolute, 6100 R,
and 60 to 70 feet per second, respectively. The corresponding inlet con-
ditions for the maximum combustion parameter and combustion efficiency
were 13,626 pounds per square foot absolute, 10780 R, and 82 feet per
second, respectively.

The combustor total-pressure loss ratio as a function of combustor
temperature ratio is shown in figure 9. As the temperature ratio in-
creased from 1.45 to 2.05, the combustor total-pressure loss ratio de-
creased from 0.07 to 0.05. This reduction in total-pressure loss ratio
results from the more rapid decrease in friction pressure loss that ac-
companies the decrease in combustor-inlet Mach number in comparison with
the increasing momentum pressure loss as combustor temperature ratio is
increased.

Turbine performance. The over-all performance of the turbine is
presented in terms of corrected turbine enthalpy drop and turbine gas-
flow parameter for lines of constant corrected turbine speed, pressure
ratio, and efficiency. The turbine performance map determined from
open stator data at a compressor-inlet Reynolds number index of 0.60 is
shown in figure 10(a). The range of engine speeds and exhaust-nozzle
areas investigated caused the turbine-inlet Reynolds number index to
vary from 1.04 to 1.41, but, as will be shown later, Reynolds number has
little or no effect on turbine performance in this range. Over the
narrow range of corrected turbine speed and pressure ratio as limited by
operating in an engine at a constant stator position, the turbine effi-
ciency varied from 0.86 to 0.88. The corrected turbine gas flow, which
can be obtained by factoring out the corrected turbine speed and the
factor 60 from the turbine gas-flow parameter, was about 28.3 pounds per
second.








NACA RM E58C12


At a compressor-inlet Reynolds number index of 0.20 and open stator
position, the Reynolds number index at the turbine inlet varied from
0.29 to 0.43 (fig. 10(b)). The turbine performance map was quite similar
to that at the higher Reynolds number index; the corrected turbine gas
flow was still about 28.3 pounds per second; the turbine efficiency
varied from 0.86 to 0.88. However, the peak turbine efficiency of 0.88
occurred over a smaller range of the performance map than at the higher
Reynolds number index.

Turbine performance at the lowest turbine-inlet Reynolds number
indices investigated (0.13 to 0.19) is shown in figure 10(c). The much
larger range of this turbine performance map resulted from the combina-
tion of closed-stator data at a compressor-inlet Reynolds number index
of 0.2 (lower right portion of map) and open-stator data at a compressor-
inlet Reynolds number index of 0.08 (top left portion of map). The tur-
bine efficiency lines must be considered only approximate because of the
relatively large Reynolds number index variation for this low Reynolds
number index range; but it is evident that the peak efficiency shifted
to a region of lower corrected work and pressure ratio than at the higher
Reynolds number indices. The minimum turbine efficiency encountered
during the investigation was about 0.82 and occurred during closed-stator
operation at a Reynolds number index of 0.20. The trend of the constant
corrected speed lines to lower corrected gas flows at the low turbine
pressure ratios indicates that the turbine nozzles were unchoked when
operating with the variable stators in the closed position. When the
turbine nozzles were choked, the corrected turbine gas flow was still
about 28.3 pounds per second, the same as at the higher turbine Reynolds
number indices.

The effect of Reynolds number on turbine performance is summarized
in figure 11. At the conditions where Reynolds number effect could be
isolated, that is, at constant values of corrected turbine speed and
pressure ratio, the turbine efficiency was not affected by Reynolds num-
ber down to a Reynolds number index of about 0.4, but dropped off about
0.02 as the Reynolds number index was reduced to 0.15. There was no
apparent Reynolds number effect on corrected turbine gas flow over the
range of Reynolds number indices investigated.

Exhaust system. The exhaust-system data are presented to allow
calculation of over-all engine performance from pumping characteristics
which are based on turbine-outlet pressure. Tailpipe total-pressure loss
data are shown in figure 12 as a function of the turbine gas-flow param-
eter wg,5T9/P5, which is a function of the turbine-outlet Mach number.
With the variable stators in the open position, the tailpipe total-
pressure loss increased from 4.5 percent of the turbine-outlet total
pressure at a gas-flow parameter of 1.02 up to about 13 percent at 1.47.
At the static sea-level military condition, the gas-flow parameter is
1.26 and the total-pressure loss ratio is about 0.07. The turbine








NACA RM E58C12


gas-flow parameter could be lowered to 0.65 with the stators fully
closed, at which point the total-pressure loss ratio was about 0.04.
Data obtained with the largest exhaust-nozzle area resulted in total-
pressure loss ratios as high as 0.35 (not shown on figure), indicating
that choked flow existed somewhere near the turbine instead of at the
exhaust-nozzle throat. The exhaust nozzle became unchoked at turbine
gas-flow parameters somewhere between 1.47 and 1.53.

The velocity coefficient of the primary exhaust nozzle with the
secondary nozzle removed is shown in figure 13 as a function of nozzle
pressure ratio. When the exhaust nozzle was unchoked (nozzle pressure
ratio po/P9 > 0.5), data scatter made the values unreliable. When the
exhaust nozzle was choked, however, the data fell about a mean value of
about 0.985.


Engine Performance

Several aspects of the over-all engine performance are discussed in
this section. Typical effects of variable stator position on net thrust
and specific fuel consumption are presented at a specific flight condi-
tion and exhaust-nozzle area. Engine pumping characteristics with open
stators are presented at selected values of Reynolds number index and
the general effect of Reynolds number on the pumping characteristics is
summarized. Net thrust and specific fuel consumption are presented for
a flight Mach number of 0.9 at rated engine conditions over a range of
altitudes above the tropopause to illustrate over-all performance losses
resulting from decreased Reynolds number index.

Some typical effects of variable stator position on over-all engine
performance are shown in figure 14 for one specific flight condition and
exhaust-nozzle area. Net thrust fell off rapidly in approximately linear
fashion when the stators were closed. This is, of course, the expected
trend on the basis of the corresponding airflow and compressor pressure
ratio reductions (figs. 7(a) and (b)). Although the thrust dropped
rapidly, the specific fuel consumption increased only slightly with
closure of the stators to about the mid position (180). However, the
specific fuel consumption increased rapidly as the stators were closed
further. This nonlinear variation of specific fuel consumption with
stator position is a result of the similar variation in compressor effi-
ciency with stator position shown in figure 7(c). Inasmuch as the vari-
able stators are generally scheduled to be open except for certain engine
transient operations at reduced speeds, the following presentation of
generalized steady-state engine performance is confined to the open
stator position.








NACA RM E58C12


Pumping characteristic maps, which consist of the variation of en-
gine pressure ratio with corrected engine speed with lines of constant
engine temperature ratio and corrected airflow, are shown in figure 15
for Reynolds number indices of 0.6, 0.2, 0.12, and 0.08. The peaks of
the lines of constant engine temperature ratio show the regions of maxi-
mum combined compressor and turbine efficiencies. The slope of the
lines of constant corrected airflow at low corrected engine speeds re-
flects the reduction in corrected airflow with increasing pressure ratio
as discussed in the compressor performance section. Over-all engine
performance may be determined for choked exhaust-nozzle operation at any
flight condition corresponding to a Reynolds number index greater than
0.08 by use of the pumping characteristic maps and several auxiliary
curves (figs. 6, 8, 12, 13, 18, and 19). A sample calculation of engine
performance using this method is presented in appendix C.

The general trend of engine pressure ratio and corrected airflow
with Reynolds number index is shown in figure 16 for several corrected
engine-speed and temperature-ratio conditions. Curves similar to these
can be constructed from the pumping maps for calculating engine per-
formance at other engine conditions and can be interpolated for inter-
mediate values of Reynolds number index.

The reduction in net thrust and increase in specific fuel consump-
tion resulting from Reynolds number effects on the engine components are
shown in figure 17 for the rated engine speed and limiting temperature
condition over a range of altitudes from about 35,400 feet tropopausee)
to 72,000 feet at a flight Mach number of 0.90. This corresponds to a
range of compressor-inlet Reynolds number indices from 0.46 to 0.08.
Increasing the altitude over this range reduced the corrected net thrust
by 14 percent, 6 percent of which was due to reduced airflow. The lower
engine pressure ratio resulting from reductions in the compressor and
turbine efficiencies accounted for about 6 percent of the thrust loss,
and about 2 percent was due to the increased tailpipe pressure loss
brought about by the above effects on the turbine-outlet Mach number.
The specific fuel consumption was increased about 16 percent as altitude
was increased over this range, 6 percent of which can be charged to com-
bustion efficiency, 8 percent to the compressor and turbine efficiencies,
and 2 percent to the higher tailpipe pressure loss.


SUMMARY OF RESULTS

The results of performance tests on the XJ79-GE-1 turbojet engine
and its components are summarized as follows:

1. At rated corrected engine speed (7460 rpm) and design compressor
pressure ratio, the corrected airflow was 159 pounds per second and the
compressor efficiency was 0.784 at a Reynolds number index of 0.6. At








NACA RM E58C12


this Reynolds number index, peak compressor efficiency was between 0.80
and 0.81 and occurred at a corrected engine speed of approximately 6900
rpm. Compressor performance was not appreciably affected by Reynolds
number at Reynolds number index values greater than approximately 0.4.
Lowering the Reynolds number index from 0.4 to 0.08 reduced the com-
pressor efficiency 0.05 and lowered the corrected airflow approximately
7 percent at the highest corrected speed (8000 rpm) at which comparisons
could be made. Increasing the engine inlet temperature from 4160 to
7000 R at a constant Reynolds number index resulted in approximately a
3 percent higher corrected airflow for a given corrected engine speed
and compressor pressure ratio.

2. Varying the variable stators from the open to closed position
(00 to 350), resulted in reductions in corrected airflow and compressor
pressure ratio on the order of 60 percent. Peak compressor efficiency
was lowered only 0.02 by closing the stators halfway, but decreased
rapidly when the stators were closed further. Net thrust fell off
rapidly when the stators were closed, but the specific fuel consumption
remained relatively low until the stators were closed more than halfway.
The planned schedule of variable stator position as a function of cor-
rected engine speed apparently safely bypassed the low-speed stall region
and passed through the regions of peak compressor efficiency.

3. Combustion efficiency varied from 0.97 to 0.90 over the range of
engine and flight conditions investigated with the variable stators in
the open position. The combustor total-pressure loss ratio varied from
0.05 to 0.07 over the range of combustor temperature ratios investigated.

4. Turbine efficiency varied only from 0.88 to 0.86 for open-stator
operation throughout the investigation. The minimum turbine efficiency
encountered was about 0.82 and occurred during closed-stator operation
at a Reynolds number index of 0.20. Turbine efficiency was not affected
by Reynolds number down to a turbine-inlet Reynolds number index of about
0.4, but dropped off about 0.02 as the turbine-inlet Reynolds number
index was reduced to about 0.15. There was no apparent Reynolds number
effect on corrected turbine gas flow over the range of Reynolds number
indices investigated.

5. An increase in altitude from the tropopause to 72,000 feet at a
flight Mach number of 0.9 (Reynolds number index reduction from 0.46 to
0.08) resulted in a 14 percent reduction in net thrust and an increase
in specific fuel consumption of 16 percent in comparison with the values
that would be obtained assuming no losses in component performance with
increasing altitude.


Lewis Flight Propulsion Laboratory.
National Advisory Committee for Aeronautics
Cleveland, Ohio, March 20, 1958








NACA RM E58C12


APPENDIX A


SYMBOLS

The following symbols are used in this report:

A area, sq ft

B balance force from thrust capsule, lb

C coefficient

F thrust, Ib

g acceleration.due to gravity, 32.17 ft/sec2

H enthalpy, Btu/lb

M Mach number

N engine speed, rpm

P total pressure, Ib/sq ft abs

p static pressure, lb/sq ft abs

R gas constant, ft-lb/(lb)(OR)

T total temperature, OR

t static temperature, OR

V velocity, ft/sec

w flow rate, Ib/sec or Ib/hr

_I-
( -1

Sy correction factor, -- ( --
T 1.4
1.4+ 1.4 1
\ 2

r ratio of specific heats

8 ratio of total pressure to NACA standard sea-level static
pressure








NACA RM E58C12


6/f-6 Reynolds number index

i efficiency

e ratio of total temperature to NACA standard sea-level static
temperature

(p ratio of absolute viscosity to viscosity of NACA standard
atmosphere at sea level

Subscripts:

a air

B combustor

b bleed

C compressor

cr critical

eff effective

f fuel

g gas

id ideal

j jet

n net

s slip joint in inlet duct

T turbine

V velocity

0 free-stream conditions

1 inlet Venturi throat

2 compressor inlet

3 compressor outlet, combustor inlet








NACA RM E58C12 15


4 combustor outlet, turbine inlet

5 turbine outlet

5a turbine outlet (GE control thermocouples)

9 exhaust-nozzle inlet








NACA RM E58C12


APPENDIX B


METHODS OF CALCULATION

Airflow. Airflow was determined from measurements of total-
pressure upstream of the bellmouth, static pressure in the inlet Venturi
throat, and temperature at the compressor inlet. These measurements
were used to calculate engine-inlet airflow from the equation,


T-1 1

wa = pA 2 )RT-


Overboard leakage airflow was calculated similarly from pressure
and temperature measurements in the compressor discharge standpipes and
was subtracted from the inlet airflow for all stations downstream of the
point of extraction. Tailpipe gas flow was obtained from the expression

g,5 al Wa,b + f/3600

Compressor efficiency. The compressor efficiency is defined as the
ratio of isentropic enthalpy rise to the actual enthalpy rise across the
compressor

(Ha,3)isentropic Ha,2
C Ha,3 Ha,2

The enthalpy values were determined from charts based on the material
of reference 1 using variable specific heats.

Combustion efficiency. The combustion efficiency is defined as the
ratio of the ideal fuel-air ratio necessary to obtain the engine tempera-
ture rise to the actual fuel-air ratio:

(wf/a 5)id
B u("f/Wa, actual

The ideal fuel-air ratio was determined from the fuel properties and the
engine temperature rise (see fig. 18 or ref. 2).








NACA RM E58C12


Turbine efficiency. The turbine efficiency is defined as the ratio
of actual enthalpy drop to isentropic enthalpy drop across the turbine:

Hg,4 Hg,5
IT 7H H
(Hg,4 Hg,5)isentropic

The turbine-inlet temperature T4 was calculated by assuming that the
turbine enthalpy drop equaled the compressor enthalpy rise. The enthalpy
values were then determined from charts based on the material of refer-
ence 1 using variable specific heats.

Jet thrust (measured). Jet thrust was determined from the thrust-
measuring system by an algebraic summation of the forces acting on the
engine:

Fj = B + As(P PO)

where B is the balance force from the hydraulic capsule. The last
term represents the momentum and pressure forces on the installation at
the labyrinth seal.

Jet thrust (calculated). Jet thrust was also calculated from the
gas flow and effective jet velocity:

Fj = g Veff


The effective velocity, which includes the effect of excess pressure not
converted to velocity for supercritical pressure ratios, was obtained
from the effective velocity parameter of reference 3 (also see fig. 19).
The ratio of measured thrust to calculated thrust is the velocity coef-
ficient CV, which can be used for all choked nozzle conditions to obtain
true jet thrust when multiplied by the calculated jet thrust.

Net thrust. Net thrust was determined by subtracting the inlet
momentum from the jet thrust:

W"a,
Fn = Fj 2 V0
Sg








18 NACA RM E58C12


APPENDIX C


SAMPLE CALCULATION OF ENGINE PERFORMANCE

FROM GENERALIZED PERFORMANCE DATA

In order to illustrate the method for obtaining over-all engine
performance from generalized performance data, a numerical example is
presented for the following flight and engine conditions:

Altitude, ft . . . 35,400
Flight Mach number, MN . . 0.9
Engine speed, N, rpm . . 7460
Exhaust-gas total temperature,, T, R . .. 1530

From these conditions the following quantities are known:

PO = 490 Ib/sq ft abs

tO = 392.40 R

From these quantities, and assuming 100 percent ram-pressure re-
covery and an NACA standard day, the following parameters may be
calculated:

V0 = 874 ft/sec

P2 = 829 Ib/sq ft abs

T2 = 4560 R

V/ 2 = 0.9373

62 = 0.3918

s82/02- 2 = 0.46

N/V = 7959 rpm

T9/T2 = 3.355

From figure 15, values of engine pressure ratio and corrected air-
flow can be obtained at a corrected engine speed of 7959 rpm and an
engine temperature ratio of 3.355 for various values of Reynolds number
index. Curves similar to those in figure 16 can be constructed and the
engine pressure ratio and corrected airflow at a Reynolds number index








NACA RM E58C12


of 0.46 can be obtained:

Ps/P2 = 2.633

wa,2z 27 = 165.75 Ib/sec

and

P5 = 2183 Ib/sq ft abs

wa,2 = 69.29 Ib/sec

The overboard bleed flow is about 1.5 percent of the inlet airflow, and
the airflow downstream of the turbine is

Wa,5 = 68.25 Ib/sec

To determine combustion efficiency, the combustion parameter wa,2T9
is calculated,

wa,2T9 = 106.0x103

and from figure 8,

71B = 0.968

From the engine temperature rise, the ideal fuel-air ratio may be de-
termined from figure 18

T9 T2 = 10740 R

(wf/3600 wa,5)id = 0.0148

Dividing by combustion efficiency to obtain actual fuel-air ratio yields

wf/3600 a, 5 = 0.01529

and

Wf = 1.044 Ib/sec or 3757 Ib/hr

To obtain the exhaust-nozzle total pressure, it is necessary to deter-
mine the tailpipe pressure loss, which is shown in figure 12 as a







NACA RM E58C12


function of Wg,5-JTS/P5,

Wg,5 = Wa,5 +wf/3600

68.25 + 1.04

= 69.29 Ib/sec

Wg,5'9/Ps -= 1.242

(P5 P9)/5 = 0.0665

and

Pg = (1 0.0665)P5 = 2038 Ib/sq ft abs

The exhaust-nozzle pressure ratio is

p/P9 = 0.2404

From figure 13, the exhaust-nozzle velocity coefficient is

Cy = 0.985

To calculate thrust, the effective velocity must be determined.
From the fuel-air ratio and exhaust-gas temperature, the ratio of spe-
cific heats is

T9 = 1.337

From figure 19, the effective velocity parameter is

Veff/Vig9 = 1.513

The effective velocity then becomes

Veff = 2452 ft/sec

and the jet thrust is

F = W CVeff = 5201 lb
j 9g







NACA RM E58C12


By subtracting the inlet momentum, the net thrust becomes

Fn = Fj- 2 V = 3319 lb


and the specific fuel consumption is

wf/Fn = 1.132 Ib/(hr)(lb thrust)


REFERENCES

1. English, Robert E., and Wachtl, William W.: Charts of Thermodynamic
Properties of Air and Combustion Products from 3000 to 35000 R.
NACA TN 2071, 1950.

2. Turner, L. Richard, and Bogart, Donald: Constant-Pressure Combustion
Charts Including Effects of Diluent Action. NACA Rep. 937, 1949.
(Supersedes NACA TN's 1086 and 1655.)

3. Turner, L. Richard, Addie, Albert N., and Zimmerman, Richard H.:
Charts for the Analysis of One-Dimensional Steady Compressible Flow.
NACA TN 1419, 1948.











NACA RM E58C12


.7 77. *- 7'. *7. '*7 i x; 7


'= ..-', ".


.7 .7-




7' -. 7..'











I,
-'"'' I.. :_ :_




























v.7.
I. .. i












i,_- .-


















,1.j .. -' I, _,_. _-_, -_ ,. "- : : .-.'-_ ,



































.- 7- [- z'- :- ,- '7-,.. .. .. -'7.'. .... .
.-- : j -' I *.... _,0 ,-.7..7 -U .: 7 7, C .' __: 7. ..r.0-, .
rit ..... ..... .. ..... ....
_... ,_ Ii








[ .:- --. .. .




-. -_ ... .-. '






,--













iil


I '





I ,: -




1.2~


i-'

-- .-


"'." z. J ,


,z x r.


,''"":


..-~. i.


..;..


ir
-yri


o .' ,, >, ,, :
I ." ~ I IJ-. .


I;.i. --ii; r~-, --.r

--r'''? ZT~_r i


,..," = .* : *. .- : ..'-0 ~3 ,













1 1.-' :-'='=" g -.' '; ..' 1 --' 'j". >









,.- .--, 7 2 .
r '' t '





c zi
C 7 1.


z
r:- r: ,.--, .s : jia..!:


.. -- ~ C '- S ,1
I

i3 i't


''



il-~
srr
.1...- r -


~ r r. -I











NACA RM E58C12


: --.. .-..- ..- .-.. .. ... .-. .. :-.- ~. .
d" 2'r i'~lc i~ ______NO nn nn n

1.770077. ;. *.i -.m .
S oin>[








,' -1 n p .-,,.,' ~l '.,' .- i'. ; -, .-., .- .. ., i




*a ca ol cae ^o e a o ao l e aoo le o a e ao lo e a e
*a" "4 h





0 I-i Hi ii
I 30071,73 33














.0->- "a07ta. 777"1 .00)0-.0.0w O~l 07077.0.0 .077 .0.0.0 .0.0 -.0 ...0 ..00.. 0.0.00.0 0.70. 0.i
--0707 0377770 770 307.* 0n. .







770u 0. t7 0 f-77*' 70.07.0-70 .00770.i0
7,4 00 .0 3 C
07734.00 07.4 _0770,..7 0 7 7 0 7 0.0 7 0 03707-070_ 7 3 3 0 _70700371_
07747. .003 ~ .0.7.77' .0707700 0.77.077 07 ~ ~ .0..0300*0 007*0' 0377 7
4.000 ~ NnwN~w~ ~~~i~n n~~ o


: -: i .....






_ -.70 -3 -
0 ~0 00
J-ajt oo o
li C
^ ___ 0 __


70700 07100.-* 3a .*- 7 00. 77.00


. .. .. 7


::=_ .n mm ,: -_.. : -. : -- :. :-.-.:.- .:.- i


7. .-.7 NN~o ... 0 0N 7 Nr-7 -7703


-o. o.. ~~~

Z'r-~? 1~~.` ~603 (lm1NI LI) O I NII


-_" ':At ':-I .',':' ,_-,I: .',, ":,::,:j,: ::5 :: I--~. ': ':




,- .. .-,7 .7 7 7 .... -. -, 0 -.. 7.7. ., .
.. 7... .- 7.77


' -.


k1

I 37


-


j-.- ool~mrl -.~.r ~~ir-
i


000ooooo


-I


"--r` ~~1z I













NACA RM E58C12


-n 4 *ir -

.43 ElB -1 ^ ^_




















4- 0 .D3 4n 0 -. -. j 4. -
I,'
i-T ) >S












Jt 3 *






I. r

^^r -^:*-










- .,



-'' I ;j




0 JD




s

| ^ 1' _-.



p~- jf -*


'a


C- .4.3r4 4434
rI ; .34.3.-i.3
4' .3



C I -





H






E-e -


- .-.- -, 4,-


.B.43 [--?^ _- ---- .4 ~ -- -




.. .... .. .....













-a -
^ L -* I i*| ;F S:.13 ..'a- u-1,- ui4:rr P; S 1 i I Ml;


...34 4 .4 ~4 ~ 43.4444.3 r
.3 ili~- 4.3,,'-'. 44444.3
r- I 43 .4 434: 4 444..3 t.,, 4'.3


--I


jl--j r,~ri-C ~~j;... 19r-I-r rII'IC Ili-~~
-n~~1 >*?1 1 --lj.-.ll li~l- ii rl--0-
I~-- I- -r -1
rr -



I:C~i~ Cir- -i--i _'C1--- 3--i-'I 'n-- r~'-il-
I- 1-.-. ~
-~r ~l'l-?T.. -I1~~ IlTIll I--il~ ~-1--


;11 I
I --I
r~~~.i




~c;-
Ir~ly


',f, .- ,:., .r:,
I .- -.,



.-,.,11


7


F I F F rT


7 : i A -A I ,I


I


~:; ir

- .-II I


;i, r-
.r.


el 7 -
:7 7






C 7. 1. 111, m; m;- ,-r j.*r


, D z x I


-;o-;m r*~nCi ;~r~, -j7F~ -i-
I--I~ 1~ ~


ri -l~-i
.i
























a c
Iti




,Z I f : -
*ia'c *


,--.. .... _.. ..
- ,,0 _. ''
o"


,. .....






.,,



6 74 074,1 0044 ,07
.7 __ (0 07-000 007-00 7-007-7-



2? 0r (
: l t-. Qt lO~r llit-f '' i[0' l?
. L < Q if''^ *^tf .t '''-


r' q -1r2l. i .. ... .. 0-


1 ;.,_: solo 1 :0 M T!:'












t-l t-. t i,.' -, T q .- 0, ='- '2 '1 0 'Cl ^r 30 --
f~Ko 444 44 o O 3 go io i









i~r0 M I oil q 11"na on~


a I





O I' .
II







^ '* "" ___ ^t ^itnm ^ ^tN^^^^









0o
--^ m-m^ o M
o : 3 r ---* r (m 'iu"i


s ,1-----







s IZL& #,-
I -






I ~ 1 -i+ .. .. l


- ;t., 1

] ....

l, I..*:4 ,

) .' ''

nt


"i






a.


o(!O0o !0 :l !0 Q
>noOW'^" mV)OHMO


7-(000 0(0(,17 04(0704

.n' .* :725. 1NNN;



....'LI li~ ... ".. ,2;


00 'S40 0 "*^ ,' 1* -'- 0-:


p-ay 0-' hows 310 H. C r ^.r-


- --'-- .-7 0 -7 .' --



,-j '"7 ". .. r.. ..,N 7 ...

.772.7.7 04 Ca.77 .7r- '7 .174.*7
-I 70 700044 ..-., 00.( 000441 000044 00000


itsr m.......


NACA RM E58C12


inau r*ocmolif o< ***i


Y-, A ,, ,
o '";.3.


tmiu>o r <


'




i





ddddd

crun~~a
--~I
~


t9~~~
rn~ton

~ntnuruiui


)I I--
'2 1


''

murvlulm

mnmran
:rI



~ulmaa

r-


i r-i
-. .~ 5.
mto~~vt r~~s~u,
~ -I-.-
----~


amrme I i:
in~o~nxvr
m~o~mm



rr:, m~o~o~p
5 """~0'0
~ ooooo


-. .





AI


nnrol~an
oorl~no
mmmcm



a~mmn
molm~~

~nmm~m


As? pil


--"


I" ,,,,,,,mo i~~ c~
0.2012 ~ m~~m~ r~r cr















7. 0 ? Id -0 a.- Ca,- .7.7.c J) a u iJ
r0 u'j f 7 '- Ti 0n .. m zi'


I P


-nrini
I-

ii 3j rSui~CC


Ii~-O ~P-
oo-~ r:r;


I -' -

_. .1 _.



1
L. ^ i' l^ c. [ i- -.




^ T; C. !






fl-
f,; ., F,















i ^ i ]:' T,* ^ i-j ^
3S


r I


,_, t---











'` ,.-- : :
......


- r -77 ,~- 7.777 7.7. *7.--.7. ;. .~


- ; ~ 7.,~- 7"7.77 -


...... .. .


















li.'-, 'L "'.. ...: 'h '9 ... ,.,' i: l Z ?'-
,d,:, -. 7 ... .... ,: ...

z ,i 7., z. L '' -- -


*
10
0U


,H S -..^'
C


o





* I, .",



Ii










I-:,
] ^ '= ______
-I .1 ^I
|~ _a3

,L T
It :: -'.
jj?" .*
IsT ?


. ...... .- n ;






i-



-c-


F-i


NACA RM E58C12


i..l r~r~r~o


L- a-- ^ 7 1- *_.t.


''""' Y~~"


-;, i..I.
;.,....-





-'I-




11.:


t


Dir*j.~

*i~li Ir
r~'


~II~



- r- .;
....


.nn-n
';~-~


n -;c;


~si. -r.
--~-


S-- -r- ,


. z ,. ,.


w. i .. ,_- -i -i ,_- j. i,-.- ,'-,- ..... ....... ooo o


-7! ^,;'! ? ^p -^ e r ^^i t' 7, T''
-ro 5





iB0 0 *"T *~i TiiT7. 2 A- -i. .1 Z. 71 a 0 t_. J -


......5"' i-"" ....... : ; "5 -*-"" ..... o




































E-




0









0)






























0
OC,
I-










E-(







* <

@,










1-


NACA RM E58C12


.., "0 0 -0 0-0.0 0000 0 .- .. -.' 00 0 z, :: '
.-mm m "; mm 1 ^f -S00 1 -







S. ... .. .. .



a9 .. .* ., .. 0b
-.0 -- 0, .-.... .: 0. "0 ,
-. 1 1 0 S0000 00o o


.: -. ..... .. ... .. ... .. .






"-"' 0 0... ... 0 .- '

t-





















c.000. I I a N 0 0 0 0 N I
-'-- H d .- ^d .- 0 ..
S. 0 0 0 00
















0 IO





," ,1'.-00 .00 0 -' 0. ',': 00. 000'!,:, -.00:e t f. -n' om 00 0> 01 0 -a-'010 0 0
t: i .--



S.. ... ... .-,oo00 0,







.I t .., ... ..- .... oo 00...

,r,: t ,-", ,-'.- 00 011 -' .- '' l.,. 0








K.... -t l' '13 '0 1 -
,1" c 7 '^'"l*'* ^ '"'' ''1 c i' r n ii -" t s n i o w 3 "iooYt f )if s .*

3n~ L.mm ~ lm


















;v


1,



,, it!
-.i.!..


- .- ; 4i 4 .In .4. a 1 .. ., ,






. ...... .3.3.4. .... O '


. 4
4 :,.. _.. ... .. 7. 3-, 4'. .. .. 2 .-- -.. --- .f .. -,'.-








-. ...- .. ... ... .. -. ..












J,- --,


























n -, -- :.-- "._
-





^ :c 3t~ a *u~ T Srui* *<*** ico u-n tlw i0 J' -''^ ~

FO


-C rl-rl~r rTTTT li:j=~~ Lin-- -L~ln'-'r N~i~ll Si--r~
ii~l'l'~ 'I~i* li~iji ~-C ilI ~5jTc~ ~3-jrl-
ilrlliril'r Ii'i'uii i~lllill il;llilil iliililr ilil*~~ C~CCC ~iiici

i-Sr -cui -'.ll'i L-lrlil -i^~j I-~ii ~j+~- ~.~CIli
r


NACA RM E58C12


4 *






L-- .....




lc l
SI- 1>

?-J i |
t2^ '' ''


n _u,
a L:'


3CC^,C no n, s ToV 210. 0..7


- 1'*wul iljj -~D a 01* r j-i ul L lr- 6iTc -T ~Cj Yf ul I.'r I fr, -C r fi '**u ar-ac"


1














NACA RM E58C12


-0- ...'., 0, ..... .. .. ... .0...-. ..0 0 .lJ..IO '0,(0 0r ..000,140 0,1-0,0r0
04.01 p. 1 0 Z.7







S 0 :- 1 0.10.00 010t- r- -- ,- : -: 1: .0

0 . . . ... .







0. 0 ,, P. -1I 1. MI 0.0, .. .. I u 000 0-0 0 0o...







~- ~N 9 00100 2. oooo0 ooo.o-
3- .. ai I M O .. 0 =i N4. ... -.....





'00 ..1 .-I



44. 0 4 a, 0 z 0 1.0 o,0-,oo .. .. 1-0,, -1oi0'0,: 1-1 0, _,o 00 (0

0 INN 0,10.10, -N ...'1 l 0N 0W.1-00 N1-..
4 ,! P :. .* .I" 1* .10 0





c 40,4 ----------- .. ... .. *4 ... ....01-0, I..... 0,0,0,0,0,--------
14 434 0, r 0 1,




Or I

3 44 0 r1; SG 0 ,, T 4 .-i wo ..-1,0,0,0 0- 0 0,01-0,0





01000 -- 0 -00
0 o41 01 M 4 -- *. 1! 0, 0 1!,


N P .. 0 .0 ., .. .- .. .0 ,, 01. I* 0=1-0 T ,



10,0000 4 010100... ....... ...... I. 0,0,0,0,0,
04- 00 10 1-.. 0 I : 00000 00000








0 P -, p p.0,


ps. -Ion










T -0-

.. p N^.p p.0,0,10 0,p:r p. I 0 = = !;, ,0*p.01 ,


* .: .> ..... 0, N 'm'o 0, 0,o -

010,010,0,0,0,01 NNNNN ww4,.N:, NOr m NII 1





01 .-. p. 0. 0 ., '. 01- 1 1 I I I-
-.10,0, N' 0 0 NN. 0,0.0,011o 01 I I 1 1. o I 14140,14


I

-- %











NACA RM E58C12


43 r-4 )04t000 (010 441id 0 *LJ- .; .S4.. .444 r i*- -
"E 3 a ,o *1 r-10104Bf* 144010444 to 44~ 10i i *iw -1I. r j- .-


U4040 3.. 44



I0 N 44 OL R- 101 1 a0I4 1 ..
Ai A 04r- 0- 0 0- 4,iO Or.0 .44, -,_2
fo 444(444 1040.10004 04444104040 404.44.41004 44 44.. __

*iSnsin o~f rjmpcom ~WOjtDoo(ii-~u '-r" ''''


.444 b.4 44.0
1 -jr '4
-N- ~
40 0.




434 4 4

I-


:44 i
H-s


I '


017h- -1g. ;.. '-.p q 0 0 '. -|


K 1. ai





it; it -














.. .,, 4 .
0. 4 i i- A







i,
^-" -^^ ^ ^ ^ :^


4 ) ^_ _

0 .44=444
aif

M ,,
rl j ) fl


H

i 1 .4 riir


J -r------
^ 4 4.

^.p


I, A 4, ,- 4 7 o z -, 4 -, 4 ... ,- ,






.4,044,0 40,0-44,0,0 01=4401,0 44 4- 4 '




1710N VAN" ,,e TOM

,, :;i: : : ii b= i "!i


iLi:'1'
S-- --'- _-4', 4.-.:..
--4 4 r1--. 44..4 _______ *
:. :-



,=j


tO ;, -/ .....y *r --! =lo K ; :I
0dN _______ _____I~N


a













NACA RM E58C12


ON O o 0
(- iiror ot.iol


,.4
44 -


cNNm(n iN~j~im
nrlNLO~I nF-nOn
I~~
On~mF ON~~n mR~Nm
~~rnn ~t~al iar6d


S. : -




- : :': Z

' : 'f,_,;


toto.NO tottotN totNtit



NNNOmt N4ONNN Oto0tN


... .. ... .. ....










I
-


m. ...... ......




oN NNo N tto. NNNO NNwN
Nt~ONO (OtN C ..t.tOO tO~tttoot


i, I~;s
..
r







L~::L



; i;
F-
L

I


t~l-.a1-


ii.



::
a
r-,ii ..
6 j:
c


~iF-mono
r-crfm r~i
~r(r(T(i

1


mF-cuin~
i~nrrunr



mr-inrln
~no~m~-r
1-.


mm-*or-pg
i-~~nc~
Ninaioila
arr-*aa~


Dmnmnnl
m~rrlrlmr



r3nolo~m~o
mrm~aom
17~1., f

























a) r. .Z
+- 1- -,









rd _____ a) 4-:
a) H 0
U- *H4r-1 _


d a 3 -P -
*r- H -P .-- / -

( 0



__ W_ __ ___ ^ _



_____ ---- ----- ----- ---- -- l -. --- ---- ---- ---- ---- ---- ----- ----



-- 7---/----


/---------





/EEEEEE


0 Un


2ap 'uoT'Tso uado .CTTnj moiJ uo-Ilsod .)qmaiusae ojeus aTqer'EA


NACA RM E58C12


1--
;o 3
rl

C.
ci
E'i
0 ,-i
0

-,-I
0
0 P4


o a.
Y'~ -H




0









O 00
O rO









o OI
00





4-1)
o o













0 I -tS
0 P
T, 1P -












+r-O
O ** +M














0 1 0 9

0 C dI
O 2 4-





a.) ca
c -i 41
O 0C
0 O -




0 +> +-
Clj 34-

















0 0 U)
O 8O
0 C -. 0)

oC,
-ii


oaa








5 .
0 10

CCI
0 0 + L
0j CO 0






IACA RM E58C12










a






.







0
o
,-
~b II ,
f~'~~I Yti~f~



,', ,. __..











34 NACA RM E58C12











I





00




1 4





Sr I
i c, lx, u





a 0











00

1_-' ,n 0 0
A--- .- 0



o N
t)


oa : 0 1U1




0 V, a1
p. I LI l.I







0 I -
u^ C I I 1CI I n _o
6 i iI I In -








a Cc
-*



6n l









0. L1 a. > 0 0 -
3 I 1 1 ND U \













4-' I 3n
3 ra rID


,--I

0
- -4
r-








NACA RM E58C12


Exhaust- Compressor
nozzle area, efficiency,
A, IC 0.70
14 sq ft percent -.75
O 2.13
-- 2.35 _
0 2.61 .78
13 A 2.94
13-- 4.70 ---- 8300-

8000
12 ---- .8) V
7700

/1 460

S/ Corrected engine
7200-- speed,

10 700 r -70-
Irpm


.0 1 0 6800



,3 6600
11: 128 136 144 152 160 168 11
Corrected airflow, w., 2-,/0/52, lb/sec

(a) Compressor-inlet Reynolds number index, 0.60; variable
stators open.


Figure 4. Compressor performance maps.








NACA RM E58C12


4 I I I I II I I
Exhaust Compressor
nozzle area, efficiency, 0.72
A, TC;
1 sq ft percent
0 2.15
o 2.35 .76-
.63 08000
A2.94

.77 7700






10 .6_ Corrected engine
speed,
1--''"- O -- --oool- O -





'660






Corrected airflow, Vwa ,,/ / 2' lb/sec

(b) Compressor-inlet Reynolds number index, 0.20; variable stators open.

Figure 4. Continued. Compressor performance maps.
Figure 4. Continued. Compressor performance maps.








NACA RM E58C12


38 42 46 50 54 5E
Corrected airflow, wa,2-2/O 2, lb/sec


62 66


(c) Compressor-inlet Reynolds number index, 0.20; variable stators closed (350).


Figure 4. Concluded. Compressor performance maps.


I I I I t I
Exhaust Compressor
nozzle area, efficiency,
A, c 'C 0.50 P
sq ft percent 050

0 2.17
5 0 2.34
5 < 2.64
1A 2.93 .55
S4.89 M
n 746800 60


6800


Corrected engine
speed,
N/rpm,
rpm


-5200-


.,,.,.,,


a-Nsooc


)









NACA RM E58C12


1. O0


.92 -- I --




. 8 8 -- -- -- -- -- -- --
0 .1 .2 .3 .4 .5 .

Compressor-inlet Reynolds number index, B2/9p2-2


Figure 5. Reynolds number effects on compressor perform-
ance. Variable stators open.


I I 1 I I I I
Corrected engine Compressor pres-
speed, sure ratio, -
/P3/P:
rpm

0 8000 12.50
O 7460 11.25
0 7200 11.25
A 6600 8.50
A 6600 8.00





---

1!


0 X

U-rd
d P




H r-d
q-4 H
, 0

Ok
SRa

ca
o 0


4- 0
as co
0 -- 0
* t C H
.d 0



4-4 O



+o ~r.
0 .d
CH





o r o
u N
U o


f-i


-


I


Jvow'























a
0




I1
0,


cc



0




o
a"


























.,.,
o







cq
F4 (- 5-


0
C, .
< ur'
sr&
0;I

0 ~


NACA RM E58C12


150




140-




130



1 Engine-inlet total
120 temperature,
T21
OR

10 525
O 0 600
700



100
6000 6200 6400 6600 6800 7000 7200
Corrected engine speed, N/-/02, rpm

(a) Constant exhaust-nozzle area, 3.16 sq ft
I I I I I
Corrected engine Compressor
pee- pressure
ratio,
rpm P3/P2

0 6600 8.56
1.04 O 7000 10.35
1 7460 11.95


1.00




.96
400 440 480 520 560 600 640 680 7
Compressor-inlet total temperature, T2, OR

(b) Constant corrected engine speed and compressor pressure ratio.

Figure 6. Effect of inlet temperature on corrected airflow. Reynolds
number index, 0.4; variable stators open.


























- 0
0


RACA RM E58C12


140+-


I I I I u t e r I
I Exhauxt-nozzle rea, I I I


A,
zq1 ft

2.17



4.o.9

Variable-st3tor position,
deg
0
35


40 ___







S--- -- ---


4800


6000 6400 6800 7200
Corrected engine speed, 1/- / rpm

(a) Corrected airflow.


8000


8400


Figure 7. Effect of variable-stator position on compressor performance at Reynolds
number index of 0.20.


d


1 *14

























- O

0


12 H-


Exhaust-nozzle area, I I i i I I I


A,
sq ft
2.17
2.34
2.63
2.93
4.88
Variable-stator position,
deg
0
18
35

Variable-stator scheduled
operating line


*" -- -- -- -- -__7- -- "4 --- -


8
Stall '
2-----po"-O






4
6 -- -- -- -/2- -- -- -- -- -


5000 5400 5800


6200 6600 7000 7400
Corrected engine speed, N/S, rpm

(b) Compressor pressure ratio.


7800


8200


8600


Figure 7. Continued. Effect of variable stator position on compressor performance
at Reynolds number index of 0.20.


NACA RM E58C12


_ g: ^ _


_I_










42 NACA RM E58C12






















.7 -___- ___
E..
.7(. __ ;q ft -O




.6( A 4.88


S.6 d --4.

4 .,
g _Varbc-_ tnorp



Ci
S 4- ----- .i,
0
it-] -- -- -- -- -- -- -
o __ :il __ __ __ _


(c I'..mrpre=or efficiency.


Fiurfe '. C:n.iudeid. Effect of variable .tator position on compressor performance
't Feyr-,olds rnurjber index o:'f 0. .












NACA RM E58C12 43





-x

0
C']




0





0
w
ri







q-ar -- 0--0-- G- -----









L. O .p C OW0 O
rl)






-O -o -- -- -
)(I


L oooa 0
0a i s d .





I H r Cl2 H


E m'



a ot




S-- -- -- >-f

IL Q) O
--- j j --- ---- ---- ---- ---- ---- --- ---- ---- ---- ---- ---- -- 0 $ -1

| ___ ____ ___ ___ ___ ___ ___ ___ ___ ___ ____ ___ S 'H


%t '.3uaTo'T.jja uroT rnqinoo0j










HACA RM E58C12


0 Ti







S0





X.
o m



P, 4-)
0







L ) 0
0 0
4-)






0 ,
4---
3 0
0




















0
ft -




u a i
1^ (L


2,/( 2d) 101. J
ssoT -nssaad--Bqoq. jopsnqmoo









IACA RM E58C12 45




o
o
0


\
)0





8 \ \
o --o \

\ \ o o
oa o \





p o | 6
C _o cm
E-) c H *

1 ) -- 0 -

0I -

-1T 0 V

_O *H tO 0
a r T o o a

_O____ __ 0 0 S
to


E- 44 _

0 H .

0
to -
Ho 0

4' 0


0
D H
3 El


qT/n,4a 1'1Ie/tHHv 'doxp AdyqO wa aon;q npl. pavai.-oo











46 NACA RM E58C12





0
0
UY)




\to
0


H u

\ o r)



S0 \ H *-" 0 p
o C; C- .j u



__ __ 0
+ d 0 GJ
oo o w
0, \ O




---)-- ," ,. 0
\ 0) H
a) rO c\i












o O
Pr A L V O1 E
A 0










t ap
0 r-----C- rd 0)
cu a) aI 4











O m p,- m -o
0\ -t
|O o r. E-













:El
H E-






-- f


---- --- --- --- --- --- --- --- --- --- -^-* -- E


qT/nq 'r10 la'P /5H7 -o.rp Xd[eqMiua auTqan4 paqoaszoo









KACA EM E58C12 47









10
0
0' 0
Ln
___ -2f- -________ _________________
< o --- -- -- --- -- -- --- --3-- --
0/ LO-

mm
0 4
H 0 w*O
___C __L ^ T"^"'__- ___t ___ ______ -^ m"
rl oF^ ( -^" Ifl



02E^ r'^ ^ **1
cr I\r~S,-'s
rtd -P
0 0P c1




__ 0j \V _-'y__ o s
4 ( P2 c(
H 4 0
SH0 "1




0 --H rb ~ ^\ M -
0 ca
co to l 4- o
to $4:b


CD rd) P,

--- -- --_ _. ^ ^ -- ^ ^ -- -- _o g0 i ^
00
U '3



o___ s ___ o ( tc (1-1
0e H~ 3f
(00 0
-o '

co-'



_-" ~^ ___ ^ \/ -pa o
-- --'3- ^ B- "' -s 0, 3~I
0 HO H
r oto
H *



___ __ __ __ K :=____ s *C'3 __ (4- v
9) '3
0 p

o a )



S____ __ ___ ___ | -__ ___\: 08 -H
%H 9
.1-4- i 0 o0
w) a)"
+1 p, ? fi- \ /rt 1
-P A
l ~0 En to VM -
Ip 01 0 *

o .1
t\ I
| -- -- -- -8-- -- \- a


qI/nTa e 'e/Jsv 'dorp lXdTLqua auTqjnq paoai.xxo











NACA RM 158C12


Sm0

ou co




; -t --- U;







000



0 o
o IC

TI
% II


























MuTq.xnr
cti a)b o


--- H














- <__ C


cJ


L L
C%] co
M 00


oas/qj



@uTqmnq paatioo


to

x
53
1 e
(D

0
a,
*^r






I


(U
cu


51

.11
EA









RACA RM E58C12 49


LO




















oto
m 0
It)
-- ,-ooo---'







tO *




o o




oo3
,r
*0 0 0I w H| 1 4-
0 1 C4H




a)?F H o



P,
ODO<^ ^r-l


Sd/(6d ld) '0oTe1 sso[ arnssaid-T iq.; adidTidi,










NACA RM E58C12







0
---- O










0
0ri 0







N .,..t
N(
p4














c0
N N
N u
0 0
HH















0 O
O "63





__.d I



- -- ---- i-- --- K) +3

,," i




r


o*
+' -3







ri-
---- p -- +3









uaTDT TIaoD
Szo-


A3












NACA RM E58C12


8
CD3
cO
0












1 08 0









to O V\) .,4
r#














w 0
__ ___ 0
r 4







0 F,

Co 0 Cto C4



'4oTrdumsuoD TCaJ 3rds pCDasxo~ 0 4


CC 0O

44 t







CC) Cl|
S0F 0






d a
D0 o0O
C) 0












n o -p \ l




ooc - $





-- I -- -- I -- I -- --o- -- o S)O t fl


qT tSg/tj 'qsnmzq4 lan patoaijo3














NACA RM E58C12


li i|!j |^-lllil I-T|ieClir ;

















3t Ri S ; ;&i ;t: = =t- ^ ;:: .;._. .: .^ .^ -. -. .37
T- '-- -T z



U3 ...






1 17
Im
.Tv













4..
-7t
fi Ml- El

0-








r
TT q l




o -I. --tnLL
GMi


cu .. 4 = --


r),.


8


c d/ -a T au- H -


id/^S 'Od aI ajnssaad-TesOa aa;83


oo
a
0





0
o
0






CD
o
o0
r-
0













m

2,
SS

0





0 -
0-


o C-




C.


^iffiii^ifPii
7 -;i i












; v ::: T


,-, L-. t-











T,


U3 CJ ^-;p










NACA RM E58C12


i 0o
-$4


km 0

0D 0










M ca
4)
h x







t LO






IP P
or4 'a
4-> OJ ,O




l-i 3 So


'dO 'd o
w-^ H
to




o H
C1 1


.08 .12 .2 .4 .6
Compressor-inlet Reynolds number index, 52/2P- 2


Figure 16. General trend of engine pumping data with
Reynolds number index.










NACA RM E58C12


Performance loss
due to losses in -



wa, 2g -
Corrected net thrust, Fn/52 = 8470 Ib 69




.9 nd-
T
-- -- -- -- -- -- -- -- -- -- -- -- (Pp Pg) -- -




1.4
(P5 P9)
-- P5 -





and



B1.1


35,000 40,000 45,000 50,000 55,000
Altitude, ft


60,000 65,000 70,000 75,000


Compressor-inlet Reynolds number index, 52/20/02


Figure 17. Effect of individual component performance on over-all engine
performance over range of altitudes. Flight Mach number, 0.9; engine
speed, 7460 rpm; exhaust-gas total temperature, 15300 R.


43 43
w0




_U)



00





4--



0? -



4'-
G3 cl
cn)
&P ::









NACA RM E58C12


1801 -- -- -i-~t t


4- -. .. ..



160-
z: _7


S temperature,
8T- -7 2 r -T
.-.t ..



140i 7 j0.-- y
b' ~ IL~L temperature, `'" '





















60C0
407
ji -1--f-

.200~


4.

-.<1Tt
-- -:-.I T





I 7
t. .:T- .. I 'r

4-1 1P .; ---~







*--- I At -r- -- i:




E4
400I -47T
I T77. ... i1 ,.-
8 00 .4.~ .... ,. ,.


41 r -E: -

In 4


7T -E4 ---
.- i... ..... ....... .....i~~
T'.

t 00 4-t- r L.IIL I~- ~ ~
--~- C ~ C l- ~ --- CL~_ --I IEI* 41: -


.004


. .




l^| li li |II I l.ll i. .l.i|..0l

tiil!


Ideal fuel-air ratio, (wf/3600wa)id


Figure 18. Ideal fuel-air ratio for fuel used in this investigation.


zuut


.008


.012


.016


.020


.024


:-f 1- Illti ~ I ii r'


hafttia-a

















r t 4 t-' r


I 4q --. ..
-T











7l?^L I : [^ il^ ^^l| l'


Itti-t


i wttlttui :tt 4i -T ti- l t4 i rt S


h tEt- I t- f1 fEA t1 Nfi^ it E


4-)
cd

09
o cs
4-) ,-
cd -H
'm
p1
U'J


o In'o

H- H- H-^


i. -r


pJJa 'na1q.eanreJrd jtqPoo-TA aA-FlD4jQJ3


NASA Langley Field, Va.


NACA RM E58C12


O0
o
to P-





C,3
m



r -



0









0
4I

4-,




01






:3











o
cO






Ela
CO


*4
Co









'I
03
0















tCO
4-'


T'-


-t 1

i:i11::

:ntra- -
-Hti -t^ t




I





UniIVH- 0I I 0 U ILU' IUA


3 1262 08106 582 2